Patent Document

CROSS-REFERENCE TO RELATED APPLICATION(S)  
       [0001]    This application claims the benefit of U.S. Provisional Applications No. 60/307,564, filed Jul. 23, 2001, and No. 60/308,181, filed Jul. 27, 2001 which are incorporated by reference herein in their entirety. 
     
    
     
       BACKGROUND OF THE INVENTION  
         [0002]    1. Field of the Invention  
           [0003]    The present invention relates to in orbit spacecraft servicing systems and methods, and more particularly, to re-supplying an orbiting spacecraft using a low cost space tug.  
           [0004]    2. Brief Description of Earlier Developments  
           [0005]    There is an ever increasing interest in the commercial exploitation of space. Part of this interest is directed at an area which is believed to be more exploitable in the near future by placing orbital platforms/spacecraft to conduct any number of commercial and scientific space activity. These activities may include for example scientific observation, space tourism, to any number of micro-gravity industrial application including material manufacturing/processing, pharmaceutical and bio-chemical processing to mention just a few. The orbital platforms may be manned or unmanned, and may be pressurized or unpressurized. Some of these aspirations have been realized in a very limited fashion, and with very limited commercial success with earlier orbital “space stations” such as Skylab, the Soviet space stations including Mir, and today with the International Space Station (ISS) which has become operational. It can be envisioned that many more orbital platforms will exist in the future, although their number will depend on how economically they will be operated. As it is on Earth, the success of any commercial space venture will be its ability to produce a profit, and hence, the operating costs must be considered. Even in the case of non-profit or scientific ventures operating costs are a main concern. As can be immediately recognized, a very significant part of the operating costs of any space based venture is the cost of boosting consumables, repair parts, spaces and other operating items used in the operation of the venture to the orbital platform. The current approach is illustrated by the manner in which the ISS is resupplied. Consumables and other resupply items for the ISS are placed as stowed payload on a reusable launch vehicle (i.e. the manned space shuttle). Otherwise, the resupply payload may also be placed in the payload module on the upper stage of an expendable launch vehicle (such as for example the Russian Progress/Proton expendable launch vehicle). In both cases, the launch vehicles are launched from fixed launch sites, such as the Kennedy Space Center during a suitable “window” when the ISS is located with respect to the launch site to ensure that the launcher delivers the payload to the ISS. In the case of the reusable launch vehicle, the launch vehicle itself docks with the ISS to deliver the payload. In the case of the expendable launch vehicle, the upper stage delivers the payload module to the ISS. Neither of these launch means have proven adequate to provide launch costs of resupply items which are satisfactory.  
           [0006]    Moreover, launching from a fixed launch site, as is presently the case, generally results in substantial waiting periods in orbit to achieve phasing for rendezvous opportunities, or requires inefficient dog-leg maneuvers to alter the orbit plane during the ascent phase. This has an adverse effect especially on delivery of priority payload to the orbital platform. The present invention overcomes the problems of the conventional resupply methods by employing are supply system having a mobile launcher and a low cost space tug or reusable transfer vehicle. The mobile launcher uses an air- or sea-launch capability that is highly advantageous for rapid call-up launch, where payload delivery would occur as soon as possible after cargo liftoff. The total time from receiving a call for launch and on-orbit delivery would be one week or less. This is not possible using the conventional launch systems. The space tug allows for maximization of the cargo mass fraction of the launch vehicle final stage, as will be described further below, which makes most efficient use of mass ferried through maneuvers to the orbital platform.  
           [0007]    Conventional launch system, as described previously, suffer a further handicap which results in reduced cargo mass fraction for the launch vehicle final stage because of the configuration of the payload module of the launch vehicle final stage. As noted before, the payload module of the final stage in the conventional launch systems is docked to the ISS to effect resupply. Accordingly, the construction/structure of the payload module is “heavy” to provide a pressurized volume and withstand the large stresses associated therewith. The pressurized volume in the payload module is needed to allow the module to be directly interfaced to the pressurized docking port of the ISS. Payload mass thus may be in the order of about 15% of the total payload module mass in conventional launch systems. The system of he present invention uses a payload module that is not pressurized. An unpressurized payload docking adapter interfaces with the payload module, allowing the unpressurized payload module to be docked to the pressurized docking port of the orbital platform. As can be realized, by using an unpressurized payload module the cargo mass fraction of the payload module may be increased significantly with corresponding benefits to the payload launch costs. The system of the present invention overcomes the problems of the prior art as will be described in greater detail below.  
         SUMMARY OF THE INVENTION  
         [0008]    In accordance with a first embodiment of the present invention, an orbiting spacecraft payload delivery system for delivering a payload to an orbiting spacecraft is provided. The system comprises a launch vehicle, a mobile launcher, and an orbiting reusable space tug. The launch vehicle has a payload module for holding a payload for the orbiting spacecraft. The mobile launcher is adapted to transport the launch vehicle. The orbiting reusable space tug ferries the payload module to the orbiting spacecraft. The mobile launcher transports the launch vehicle to a predetermined location disposed within a predetermined distance from a ground track of the orbiting spacecraft for launching the launch vehicle from the predetermined location.  
           [0009]    In accordance with another embodiment of the present invention, an orbiting spacecraft payload delivery system for delivering a payload to an orbiting spacecraft in a predetermined orbit plane is provided. The payload delivery system comprises a launch vehicle, a mobile launcher, and a reusable orbiting tug. The launch vehicle has a payload module for holding the payload for the orbiting spacecraft. The mobile launcher is adapted to transport the launch vehicle. The reusable orbiting tug ferries the payload module to the orbiting spacecraft after the payload module is released from the launch vehicle. The mobile launcher transports the launch vehicle to a predetermined location for launching the launch vehicle from the predetermined location. The predetermined location is in the orbit plane of the orbiting spacecraft. 
       
    
    
     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 diagram of an in orbit spacecraft resupply system incorporating features of the present invention and an orbiting spacecraft  100 ;  
         [0012]    [0012]FIG. 2 is a 2-D graphical representation of the ground tracks of the orbiting spacecraft  100  in FIG. 1;  
         [0013]    [0013]FIG. 3 is a schematic top view illustrating a portion of the trajectory path of a launch vehicle of the system in FIG. 1 when launched to deliver a payload to the orbiting spacecraft;  
         [0014]    [0014]FIG. 4 is a schematic side view illustrating the orbit paths of the launch vehicle (LV), and space tug of the system in FIG. 1, to resupply the orbiting spacecraft in accordance with a first method of the present invention;  
         [0015]    [0015]FIGS. 5 and 5A are respectively a schematic perspective view, and a schematic elevation view of a launch vehicle final stage (LVFS) of the system in FIG. 1;  
         [0016]    [0016]FIG. 6 is a perspective view of the reusable space tug of the system in FIG. 1;  
         [0017]    [0017]FIGS. 7A and 7B are respectively perspective views in opposite orientations of the LVFS and space tug stack, and a payload docking adapter of the system in FIG. 1;  
         [0018]    [0018]FIG. 8 is a schematic diagram depicting a rendezvous portion of the orbit path of the LVFS/space tug stack with the orbiting spacecraft;  
         [0019]    [0019]FIG. 9 is a is a schematic side view illustrating the orbit paths of the LV, and space tug of the resupply system to resupply the orbiting spacecraft in accordance with a second method of the present invention;  
         [0020]    FIGS.  9 A, and  9 B are graphs respectively illustrating the difference in right ascension of the ascending node (RAAN) of the orbits of the space tug and orbiting spacecraft, and the difference in the argument of latitude (ARGLAT) between tug and orbiting spacecraft when operated according to the second method; and  
         [0021]    [0021]FIG. 10 is a schematic elevation view of a LVFS in accordance with the prior art. 
     
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT  
       [0022]    Referring to FIG. 1, there is shown a schematic profile view of a system  10  incorporating features of the present invention for resupplying an orbiting spacecraft  100 . 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.  
         [0023]    Referring still to FIG. 1, the orbiting spacecraft  100  is depicted as a representative spacecraft in earth orbit, though the present invention is equally applicable with any suitable type of spacecraft. The spacecraft  100  generally has a bus or service module  100 A and a command or payload module  100 B connected to the service module. The service module  100 A generally may include a power system  100 P for generating and supplying power to the spacecraft, attitude control system (not shown) with transfers and/or momentum wheels for attitude control as well as for station keeping, and any other suitable support systems for operating the spacecraft and command module. The spacecraft  100  may be a manned spacecraft that may be continuously and permanently manned such as for example the international space station (ISS), or manned only periodically. In this case, the command/service module may includes a habitation areas for the crew as well as suitable support systems for operating the habitation area. The spacecraft  100  may also be an unmanned platform with no capability for on board personnel when in orbit. In this case however, the command module of the spacecraft may include any desired automated systems for performing any desired processing. For example, the command module may include a fully automated system for material processing (e.g. forming complex chemical compounds, forming material alloys) in microgravity. In any event, it is desired that the spacecraft  100  be periodically resupplied in orbit. Accordingly, the command module, and/or the service module, has a docking port (shown schematically at  100 C in FIG. 1) through which resupply payload (e.g. consumables for the crew or for the automated manufacturing process) may be brought into the spacecraft. The resupply payload is launched to the spacecraft  100  with the system  10  shown in FIG. 1 as will be described further below. The system  10  and its operation will be described in the case of the spacecraft  100  being in LED in general with an orbit having the parameters of the ISS orbit in particular, for example purposes only. The present invention is equally applicable to the case of spacecraft in any LEO orbit or in MEO or GEO orbits. The spacecraft  100  in this exemplary case has a circular orbit, represented at O in FIG. 4, with an altitude of about 450KM and an inclination of about 51.6° (which is very close to the maximum altitude of the ISS).  
         [0024]    As seen in FIG. 2, the in-orbit spacecraft resupply system  10  generally comprises a launcher  12 , a launch vehicle  14 , and a reusable space tug  56 . The launcher  12 , which is a mobile launcher in this case, holds and deploys the launch vehicle  14  for launch as will be described below. The launch vehicle  14 , has a final stage (LVFS)  20  with a payload module  22  holding a payload destined for the orbiting spacecraft  100 . The launch vehicle  14  boosts the LVFS  20  into orbit. FIG. 1 shows the LVFS  20  in a number of positions I 1 -I 4  along its path to rendezvous with the target spacecraft  100 . The reusable space tug  16  is in orbit. The space tug  16  is a low cost spacecraft with appropriate guidance and control capability to rendezvous and capture of the LVFS  20  and to captive-carry The LVFS to the spacecraft  100  to deliver the payload. The space tug  16  is a spacecraft which may not have a refueling capability. The system  10  may also include a payload docking adapter  200  for interfacing the payload module of the LVFS to the spacecraft  100 . The adapter is disposed in orbit, and if desired may be carried by the spacecraft  100 .  
         [0025]    Referring now to FIG. 2, there is shown a 2-D graphical representation of the ground track of spacecraft  100  in orbit O over a representative period of time (such as for example 26 hours). FIG. 2 also shows a representative fixed ground launch site CAPE which may be used by conventional launch systems. As can be seen in FIG. 2, in the time span represented therein, the fixed site CAPE is not traversed by the ground track of spacecraft. Moreover, it is expected that the ground track would not traverse site CAPE for a period of days. Accordingly, as was noted before, any launch from fixed site CAPE during the period represented in FIG. 2, to resupply a spacecraft in orbit O with a conventional launch system, would employ a significant plane change during ascent, or a phasing delay of a few days in orbit, or both. This would adversely impact a “rapid call-up” missing to bring priority or “just in time” resupplies to the spacecraft. The launch and delivery system  10  of the present invention overcomes this problem.  
         [0026]    In greater detail now, and with reference again to FIG. 1, the mobile launcher  12  is shown for example purposes only as being in general an airborne launcher and in particular as being an airplane, though any suitable type of aircraft may be used including a lighter than air vehicle. In alternate embodiments, the mobile launcher may be sea borne or water borne based (e.g. ship, towed platform, submersible) or may be ground borne (e.g. ground transporter/launcher). An airborne mobile launcher  12  provides the greatest range over a short time period, and flexibility to reach a desired location traversed by the ground track of the spacecraft  100  in orbit O. The mobile launcher  12  may be such as provided for example by Coleman Aerospace. The Mobile Launcher  12  may include a suitable cargo/transport aircraft carrier and a stowage/deployment system for the launch vehicle. The stowage/deployment system is used to stow the launch vehicle aboard the aircraft carrier and to deploy the launch vehicle rapidly from the aircraft carrier during flight. After separation from the carrier, the deployment system stabilizes the launch vehicle and releases the launch vehicle when still airborne for launch. The aircraft carrier and deployment system of the mobile launcher  12  cooperate to deliver the launch vehicle  20  for launch from an altitude of 15,000 to 20,000 feet. In alternate embodiments, the mobile launcher may deliver the launch vehicle for launch from any desired altitude.  
         [0027]    Referring still to FIG. 2, the launch vehicle  14  is depicted as a representative expendable launch vehicle which includes a number of stages as desired, and at least a final stage (LVFS)  20  that has the payload module  22 . In the case of the airborne mobile launcher  12  shown in FIG. 1, the launch vehicle  14  is sized to be carried inside or outside of the airborne carrier. In alternate embodiments, the launch vehicle may have any other desired size suited for transport and launch from a desired mobile launcher. In the case shown in FIG. 1, a suitable launch vehicle may be the small Expendable Launch Vehicle concept from Coleman Aerospace. In alternate embodiments, the launch vehicle may be a reusable launch vehicle with one or more reusable stages. The launch vehicle  14 , may be for example, a solid booster launch vehicle. The main space of the launch vehicle may include suitable positioning and guidance systems (GPS/INS) to enable the launch vehicle  14  to attain the appropriate altitude when released from the deployment system of the mobile launcher to launch into orbit O of the orbiting spacecraft  100  as will be described further below.  
         [0028]    Referring now to FIGS.  5 - 5 A, there is shown a perspective view and a schematic elevation view of the launch vehicle final stage (LVFS)  20 . The LVFS  20  is depicted as a representative final stage of the launch vehicle, and in alternate embodiments, the final stage may have any suitable configuration. As seen in FIGS.  5 - 5 A, the LVFS  20  generally comprises a service module  32  and payload module  22  (indicated in phantom). The service module  32  includes structures and system for coupling the LVFS to the launch vehicle during launch. The LVFS  20  also has a thruster  32 T for injecting the LVFS into orbit (indicated by position Z 1  in FIG. 1). In this embodiment, the LVFS thruster or maneuvering system may be provided with minimal propellant for on-orbit maneuvering. Thus, the LVFS  20  generally has no significant on-orbit maneuvering capability and remains substantially passive once injected into orbit (position Z 1 ). As can be realized, this allows the LVFS to carry a larger payload mass. The service module  32  may have a deck  32 D for supporting the payload in the payload module. As seen best in FIG. 5, the LVFS  20  may have an intermediate section  32 I disposed between the payload module  22  and service module  32 . The intermediate section  32 I may have a general disk shape with surfaces  32 S configured to form an airtight seal when coupled wit the adapter  200  at the orbiting spacecraft  100  as will be described further below. As seen in FIGS.  5 - 5 A, the payload module  22  is not a module in the conventional sense. Rather, the payload module  22  generally comprises a skeletal support structure  30 , depicted in a representative manner in FIGS.  5 - 5 A, and cargo canisters  36 . The representation of both the support structure  30  and cargo canisters  36  in FIGS.  5 - 5 A are merely exemplary in nature, and in alternate embodiments the support structure and cargo canisters of the LVFS may have any suitable configuration and size. The support structure  30  depends from the support deck  320  of the service module  32  and has mounting supports  30 M for holding the cargo canisters  36 . The support structure  30  may have longitudinal support posts, columns or trusses  30 P (only one representative post  30 P is shown in FIGS.  5 - 5 A for example purposes) to connect the mounting supports  30 M to the service module  32 . The mounting supports  30 M extending from the support posts  30 P are shown as being attached to the cargo canisters  36  from an exemplary location (at the top of the canister) though in alternate embodiments, the canisters may be held from any number of suitable locations using any suitable holding means such as clamping, and mechanical fastening with threaded fasteners. As can realized, the cargo canisters  36  are sized in length and width to be admitted through the resupply port of the orbiting spacecraft  100  (see FIG. 1). The ends of the canisters may be removable to provide access to the interior of the canisters, otherwise the canisters may be provided with access panels. The structure of the cargo canisters is capable of maintaining a pressurized interior, and the removable ends or access panels form an airtight seal when closed. The cargo canisters  36  are shown as having a generally cylindrical shape with hemispherical heads at opposite ends for example purposes only, and in alternate embodiments the cargo canisters may have any suitable shape. The cargo canisters  36  in the payload module  22  are also shown as being substantially the same, for example purposes, and in alternate embodiments, the payload module  22  may have different sizes of canisters to suit the payload demands. The support structure  30  may hold any desirable number of canisters (four canisters are shown in FIGS.  5 - 5 A for example purposes only) that may be included within the surrounding fairing covering the LVFS  20  or at least the payload module  22 . As can be realized from FIGS.  5 - 5 A, with the exception of cargo canisters  36 , the payload module  22  of the LVFS  20  may be unpressurized. The structure of the payload module  22  of LVFS  20  is thus very simplified, and much lighter in comparison to the conventional payload module structure  322  of a conventional launch vehicle final stage  320  as illustrated schematically in FIG. 10. The payload module  22  of LVFS  20  also does not employ a heavy hatch for mating with the docking port on the orbiting spacecraft. This reduction in mass of the payload module structure in comparison to conventional payload modules provides a concomitant improvement in the payload mass capability of the LVFS  20  compared to conventional final stages as seen in FIG. 5, the LVFS  20  may also include a small compressed air tank  38 , that may be located in the payload module  22  if desired, to provide pressurization air when the LVFS  20  is mated to the adapter  200  as will be later described.  
         [0029]    Referring now to FIG. 6, there is shown a perspective view of the orbiting space tug  16  of the resupply system  10 . The tug  16  is illustrated in FIG. 6 as a representative small spacecraft, and in alternate embodiments any suitable spacecraft may be used.  
         [0030]    In this embodiment, the space tug  16  is shown in FIG. 6 as being similar to or a derivative of the Loral LS-400 spacecraft bus for example purposes. In general, the tug  16  has a bus  40  with a power system  42 , a propulsion system  44 , an altitude control system  46  and a command and guidance system  48 . The bus  40  also includes docking or grappling clamps  50  for capturing and holding the LVFS during ferry maneuvers to the orbiting spacecraft  100 . The power system  42  may include solar panels or any other means for generating power which is distributed by the power system to any of the other support (propulsion, altitude control, command and guidance) systems. The propulsion system  44  may include a thruster(s) with a desired propellant to perform the ferry maneuvers as will be described further below. The altitude control system  46  may include momentum wheels and/or thrusters to provide 3-axis stabilization of the space tug when mated to the LVFS. The command and guidance system includes communication electronics/antennas, suitable position/altitude sensors and processors to provide the space tug  16  with appropriate guidance and control capability and appropriate redundancy to rendezvous and dock with the LVFS  20 , captive-carry the LVFS to the orbiting spacecraft  100  and deliver the LVFS  20  to the docking port of the orbiting spacecraft  100 . Table 1 below, lists exemplary operating parameters for three spacecraft bus types that may be used as the space tug of the resupply system  10  (see FIG. 1).  
                                 TABLE I                           Orbiting Spacecraft is at 450 km altitude.            Description   Type 1   Type 2   Type 3                   1 st  Generation   2 nd  Generation   Space-based           Tugboat   Tugboat   Upper stage       Orbit altitude   450 km   200 km   80 km       used for launch       vehicle target       Rapid call-up   Yes   LV must target   Yes               450 km,               otherwise 2-day               delay       Dry mass   300 kg   1000 kg   1000 kg       Propellant   80 kg   2200 kg   21000 kg       capacity       Refueling   No   Yes   Yes       necessary       Propellant   Monoprop.   Bi-propellant   Cryogenic           Hydrazine   (NTO/MMH)   (LOX/LH2)       Specific Impulse   220 s   310 s   440 s       Range of   400-500 km   200-500 km   80-500 km       altitude       Typical ΔV per   20 m/s   340 m/s   4200 m/s       cargo retrieval       mission       Typical cargo   1500 kg   1660 kg   11000 kg       mass       Typical ferried   2070 kg   2230 kg   11600 kg       mass       Typical   20 kg   500 kg   20000 kg       propellant mass       per mission       Number of   4 (not   4 per refueling   1 per       missions   refueled)       refueling       Spacecraft bus   L400 (Loral)   L1300 (Loral)       Aerodynamic   No   No   Ballute       decelerator?       Time for   3 hours   3 hours   1 minute       rendezvous                  
 
         [0031]    As can be realized from Table I, the Type 1 spacecraft bus provides a low cost space tug with modest maneuvering capability. In this case, the space tug  16  would not be refuelable and may be de-orbited after completing a number of missions (the number of missions shown in Table I, along with the other figures indicated in the table for all spacecraft types are merely exemplary and in alternate embodiments various parameters may be different as desired). The Type 2 and Type 3 spacecraft buses have a progressively larger maneuvering capability and may be refueled between missions as indicated in Table I. Accordingly, the Type 2 and Type 3 spacecraft represent more costly spacecraft buses for the space tug in comparison to the Type 1 spacecraft bus. A suitable means for refueling the Type 2 or Type 3 spacecraft bus is described in U.S. patent application Ser. No. 09/598,128, filed on Jun. 21, 2000 which is incorporated by reference herein in its entirety. FIG. 4 is a schematic side view that illustrates a launch and rendezvous profile used with the resupply system  10  in FIG. 1 in the case when the space tug  16  is substantially a Type 1 spacecraft bus. FIG. 9 is another schematic view that illustrates another launch and rendezvous profile of the resupply system in the case when the space tug was a Type 2 spacecraft bus. FIG. 11 is another schematic diagram illustrating still another launch and rendezvous profile of a resupply system  10 ′ having a space tug  16 ′ with a Type 3 spacecraft bus.  
         [0032]    The resupply system  10  in FIG. 1, allows for a rapid call-up launch for resupplying the orbiting spacecraft  100  in orbit O. The system  10  allows payload delivery to occur as soon as possible after cargo liftoff. The total time from receiving a request for the payload to on-orbit delivery of the payload may be about one week or less. This would not be feasible with conventional launch systems. As noted before, FIG. 2 illustrates that in a given period (in this case a representative period of 26 hours) the ground track of spacecraft  100  in orbit O does not traverse a given ground site CAPE. There are however, a number of segments G 1 -G 4  of the ground track that pass sufficiently close to the given site CPAE so as to be within the range of the mobile launcher  12  in the event the mobile launcher is based proximate site CAPE. Both northbound G 1 -G 2 , and southbound segments G 3 -G 4  pass within the range of the mobile launcher  12 . This in effect doubles the launch opportunities to resupply the spacecraft  100  using system  10  as will be seen further below. Referring now to FIG. 3, there is shown a schematic top view illustrating a representative ground track segment GO of the spacecraft  100  and the trajectory path of the launch vehicle  14  to boost the LVFS  20  into orbit. The ground track segment G 0  is shown for example purposes only as being a northbound segment. As noted before, the system  10  is capable of launching the LVFS  20  for rendezvous with the orbiting spacecraft  100  when the spacecraft has a southbound ground track segment as will be evident below. As shown in FIG. 3, the mobile launcher  12  of system  10 , carrying the launch vehicle  14  travels (from a given basing site which may be anywhere within the range of the mobile launcher from the given ground track segment G 0 ). Along the path indicated by arrow FP to a designated launch point LP. The launch point LP intersects the ground track segment GO of the spacecraft  100  (i.e. is in the orbit plane of the spacecraft  100 ). As can be realized, the launch point LP is established to define an angular offset (not shown) from the orbiting spacecraft  100  of appropriate angle at the time of main engine start of the launch vehicle  14 . As the mobile launcher  12  reaches the launch point LP at the appropriate time, the launch vehicle is deployed from the mobile launcher and disposed in the proper orientation for launch. The launch azimuth of the launch vehicle is not restricted. This allows the launch vehicle  14  to launch the payload in the orbit plane of the orbiting spacecraft. The ascent trajectory indicated by arrow A 7  in FIG. 3) follows (i.e. is coplanar with) the ground track G 0  of the orbiting spacecraft  11 . The unrestricted launch azimuth of the launch vehicle  14  allows the launch vehicle to along the northbound track G 0  or along a southbound track.  
         [0033]    [0033]FIG. 4 shows a side view of the launch and rendezvous of which only a portion is shown in the top view of FIG. 3. At the time of main engine start of the launch vehicle  14  (at position LP in FIG. 3) the space tug  16  is at position R 1  along the orbit path O. In this case, the orbit of space tug  16  when in position R 1  is in substantially the same plane, and has substantially the same parameters (e.g. inclination i, RAAN, inclination, eccentricity, altitude) as the orbit O of spacecraft  100  to be resupplied. In addition, the angular offset between the spacecraft  100  and the space tug  16  when the space tug is at position R 1  is such as to allow performance of the ferrying mission as will be noted below. As shown in FIGS. 3 and 4, the launch vehicle boosts the LVFS  20  to a position I 1  where the engine  32 T of the LVFS fires to inject the LVFS into orbit. The LVFS achieves orbit in position  22  as shown in FIGS. 3 and 4. The LVFS  20  at this point has an orbit which is also in the same orbit plane as both spacecraft  100  and space tug  16 , and has the same orbit parameters as orbit O of the spacecraft  100  and space tug. FIGS. 1 and 4 show in this case that the LVFS  20  is offset back (i.e. opposite the direction of orbit indicated by arrow D) from the spacecraft  100 , but ahead (i.e. in the direction of orbit) of the space tug  16 . In alternate embodiments, the positional order between the spacecraft, LVFS, and tug may be reversed if desired with the tug forward most in the orbit direction, then the LVFS, and the spacecraft last. The order shown in FIGS. 1 and 4 may facilitate rendezvous with the spacecraft  100  from below which may be desirable in the case of the ISS. As noted before, after achieving orbit, the LVFS  20  has substantially no maneuvering capability. The tug  16  propulsion system  44  (see FIG. 6) is thus used to rendezvous and dock with the LVFS  20 . This is schematically illustrated at position I 3  in FIGS. 1 and 4. In order to rendezvous with the LVFS  20  in the case, the tug  16  may generate a velocity increment of about 5 m/s. The tug  16  captures and docks to the LVFS  20  using capture system  50 .  
         [0034]    Referring now to FIGS.  7 A- 7 B, there is shown respectively to opposing perspective views of the LVFS and tug stack  16 ,  20 . In this embodiment, the tug  16  captures the LVFS from the rear end  32 R of the LVFS. Accordingly, the stack  16 ,  20  has the tug positioned at one end, and the payload module  22  of the LVFS located at the other end of the stack. In alternate embodiments, the tug may hold the LVFS in any other suitable arrangement (e.g. such as a piggyback arrangement). The propulsion system  44  of the tug  16  is then used to reposition the LVFS tug stack  16 ,  20  and dock the payload module  22  to the orbiting spacecraft (indicated respectively at positions I 3 A and I 4  in FIGS. 1 and 4). The tug propulsion system may generate a velocity increment of about 10 mk to reposition the stack and rendezvous with spacecraft  100  in the case. FIG. 8 shows a schematic side vie of the LVFS/tug stack  16 ,  20  in a number of positions I 3 A-I 4  as the stack approaches and docks with orbiting spacecraft. In FIG. 8, the orbiting spacecraft is depicted as the I 3 C which has its docking port  100 C facing the earth. In positions I 3 B and I 3 C the stack  16 ,  20  approaches the spacecraft from behind and below. Relative motion may be about 4 km/hr. At position I 3 D the tug  16  orients the stack to be aligned generally with the docking port  100 C of the spacecraft. The tug  16  then moves the stack in for docking (positions I 3 D,  83 E). In this case, spacecraft  100  is provided with a remote manipulating system (RMS) which may be used as shown in FIG. 8 to capture the stack and assist in docking the payload module to the docking port  100 C (positions I 3 F, I 4 ). In alternate embodiments, the tug  16  may maneuver the stack directly into contact with the spacecraft for docking.  
         [0035]    The payload module  22  of the LVFS  20  in this case is docked to the docking port  100 C of this spacecraft using unpressurized payload docking adapter  200 . FIGS.  7 A- 7 B show opposing perspective views of the adapter  200 . As seen in FIGS.  7 A- 7 B, the adapter  200  has a substantially cylindrical body with a hatched port  202  at one end and an LVFS interface port  204  at the other end. The LVFS port  204  may be configured to be open when not mated to the LVFS. Thus, the adapter  200  is unpressurized when not being used. The hatched port  202  is configured to be sealably coupled to the resupply port  100 C of the spacecraft. This allows operators inside the spacecraft access to the hatch of the hatched port  202 . The adapter  200  includes an appropriate vent  208  which is normally closed. The exterior of the adapter  200  may have a number of grappling fixtures allowing the RMS to grab and maneuver the adapter  200  for coupling with the port  100 C. In this embodiment, the adapter may be stowed at a suitable location on the spacecraft and then berthed to the port  100 C prior to a payload delivery mission. The LVFS interface port  204  may have movable doors  206  (two doors are shown, but any number of doors may be used). The doors  206  may be moved in the direction indicated by arrows Y between disengaged and engaged positions. In the disengaged position, the doors  206  are moved sufficiently apart to allow the payload module  72  (i.e. the canisters and support structure) of the LVFS  20  to enter into the adapter  200  in the direction indicated by arrow Y in FIG. 2B. In the engaged position, the doors  206  are positioned to be seated against surface  32 S of the LVFS thereby forming an airtight seal. The air canister  38  (see FIG. 5) on the LVFS may then be used to pressurize the adapter  200 . After being pressurized, the hatch in port  202  may be opened to access the payload canisters  36 . The payload canisters  36  may be removed as noted before from the LVFS and brought through the adapter into the spacecraft for unloading the payload. The empty canisters may then be filled with any desirable materials for removal from the spacecraft  100 .  
         [0036]    [0036]FIG. 9 shows an alternate approach for launch and rendezvous with the orbiting spacecraft. This involves the tug performing an out of plane maneuver as depicted in the graphs shown in FIGS.  9 A- 9 B. This consumes a larger amount propellant than the approach shown in FIG. 4, and would employ a tug having a Type 2 or Type 3 bus.  
         [0037]    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.

Technology Category: b