Abstract:
A space module has an outer structure designed for traveling in space, a docking mechanism for facilitating a docking operation therewith in space, a first storage system storing a first propellant that burns as a result of a chemical reaction therein, a second storage system storing a second propellant that burns as a result of electrical energy being added thereto, and a bi-directional transfer interface coupled to each of the first and second storage systems to transfer the first and second propellants into and out thereof. The space module can be part of a propellant supply architecture that includes at least two of the space modules placed in an orbit in space.

Description:
[0001]    Pursuant to 35 U.S.C. §119, the benefit of priority from provisional application 60/309,977, with a filing date of Aug. 3, 2001, is claimed for this non-provisional application. 
     
    
     ORIGIN OF THE INVENTION  
       [0002] The invention described herein was made by employees of the United States Government and may be manufactured and used by or for the Government for governmental purposes without the payment of any royalties thereon or therefor. 
     
    
     
       BACKGROUND OF THE INVENTION  
         [0003]    1. Field of the Invention  
           [0004]    This invention relates to the storage, transportation, and supply of propellants in space. More specifically, the invention is a reusable space module that can warehouse, transport, and supply high-thrust propellant(s) needed for time critical space travel and low-thrust propellant(s) usable for non-time critical space travel.  
           [0005]    2. Description of the Related Art  
           [0006]    Many engineering problems must be solved in order to provide safe, affordable, and efficient in-space transportation of both personnel and equipment. These challenges directly impact the commercialization of space. In particular, cost is the single largest obstacle. One method of reducing costs is to develop reusable transportation systems for both Earth-to-orbit systems and in-space systems. Without reusable systems, sustained exploration or large-scale development beyond Low Earth Orbit (LEO) using in-space transportation systems will not be economically feasible.  
           [0007]    Significant challenges must be overcome for in-space transportation to become a reality. For example, reusable in-space transportation systems that are affordable must provide good fuel efficiency, be capable of mass production, and be compatible across a broad spectrum of applications. Another significant challenge involves how to minimize the in-space travel time for manned missions. The risk associated with human missions can be significantly reduced by decreasing the time that the crew is in transit.  
           [0008]    Currently, the primary propulsion method for manned and unmanned missions utilizes chemical propulsion systems. This trend is expected to continue for the foreseeable future. While nuclear thermal propulsion systems show promise as part of a highly efficient transportation system of the future, the major drawback of a nuclear thermal propulsion system is the inherent public concerns that accompany the use of these systems near the Earth. The chemical propulsion system provides a reasonably high thrust thereby making it a logical choice for time critical space travel. However, one significant drawback to chemical systems is the relatively low specific impulse of this form of propulsion. That is, chemical propulsion systems require large propellant quantities to provide the velocity changes needed to complete a mission. The need for a large amount of propellant translates into a larger, heavier space vehicle which, by itself, exacerbates the need for even more propellant. As a result, chemical propellant-based space vehicles are large and expensive.  
         SUMMARY OF THE INVENTION  
         [0009]    In accordance with the present invention, a space module has an outer structure designed for traveling in space. At least one docking mechanism is coupled to an end of the outer structure for facilitating a docking operation therewith in space. A first storage system mounted within the outer structure stores a first propellant that burns as a result of a chemical reaction therein. A second storage system mounted within the outer structure stores a second propellant that burns as a result of electrical energy being added thereto. A bi-directional transfer interface is coupled to each of the first and second storage systems to transfer the first and second propellants into and out thereof.  
           [0010]    The space module can be part of a propellant supply architecture that is used for in-space transportation. In this architecture, at least two of the space modules are transported to an orbit in space. A first of the space modules is coupled to a first space vehicle via its docking mechanism. The first space vehicle travels to a location in space using the first propellant (e.g., high thrust propellant when the first space vehicle is a manned vehicle) to generate thrust. A second of the space modules is coupled to a second space vehicle via its docking mechanism. The second space vehicle travels to the same location in space as the first space vehicle, but uses the second propellant (e.g., low thrust propellant when the second space vehicle is an unmanned vehicle) to generate thrust. When it is time for the first (manned) space vehicle to return, the first and second space vehicles swap space modules thereby providing i) a full supply of first propellant for the first space vehicle&#39;s return trip, and ii) a full supply of second propellant for the second space vehicle&#39;s return trip. 
       
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0011]    [0011]FIG. 1 is a top-level diagram of a reusable space module for the storage, transportation, and supply of multiple propellants in accordance with the present invention;  
         [0012]    [0012]FIG. 2 is a block diagram illustrating separate storage tanks for storage of a chemically initiated propellant&#39;s fuel and oxidizer;  
         [0013]    [0013]FIG. 3 is a schematic view of one embodiment of the space module configured for transportation and deployment in space by means of a shuttle-class reusable launch vehicle, and configured to withstand translational loading imported by a space vehicle(s) coupled thereto;  
         [0014]    [0014]FIG. 4 is an isolated side view of the frame of the outer structure for the space module illustrated in FIG. 3; and  
         [0015]    [0015]FIG. 5 depicts a space transportation scenario utilizing a propellant supply architecture based on the space module of the present invention. 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0016]    Referring now to the drawings, and more particularly to FIG. 1, a top-level diagram of a space module for the storage, transportation, and supply of multiple propellants is shown and referenced generally by numeral  10 . In general, space module  10  is a facility that combines propellant storage and supply in an in-space structural package that can be coupled to and used by space vehicles as their propellant supply during in-space travel.  
         [0017]    Because the present invention is not intended to be limited by a specific configuration or use in a specific in-space propellant storage/supply architecture, space module  10  will be described relative to its general inventive features. Then, by way of illustrative example, a more specific embodiment of the space module will be described in which transportation to space and deployment thereof in space is accomplished with a Shuttle-class reusable launch vehicle (RLV). Finally, an in-space propellant supply architecture will be described in which multiple space modules are used to supply propellant for an in-space mission.  
         [0018]    Space module  10  has an outer structure  12  constructed for in-space travel and the protection of the contents thereof from hazards of the space environment such as debris, excessive heat and/or radiation, etc. At a minimum, the contents of structure  12  include a store  14  of chemically initiated propellant, a store  16  of electrically initiated propellant, and a bi-directional propellant transfer interface  18 . At one end of structure  12  is a docking mechanism  20  to which a visiting module (e.g., space vehicle, satellite requiring refueling, or other in-space facility) can be mechanically coupled. Additional docking mechanisms can be provided on outer structure  12  as will be described later below. To facilitate the transfer of propellant both to and from each of stores  14  and  16 , transfer interface  18  is accessed at docking mechanism  20 . In general, transfer interface  18  provides individualized and standardized connections for the chemically and electrically initiated propellants in stores  14  and  16 , respectively.  
         [0019]    As used herein and as is well known in the art, a chemically initiated propellant is any liquid-form mono or bipropellant, solid propellant, or combination of liquid and solid propellant that burns as a result of a chemical reaction therein. Such chemically initiated propellants (e.g., liquid hydrogen/oxygen, hydrazine, kerosene, etc.) generally provide high thrust at the expense of a low specific impulse which is a measure of efficiency. Thus, chemically initiated propellants are preferred for use in space travel that is time critical as is the case for manned space missions. In contrast, electrically initiated propellants generally produce low thrust, but have a high specific impulse or efficiency. Thus, electrically initiated propellants could be used in space travel that is not time critical, e.g., ferrying of equipment, unmanned space missions, transportation of the chemically initiated propellant, etc. As used herein and as is well known in the art, electrically initiated propellants require the addition of electric energy (e.g., spark discharge, propellant ionization, solid resistive elements in the propellant, etc.) to initiate the burning thereof.  
         [0020]    A well known example of chemically initiated propellants is a liquid bipropellant having a fuel (e.g., liquid hydrogen) and an oxidizer (e.g., liquid oxygen) maintained in different storage systems until the burning thereof is desired. Accordingly, in such instances, store  14  provides a tank  14 A for the fuel and a separate tank  14 B for the oxidizer as illustrated in FIG. 2. Each of tanks  14 A and  14 B would be constructed as needed to safely store the contents thereof until they are to be supplied via transfer interface  18 . For long-term, in-space storage, it may be necessary to cryogenically cool the liquid fuel and oxidizer to maximize their densities and prevent their boil off during prolonged periods of time. Accordingly, one or more cryogenic cooling system(s)  22  may be provided onboard outer structure  12 . Cryogenic cooling system(s)  22  can be used to cool any of the propellants in store  14 .  
         [0021]    The electrically initiated propellant in store  16  can be any low thrust, high specific impulse propellant such as xenon, krypton, argon, helium, and even hydrogen. However, for current state-of-the-art electric propulsion systems, xenon and krypton are usually used because they provide sufficiently high thrust levels. If desired, the propellant can be cooled by cryogenic cooling system(s)  22 . For example, the cryogenic cooling of xenon will keep it in its higher density liquid state which means that a greater amount of the propellant can be stored in a given volume than if it were maintained in its gaseous state.  
         [0022]    As will be explained further below, one use of space module  10  involves the deployment thereof by itself in an orbit where a visiting vehicle can be docked thereto via docking mechanism  20 . In this use, since space module  10  must be able to maintain its proper orientation in orbit and/or position itself for a docking operation, an attitude control  24  is coupled to outer structure  12 . Attitude control  24  can be realize by any of a variety of reaction-based (e.g., thrusters) designs or momentum exchange devices (e.g., control moment gyros, flywheels, etc.) Thus, the particular choice of attitude control  24  is not a limitation of the present invention.  
         [0023]    Space module  10  could also incorporate other devices/systems to support its functions. For example, power for space module  10  could be beamed thereto via space or collected by solar photovoltaic arrays (not shown) extending out therefrom. Storage of such power could be achieved using batteries. Another option for power storage is to use flywheels if attitude control  24  is a flywheel-based system. Another device/system not illustrated for sake of clarity is the plumbing and management system used to couple the contents of each of stores  14  and  16  to transfer interface  18 . Other devices/systems that would typically be included onboard space module  10  include communications systems, navigation systems which would be coupled to attitude control  24 , and other sensors/computers typically found onboard spacecraft.  
         [0024]    Referring now to FIG. 3, a more specific embodiment of the present invention&#39;s space module  10  is shown and is illustrated generally by numeral  50 . Space module  50  is sized and constructed for transport in and deployment from the cargo bay of a Shuttle-class RLV, and is further constructed to maintain its structural integrity when coupled to space vehicles that impact axial and translational loads transferred to space module  50  during space travel. In addition, space module  50  has docking mechanisms  60 A and  60 B provided at opposing longitudinal ends thereof. In this way, a space vehicle (i.e., with engines) could be coupled to docking mechanism  60 A (where propellant transfer interface  58  terminates) while a payload (e.g., crew compartment, equipment, another one of space modules  50 , etc.) could be coupled to docking mechanism  60 B. For these reasons, the primary structure of space module  50  must be strong enough to support and protect its contents and provide a load transfer path for modules/vehicles coupled thereto.  
         [0025]    As shown in FIG. 4, the primary or skeletal frame of space module  50  is referenced generally by numeral  70 . Frame  70  defines two main cylindrical sections  70 B and  70 C and two tapered sections  70 A and  70 D disposed at opposing longitudinal ends of frame  70 . In general, each of these sections begins and ends with ring-shaped beams which are structurally coupled to one another by longerons (e.g., I-beams). More specifically, upper tapered frame section  70 A has a docking ring  71  (to which docking mechanism (DM)  60 B is coupled) and an end ring  72  coupled together by longerons  76 . Main section  70 B has end ring  72  and a mid-span ring  73  coupled together by longerons  77 . Main section  70 C has mid-span ring  73  and an end ring  74  coupled together by longerons  78 . Lower tapered section  70 D has end ring  74  and a docking ring  75  (to which docking mechanism (DM)  60 A is coupled) coupled together by longerons  79 . The longerons that span each section define the main axial load carrying structure. Therefore, the longerons are typically made from a very rigid material such as a magnesium metal matrix having long fiber carbon embedded therein, although it is to be understood that the choice of material is not a limitation of the present invention. Each of rings  71 - 75  can also be made from the same rigid material. Note that mid-span ring  73  can also serve as a connection point for trunion fittings (not shown) used by a Shuttle-class RLV. The tapering of each of sections  70 A and  70 D to the dimensions of docking rings  71  and  75 , respectively, improves axial and translational load transfer from vehicles/modules docked to space module  50 .  
         [0026]    Referring again to FIG. 3, a skin  80  is operatively connected to frame  70  to protect the contents of space module  50 . Because space module  50  will be relatively “parked” in orbits such as Low Earth Orbit, it will be subject to impact from micrometeriods and other forms of orbital space debris. Thus, skin  80  forms a shield that is generally a multi-layer structure that can be impacted without permitting an impacting projectile pass therethrough. Some of the layers could be selected to provide thermal protection if necessary. One example of skin  80  has layers of ceramic cloth interleaved with low density open-cell foam, all of which is supported on a multi-layer strength material (e.g., KEVLAR) wall that is attached directly to the longerons of frame  70 . Another example of skin  80  is the well-known Whipple debris shield with thermal insulation incorporated therein. Skin  80  could also be made from combinations of these or other technologies. Accordingly, it is to be understood that the particular choice of skin  80  is not a limitation of the present invention.  
         [0027]    The diameter of section  70 C is made slightly smaller than that of section  70 B in order to accommodate various stand-off devices/systems used to support space module  50 . By way of non-limiting examples, such support devices/systems may include radiator(s)  82 , photovoltaic array(s)  84  which are collapsed against skin  80  during transport and would typically extend out from space module  50  when it is deployed in orbit, and panel(s)  86  for accessing computers, data recorders, and communications hardware (not shown for clarity of illustration) maintained onboard space module  50 .  
         [0028]    Space module  50  is illustrated with three storage tanks for storing liquid hydrogen fuel, liquid oxygen oxidizer, and liquid xenon. Specifically, tank  52  stores liquid hydrogen, tank  54  stores liquid oxygen, and tank  56  stores liquid xenon. Although not illustrated in FIG. 3 for sake of clarity, each of tanks  52 ,  54 , and  56  would be coupled to a cryogenic cooling system (e.g., cryogenic cooling system  22 ) in order to prevent boil-off during long periods of storage. Also not illustrated for sake of clarity are the attitude control system and any plumbing system used to couple each of tanks  52 ,  54 , and  56  to transfer interface  58 . Because the density of liquid oxygen and liquid xenon is considerably greater than that of liquid hydrogen, tanks  54  and  56  are positioned closer to transfer interface  58  than tank  52 . This positioning scheme places the center of gravity of space module  50  as close as possible to the space vehicle (not shown) that will be coupled to docking mechanism  60 A and that will be burning one of the liquid hydrogen/liquid oxygen or liquid xenon propellants.  
         [0029]    In operation, the space module of the present invention is filled with each of its propellants by means of its bi-directional propellant transfer interface (e.g., interface  18  or  58 ) and then deployed in space by means of either a launch vehicle (e.g., reusable or expendable) coupled to the module that uses one of its propellant stores, or by means of an RLV (or an expendable launch vehicle) that ferries the fully-loaded space module as a piece of cargo. It is to be understood that, depending on the particular application, the space module could be partially filled with one or more of its propellants. Further, other applications may exist in which the space module is deployed in an empty state, i.e., no propellants. Regardless of the amount of propellant stored or type of deployment vehicle used, once the space module is deployed in an orbit, its onboard attitude control (e.g., attitude control  24 ) maintains the proper attitude of the space module for docking. Onboard position determination systems (e.g., a GPS-based system) may be used to position the space module (using its attitude control) for docking with a space vehicle. Once coupled to a space vehicle, one of the chemically initiated or electrically initiated propellants is supplied to the space vehicle via the space module&#39;s transfer interface  18 . Individual units of the space module  10  could be used to re-supply propellant to geostationary satellites or space stations. They could also be used to supply both types of propellant for one-way mission vehicles. They further could be used as propellant depots to which small space vehicles come to re-fuel.  
         [0030]    In addition to its stand alone utility, the space module  10  of the present invention can be used as an integral element of an in-space propellant supply architecture, one example of which is illustrated in FIG. 5. Two fully loaded space modules  50 A and  50 B (each of which is similar to space module  50 ) are deployed about the Earth  100  in a Low Earth Orbit  102 . The fully loaded nature of the space modules is indicated in FIG. 5 by tanks  52 ,  54 , and  56  being fully cross-hatched. The space modules  50 A,  50 B could also be initially stationed at various locations in space other than Low Earth Orbit  102 . Accordingly, it is to be understood that the particular choice of deployment location is not a limitation of the present invention.  
         [0031]    A manned space vehicle  90  is coupled to space module  50 A and an unmanned space vehicle  92  is coupled to space module  50 B. Manned space vehicle  90  uses the high thrust, chemically initiated propellant in tanks  52 ,  54  in order to propel itself toward an outer space destination as quickly as possible. For example, when space vehicle  90  reaches a destination  104 ′, the fuel and oxidizer in tanks  52  and  54 , respectively, are substantially depleted as indicated by the reduced cross-hatched portion thereof at destination  104 ′. However, note that the electrically initiated propellant in tank  56  is still at full capacity. In contrast, unmanned space vehicle  92  uses the low thrust, electrically initiated propellant in tank  56  in order to propel itself toward approximately the same destination as space vehicle  90 . Thus, when space vehicle  92  reaches destination  104 ′, the propellant in tank  56  is substantially depleted as indicated by the reduced cross-hatched portion thereof at destination  104 ′. However, the chemically initiated propellants in tanks  52  and  54  are still at full capacity. Accordingly, when it is time for manned space vehicle  90  to return to orbit  102 , space module  50 A is released from space vehicle  90 , and space module  50 B is released from space vehicle  92 . Then, space vehicle  90  re-docks to space module  50 B (which is fully loaded with chemically-initiated propellant), and space vehicle  92  re-docks with space module  50 A (which is fully loaded with electrically initiated propellant). The switching of space modules between space vehicles typically occurs in the vicinity of original destination  104 ′. For purpose of illustration, the switching of space modules/vehicles is complete at a destination  104 ″ which is generally&#39; going to be in the vicinity of destination  104 ′. Manned space vehicle  90  can then quickly return to orbit  102  using the high thrust provided by the full load of chemically initiated propellant onboard space module  50 B, while space vehicle  92  returns at a slower pace using the full load of electrically initiated propellant onboard space module  50 A.  
         [0032]    The above-described, in-space propellant supply architecture can save a significant amount of chemically initiated propellant when compared to a manned space vehicle that must be configured with enough propellant to get its crew to a destination in space and then return its crew from the destination in space, primarily because there is a substantial mass penalty for carrying enough propellant to support transport for a round trip. In contrast, the propellant supply architecture of the present invention can utilize smaller space modules because the high thrust propellant is designed for a one-way trip.  
         [0033]    The space module  10  provides a reusable means of providing propellant for in-space operations such as re-fueling of orbiting vehicles, providing manned and unmanned space vehicles with both high thrust and low thrust propellant supplies, and warehousing of propellants for long periods of time. Each space module can be re-filled with each type of propellant at a depot maintained, for example, in a Low Earth Orbit. Use of the space module in a propellant supply architecture as described herein will greatly reduce the mass penalty associated with existing manned space travel that use chemically initiated propellants.  
         [0034]    Although only a few exemplary embodiments of this invention have been described in detail above, those skilled in the art will readily appreciate that many modifications are possible in the exemplary embodiments without materially departing from the novel teachings and advantages of this invention. Accordingly, all such modifications are intended to be included within the scope of this invention as defined in the following claims. In the claims, means-plus-function and step-plus-function clauses are intended to cover the structures or acts described herein as performing the recited function and not only structural equivalents, but also equivalent structures. Thus, although a nail and a screw may not be structural equivalents in that a nail employs a cylindrical surface to secure wooden parts together, whereas a screw employs a helical surface, in the environment of fastening wooden parts, a nail and a screw may be equivalent structures.  
         [0035]    What is claimed as new and desired to be secured by Letters Patent of the United States is: