Abstract:
A demountable docking interface mechanism and method for docking and undocking two spacecraft utilizing the demountable docking interface mechanism.

Description:
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates generally to docking mechanisms for spacecraft. 
     2. Description of the Related Art 
     The ability to dock multiple spacecraft has an enormous benefit for many spacecraft applications. Spacecraft that do not fit in a single launch vehicle can be launched independently and assembled in orbit. Furthermore, for instance, a broken module of a spacecraft can be jettisoned and replaced by a new module without having to replace the entire spacecraft. 
     A current limitation of all existing docking interface designs is the limited reliability and high complexity of their designs. In particular, the current Orbital Express docking interface design (Starsys Research Corporation, Boulder, Colo.) uses motors, gears, and cams to reach mechanical connection loads that are high enough to transfer fluid. Additional moving parts may decrease reliability while increasing the mass of the design. Reliability is a primary concern of all spacecraft missions. There is not a chance to repair the docking interface after the spacecraft has been launched. 
     Other docking interface designs such as Michigan Institute of Technology&#39;s Space System Lab and Michigan Aerospace Corporation do not allow for the transfer of fluid. Further, none of the current docking interface designs are scalable down to a CubeSat size spacecraft (10 cm×10 cm×10 cm). 
     A CubeSat is a spacecraft with dimensions of 10×10×10 centimeters (i.e., a volume of one liter), weighing no more than one kilogram, and typically using commercial off-the-shelf electronics components. CubeSats provide a mean for universities, companies and other organizations throughout the world to enter the realm of space science and exploration. Most CubeSats carry one or two scientific instruments as their primary mission payload. Miniaturized spacecrafts have application in low data rate communications systems, gathering data from multiple points and in-orbit inspection of larger spacecrafts. 
     SUMMARY OF THE INVENTION 
     Embodiments in accordance with the invention include a demountable docking interface mechanism and method for docking and undocking two spacecraft utilizing the demountable docking interface mechanism. Embodiments in accordance with the invention are best understood by reference to the following detailed description when read in conjunction with the accompanying drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a side view block diagram illustrating a first docking interface mechanism attached to a first spacecraft and demountably coupled to a second docking interface mechanism attached to a second spacecraft in accordance with one embodiment. 
         FIG. 2  is a perspective view of one of the docking interface mechanisms of  FIG. 1  in accordance with one embodiment. 
         FIG. 3  is a cutaway perspective view of an alignment pin included in the docking interface mechanism of  FIG. 2  in accordance with one embodiment. 
         FIG. 4  is a cross sectional view of the alignment pin included in the docking interface mechanism of  FIG. 2  in accordance with one embodiment. 
         FIG. 5  is a partial cross sectional view showing the shape memory alloy coupling of an alignment pin of  FIG. 2  approaching a cup pipe within a corresponding alignment cup in accordance with one embodiment. 
         FIG. 6  is a partial cross sectional view showing the shape memory alloy coupling of the alignment pin of  FIG. 5  in a soft mate configuration with the cup pipe within the corresponding alignment cup of  FIG. 5  in accordance with one embodiment. 
         FIG. 7  is a partial cross sectional view showing the shape memory alloy coupling of the alignment pin of  FIG. 6  in a hard mate configuration with the cup pipe within the corresponding alignment cup of  FIG. 6  in accordance with one embodiment. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIG. 1  is a side view block diagram illustrating a first docking interface mechanism  10  attached to a first spacecraft  12  and demountably coupled to a second docking interface mechanism  10 ′ attached to a second spacecraft  14  in accordance with one embodiment. In one embodiment, first docking interface mechanism  10  is a permanent component of first spacecraft  12 , i.e., permanently formed in or on first spacecraft  12 . In another embodiment, first docking interface mechanism  10  is detachable from first spacecraft  12 , allowing reuse of first docking interface mechanism  10 . In one embodiment, second docking interface mechanism  10 ′ is a permanent component of second spacecraft  14 , i.e., permanently formed in or on second spacecraft  14 . In another embodiment, second docking interface mechanism  10 ′ is detachable from second spacecraft  14 , allowing reuse of second docking interface mechanism  10 ′. 
     For purposes of description, in one embodiment, first docking interface mechanism  10  and second docking interface mechanism  10 ′ are substantially identical, and are used to demountably couple first spacecraft  12  and second spacecraft  14 . Herein the use of the prime indicator symbol (′) is used to differentiate substantially identical structures of second docking interface mechanism  10 ′ from those of first docking interface mechanism  10 . 
     Referring to  FIG. 1 , in one embodiment, a first alignment pin  16  extends from first docking interface mechanism  10  into a first alignment cup  20 ′ of second docking interface mechanism  10 ′; and, a first alignment pin  16 ′ extends from second docking interface mechanism  10 ′ into a first alignment cup  20  of first docking interface mechanism  10 . Additionally, although not shown in the side view of  FIG. 1 , in one embodiment, a second alignment pin  18  extends from first docking interface mechanism  10  into a second alignment cup  22 ′ of second docking interface mechanism  10 ′; and, a second alignment pin  18 ′ extends from second docking interface mechanism  10 ′ into a second alignment cup  22  of first docking interface mechanism  10 . First docking interface mechanism  10  is further described herein in detail, and also simply referred to as docking interface mechanism  10 . Those of skill in the art can understand that the descriptions of docking interface mechanism  10  are also applicable to second docking interface mechanism  10 ′. 
     Herein the term spacecraft refers to any vehicle or object intended for transit through, or outside earth&#39;s atmosphere. Further herein the term space refers to the area outside Earth&#39;s atmosphere, unless otherwise specified. 
       FIG. 2  is a perspective view of docking interface mechanism  10  in accordance with one embodiment. In one embodiment, docking interface mechanism  10  includes a frame  24  in which first alignment cup  20  and second alignment cup  22  are formed. In one embodiment first alignment cup  20  includes an alignment cup pipe  26 ; and second alignment cup  22  includes an alignment cup pipe  28 . In one embodiment, first alignment cup  20  and second alignment cup  22  are substantially identical. 
     Frame  24  further includes first alignment pin  16  and second alignment pin  18 , a first data and power connector  34  and a second data and power connector  36 , and four thrusters, respectively, thrusters  30 ,  32 ,  38 , and  40 . In one embodiment first alignment pin  16  and second alignment pin  18  are substantially identical. In one embodiment, connectors  34  and  36  are used to pass electrical power and electrical data signals between first spacecraft  12  and second spacecraft  14  when docked. In one embodiment, frame  24  also includes one electromagnet having a first pole  44  and a second pole  46 . The electromagnet with poles  44 ,  46  provide active assistance in close proximity operation alignment. 
     Referring to  FIGS. 1 and 2 , first alignment cup  20  and second alignment cup  22  of first docking interface mechanism  10  assist in guiding corresponding first alignment pin  16 ′ and second alignment pin  18 ′ of second docking interface mechanism  10 ′ into position for docking, e.g., coupling, as first spacecraft  12  and second spacecraft  14  approach one another. Similarly, first alignment cup  20 ′ and second alignment cup  22 ′ of second docking interface mechanism  10 ′ assist in guiding corresponding first alignment pin  16  and second alignment pin  18  of first docking interface mechanism  10  into position for docking. 
     The alignment pins  16 ,  18  of docking interface mechanism  10  are further described herein with reference to  FIGS. 3-7 . In the present illustrations, as alignment pins  16 ,  16 ′,  18 ,  18 ′, are, in one embodiment, substantially identical, only alignment pin  16  is described. 
       FIG. 3  is a cutaway perspective view of first alignment pin  16  included in docking interface mechanism  10  in accordance with one embodiment.  FIG. 4  is a cross sectional view of alignment pin  16  in accordance with one embodiment. Referring now to  FIGS. 3 and 4  together, in one embodiment, first alignment pin  16  includes a sleeve  50  that has a cylindrical portion  52  and a tapered outer end portion  54 . In one embodiment, end portion  54  is frustoconical in shape. A shape memory alloy (SMA) coupling  56  is located inside sleeve  50  with a heater  58  being arranged to heat SMA coupling  56 . An alignment pin pipe  60  is located inside sleeve  50  to provide means for transferring fluids between first spacecraft  12  and second spacecraft  14  as well as a structural load path. 
     In one embodiment, SMA coupling  56  is formed of a nickel-titanium alloy (NiTi) which exists in two crystalline phases known as martensite and austenite. The NiTi alloy is in a martensite phase at temperatures at or below an associated martensitic temperature, i.e., a martensitic temperature associated with the NiTi alloy. Heating the NiTi alloy to an austenitic temperature, i.e., an austenitic temperature of the NiTi alloy, causes a phase change from the martensite phase to an austenite phase. The dimensions of SMA coupling  56  decrease during the phase change from martensite to austenite. 
     Operation of spacecraft docking with docking interface mechanism  10  is described with reference to  FIGS. 5-7 . For clarity of description spacecraft docking is described with reference to first alignment pin  16  of first spacecraft  12  and first alignment cup  20 ′ of second spacecraft  14 . Those of skill in the art can understand that the description is applicable to the remaining alignment pins  16 ′,  18 ,  18 ′ and alignment cups  20 ,  22 ′,  22  earlier described. 
       FIGS. 5 ,  6 , and  7  notionally represent the coupling procedure, e.g., docking procedure, of docking interface mechanisms  10 ,  10 ′ in accordance with one embodiment.  FIG. 5  is a partial cross sectional view of an alignment cup pipe  26 ′ internal to first alignment cup  20 ′ of second docking interface mechanism  10 ′ approaching SMA coupling  56  and an alignment pin pipe  60  of alignment pin  16  of first docking interface mechanism  10  in accordance with one embodiment. 
     Referring to  FIGS. 1-5 , first spacecraft  12  and/or second spacecraft  14  maneuvers into a docking position using integrated thrusters, such as integrated thrusters  30 ,  32 ,  38 ,  40  of first docking interface mechanism  10 , and/or integrated thrusters  30 ′,  32 ′,  38 ′,  40 ′ of second docking interface mechanism  10 ′. Alignment pins  16 ,  16 ′,  18 ,  18 ′ and alignment cups  20 ,  20 ′,  22 ,  22 ′, earlier described with reference to first docking interface mechanism  10  and second docking interface mechanism  10 ′, assist in guiding the fine alignment process. 
       FIG. 6  is a partial cross sectional view showing first docking interface mechanism  10  in a soft mate configuration with second docking interface mechanism  10 ′ in accordance with one embodiment. In  FIG. 6 , first alignment pin  16  of first docking interface mechanism  10  is inserted into first alignment cup  20 ′ of second docking interface mechanism  10 ′ in a soft mate condition. In one embodiment, electromagnet poles  44 ,  46  of first docking interface mechanism  10  and/or electromagnet poles  44 ′,  46 ′ of second docking interface mechanism  10 ′ are activated to assist in alignment. Herein in one embodiment a soft mate condition occurs when insertion of an alignment pin, e.g., alignment pin  16 ,  16 ′,  18 ,  18 ′, into an alignment cup, e.g., alignment cup  20 ,  20 ′,  22 ,  22 ′ is completed and the associated SMA coupling, e.g., SMA coupling  56 , is below an associated austenitic temperature. 
       FIG. 7  is a partial cross sectional view showing first docking interface mechanism  10  in a hard mate configuration with second docking interface mechanism  10 ′ in accordance with one embodiment. Herein in one embodiment a hard mate condition occurs when an SMA coupling of an alignment pin, such as SMA coupling  56  of first alignment pin  16  is inserted within an alignment cup, such as first alignment cup  20 ′, is heated to an associated austenitic temperature forming a rigid connection so that data, power and fluids may be transferred into or out of the spacecraft, for example, via the transfer path created by connection of alignment cup pipe  26 ′ with alignment pin pipe  60 . 
     In one embodiment, power is applied to heating element  58  surrounding SMA coupling  56  to heat SMA coupling  56  to an associated austenitic temperature, so that SMA coupling  56  contracts to form a rigid mechanical connection and sealed fluid connection between first spacecraft  12  and second spacecraft  14 . In one embodiment, the heat is generated by heating element  58 . Generally viewed, the contraction of SMA coupling  56  results in SMA coupling  56  clamping down on alignment cup pipe  26 ′ to form the rigid connection. 
     While docked, e.g., after the rigid mechanical connection is completed, in one embodiment, first spacecraft  12  and second spacecraft  14  have the capability to share data, power, and fluid, which can be used for propulsion. 
     In one embodiment, SMA coupling  56  is heated with redundant resistive heating elements, such as heating element  58 . In one embodiment, sleeve  50  encases SMA coupling  56  to minimize the geometry required on the shape memory alloy. In one embodiment, alignment pin pipe  60  is formed of a material having a similar coefficient of thermal expansion as SMA coupling  56 . In one embodiment, the material of alignment pin pipe  60  has a much higher hardness than the alloy of SMA coupling  56  for reusability. In one embodiment, alignment pin pipe  60  is formed of the material stainless steel 440C. 
     Although not illustrated, it can be understood by those of skill in the art that in undocking, SMA coupling  56  releases the mechanical connection when the temperature of SMA coupling  56  decreases to the associated martensitic temperature. In one embodiment, this associated martensitic temperature can be reached by using a Peltier-effect cooler. Thus in some embodiments, docking interface mechanism  10  further includes a cooling element, such as a Peltier-effect cooling element. 
     As described herein embodiments in accordance with the invention provide a docking interface mechanism for demountably coupling a first spacecraft with a second spacecraft. In some embodiments, the docking interface mechanism provides spacecraft active/passive assist in the close proximity operations of docking by adjusting the attitude of a spacecraft. In some embodiments, the docking interface mechanism reduces spacecraft maneuvering. In some embodiments, the docking interface mechanism provides connections for transferring electrical power and for transmitting control and data signals between spacecraft. 
     In some embodiments, the docking interface mechanism provides fluid (liquid/gas) transfer at pressures up to 3000 psi. In some embodiments, the docking interface mechanism does not require electrical power to maintain a mechanical connection in the docking process. In some embodiments, the docking interface mechanism includes integrated thrusters. In some embodiments, the docking interface mechanism includes integrated light emitting diodes and/or cameras. In some embodiments, the docking interface mechanism is scalable down to a 10 cm×10 cm footprint. In some embodiments, the docking interface mechanism utilizes less than three (3) Watts of power for actuation. In some embodiments, the docking interface mechanism is reusable for multiple docking/undocking cycles. 
     Herein although embodiments in accordance with the invention have been described with reference to spacecraft, the invention is not limited to spacecraft in orbit. Embodiments in accordance with the invention are applicable in separating two space vehicles after launch and may be used in any application where a releasable connection is made between two structures, including terrestrial, aerial and underwater applications.