Abstract:
The invention is a micro-satellite assembly. In detail, the invention includes first and second flat structural members containing the satellite payload. First and second tubular elements connect first and second structural members such that they are in a spaced relationship. A plurality of solar panels are movably to the tubular elements between the first and second structural elements, movable from a stored position between the structural elements to an deployed position external of these structural members. A mechanism is provided for biasing the plurality of the solar panels to the deployed position. A second mechanism is used to releasably secure the plurality of solar panels in the stored position.

Description:
RELATED APPLICATIONS 
     This application is a continuation-in-part of co-pending Provisional Patent Application Serial No. 60/145,164, “Multifunctional Structure Nano-Satellite”, filed Jul. 22, 1999. 
    
    
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The invention relates to the field of spacecraft, such as satellites or space probes and, in particular, to a miniature low cost spacecraft that can be packaged with other similar vehicles. 
     2. Description of Related Art 
     In an attempt to reduce the cost of space exploration and orbital earth sensor satellites, small single purpose satellites have been developed. However, there has been little effort to standardize the design of the buss, or payload carrying structure. U.S. Pat. No. 5,386,953 “Spacecraft Designs For Satellite Communication System” by J. R. Stuart discloses a hemispherical shaped satellite design using a tubular truss assembly to support a series of antennas that allows stacking of a series thereof. Another attempt also disclosed in the above referenced patent is to use an inflatable torus shaped structure. However, neither of these approaches is suitable for use with micro-sized satellites. First of all the truss assembly is a high cost structure for use with a really small satellite. The inflatable satellite requires a storable gas and a control system for inflation, again adding cost. 
     Thus, it is a primary object of the invention to provide a micro-satellite design. 
     It is another primary object of the invention to provide a micro-satellite design that can be stored in a very small volume. 
     It is a further object of the invention to provide a micro-satellite design that can be efficiently stacked such that a single booster can be used to launch a large number of them. 
     SUMMARY OF THE INVENTION 
     The invention is a micro-satellite assembly. In detail, the invention includes first and second flat structural members containing the satellite payload. Preferably, the first and second structural members are circular shaped having peripheral surfaces and equal diameters. First and second tubular elements connect the first and second structural members such that they are in a spaced relationship. Preferably, the tubular elements are connected at the peripheral surfaces and spaced 180 degrees apart. A plurality of solar panels are movably to the tubular elements between the first and second structural elements, movable from a stored position between the structural elements to a deployed position external of these structural members. Preferably there are four solar panels with a first pair mounted to the first tubular element and the second pair mounted to the second tubular element. Each individual solar panel of each pair is rotatable from the stored position to the deployed position in opposite directions. 
     A mechanism is provided for biasing the plurality of the solar panels to the deployed position, which includes a spring coiled about the tubular element having a first end attached thereto and a second end attached to the individual solar panel. A second mechanism is used to releasably secure the solar panels in the stored position. A third mechanism is included for releasably securing the solar panels in the deployed position. 
     The novel features which are believed to be characteristic of the invention, both as to its organization and method of operation, together with further objects and advantages thereof, will be better understood from the following description in connection with the accompanying drawings in which the presently preferred embodiment of the invention is illustrated by way of example. It is to be expressly understood, however, that the drawings are for purposes of illustration and description only and are not intended as a definition of the limits of the invention. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is a perspective view of the subject micro-satellite with the solar panels retracted. 
     FIG. 2 is a perspective view of the subject micro-satellite illustrated in FIG. 1 with the solar panels deployed. 
     FIG. 3 is a partial cross-sectional view of FIG. 1 taken along the line  3 — 3 , illustrating the biasing mechanism for the solar panels. 
     FIG. 4A is a partial cross-sectional view of FIG. 1 taken along the line  4 A— 4 A, illustrating the mechanism for releasably restraining the solar panels in the stored position. 
     FIG. 4B is an enlarged partial cross-sectional view of FIG. 1 taken along the line  4 B— 4 B illustrating the solar panel restraining system for locking the solar panels in the stored position. 
     FIG. 5 is an enlarged partial view of FIG. 2 illustrating the solar panel restraining system for locking the solar panels in the deployed position. 
     FIG. 6 is a side elevation view of a plurality of the subject micro-satellites stacked for mounting in a lunch booster. 
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENT 
     Referring to FIGS. 1-3,  4 A and  4 B, the micro-satellite, generally indicated by numeral  10  includes two circular structural members  12  and  14 , in which is mounted the payload  16  of the satellite. The two structural members  12  and  14  have externally facing surfaces  16 A and  18 A and inward facing surfaces  16 B and  18 B and are separated by a pair of tubular elements  20  and  22  connected to the peripheral surfaces  24  and  26  of the members  12  and  14 . Each structural member would incorporate a propulsion ring  27 A and  27 B providing station keeping control, for example multiple one shot jets. Solar Panels  24 A and  24 B are pivotally connected to the tubular element  20  and solar panels  26 A and  26 B are pivotally connected to tubular element  22  in a manner to be subsequently described. The solar panels all have a width indicated by numeral  27 . 
     The tubular element  20  includes cup shaped ends  28  and  30 , with holes  32  and  34 , respectively, therethrough and with the interior of the cup portions indicated by numerals  35  and  36 . A pin  37  having shoulders  38  and  40  and threaded shafts  42  and  44  that extend through the holes  32  and  34  and retained by nuts  46  and  48 . Thus the pin  37  space the two structural members  12  and  14  apart by a distance indicated by numeral  50 . Solar panels  24 A and  24 B include lugs  58  and  60  with holes  62  and  64  therethrough are rotatably mounted on the pin  37 . Lug  60  includes a spacer portion  65  having a width  66 . A spacer  67  having a width  68  slightly larger than the width  27  of the solar panels also having a hole  69  therethrough is mounted on the pin  37 . 
     A spring  70  is mounted in the cup portion  35  of the cup shaped end  28  and is wrapped about the pin  37  having a first end  72  engaged with hole  74  in the cup shaped end  28  and the second end  76  engaged with hole  78  in the lug  58  of the solar panel  24 A. A spring  80  is mounted in the cup portion  36  of the cup shaped end  30  and is wrapped about the pin  37  having a first end  82  engaged with hole  84  in the cup shaped end  30  and the second end  86  engaged with hole  88  in the lug  60  of the solar panel  24 B. Thus the springs  70  and  80  bias the solar panels  24 A and  24 B in opposite directions. 
     Tubular element  22  is similar to tubular element  20 . The tubular elements  22  therefore includes cup shaped ends  28 ′ and  30 ′, with holes  32 ′ and  34 ′, respectively, therethrough and with the interior of the cup portions indicated by numerals  35 ′ and  36 ′. A pin  37 ′ having shoulders  38 ′ and  40 ′ and threaded shafts  42 ′ and  44 ′ that extend through the holes  32 ′ and  34 ′ and retained by nuts  46 ′ and  48 ′. Thus the pin  37 ′ also space the two structural members  12  and  14  apart by a distance indicated by numeral  50 . Solar panels  26 A and  26 B include lugs  58 ′ and  60 ′ with holes  62 ′ and  64 ′ therethrough is rotatably mounted on the pin  37 ′. Lug  60 ′ includes a spacer portion  65 ′ having a width  66 ′. A spacer  67 ′ having a width  68 ′ slightly larger than the width  25 ′ of the solar panels also having a hole  69 ′ therethrough is mounted on the pin  37 ′. 
     A spring  70 ′ is mounted in the cup portion  35 ′ of the cup shaped end  28 ′ and is wrapped about the pin  37 ′ having a first end  72 ′ engaged with hole  74 ′ in the cup shaped end  28 ′ and the second end  76 ′ engaged with hole  78 ′ in the lug  58 ′ of the solar panel  26 A. A spring  80 ′ is mounted in the cup portion  36 ′ of the cup shaped end  30 ′ and is wrapped about the pin  37 ′ having a first end  82 ′ engaged with hole  84 ′ in the cup shaped end  30 ′ and the second end  86 ′ engaged with hole  88 ′ in the lug  60 ′ of the solar panel  26 B. Thus the springs 70 ′ and  80 ′ bias the solar panels  26 A and  26 B in opposite directions. With the solar panels  24 A, B and  26 A, B are offset from each other and can be overlapped with each other when in the stored position shown in FIG.  2 . 
     Referring particular to FIGS. 4A and 4B, the solar panels  24 A, B and  26 A, B are retained in the stored position shown in FIGS. 1 and 3 by means of retainers  98 A and  98 B mounted between the structural members  12  and  14 . The retainers  98 A, B include a rod  99  pivotally mounted at a forked first end  100 A to a pin  101  mounted in a groove  102  in structural member  14 . The second end  100 B of the rod  99  extends into a groove  103  in the structural element  12  and includes an indentation  104 . The rod  99  is retained by a solenoid operated pin puller device  105  having a pin  106  engaged with the indentation  104 . Thus when the device  105  is actuated by electrical current from a power source (not shown) the pin  106  is retracted from the indentation  104  in the second end  100 B of the rod  99 . At this point, the solar panels  24 A, B and  26 A, B, which as previously discussed are spring biased to the deployed position, can rotate the rods  99  about the pin  101  and to fully deploy as the rods are pushed clear of the satellite. Such devices are old in the art and need not be discussed in further detail. It should be obvious that other restraining systems could be used, for example ones using shape memory alloys. 
     When release of the solar panels is required, the devices  102  mounted in the structural member  12  are actuated, releasing the solar panels such that they can move the deployed position shown in FIG.  2 . Referring to Figure, to insure that the solar panels  24 A, B and  26 A, B once deployed, remain deployed, each have a spring biased retainer or détente  110  mounted in the lugs  58  and  60 , and  58 ′ and  60 ′. The détente  110  includes a hole  112  in the lugs  58 ,  60 ,  58 ′ and  60 ′ incorporating a pin  114  biased by a spring  116  on one side thereof. Each cup shaped end  28 ,  30 ,  28 ′ and  30 ′ include and indentation  118 . Thus as the springs  70  and  80  rotate the solar panels to the deployed positions, pin  114  become aligned with and engage the indentations  118  when fully the solar panels are deployed and become locked in place. 
     Referring to FIG. 6, it can be seen that when the micro-satellites have the solar panels  24 A, B and  26 A, B in the stored position, they can be easily stored one on top of each other in a small volume. Thus a launch booster can carry a large number into orbit. Furthermore, the solar panel deployment mechanism, retention mechanism for releasably retaining the solar panels in the stored and deployed positions may very from those illustrated. For example, the solar panel extension mechanism could employ shape memory alloy materials to accomplish deployment. 
     While the invention has been described with reference to a particular embodiment, it should be understood that the embodiment is merely illustrative, as there are numerous variations and modifications, which may be made by those skilled in the art. Thus, the invention is to be construed as being limited only by the spirit and scope of the appended claims. 
     INDUSTRIAL APPLICABILITY 
     The invention has applicability to the spacecraft industry.