Abstract:
A solar array according to this invention includes a solar blanket which is folded into adjacent panels that are hinged together in an accordion-folded mode at parallel hinges. A pair of foldable spines is fixed to the panels and runs the length of the array. The hinges are included in the spine. The spines are mounted at one end to a base plate, and also at the other end to a tip plate. The base plate is intended to be attached to the spacecraft structure with a yoke which will provide required standoff for rotational clearance during sun-tracking. 
     A pantograph deployment structure extends between the two plates. It can be retracted for storage and extended to deploy the blanket. A conductive harness is attached to the blanket to collect current from the panels.

Description:
This application is a continuation of U.S. application Ser. No. 09/436,435, now abandoned filed Nov. 8, 1999, and a continuation-in-part of U.S. application Ser. No. 09/400,665, now abandoned filed Sep. 20, 1999. 
    
    
     FIELD OF THE INVENTION 
     A solar array to power satellite vehicles which is stowable in a small volume for shipment and launch, and that is deployable when aloft to expose a large surface area of solar collectors. 
     BACKGROUND OF THE INVENTION 
     Deployable solar arrays are typically contained in a small envelope when their space vehicle is launched. They are later deployed to an extended configuration to expose areas of solar collectors. Examples of such arrays are shown in the following United States patents: 
     Avilov U.S. Pat. No. 3,460,992 
     Harvey et al U.S. Pat. No. 5,296,044 
     Everman et al U.S. Pat. No. 5,487,791 
     A review of these patents will disclose remarkable efforts to reduce the weight and increase the reliability of these arrays. Cost, while important, has been and still is subordinate to reliability. The failure of an array to deploy and to survive for its full design life can result in loss of value of the entire craft and its payload. The cost of the craft and its payload is many times that of the array, especially when the payload is unique and employed for very advanced applications. 
     Because of this, and because of the relatively small number of vehicles involved, the design and manufacture of solar panels and their supporting structure has tended toward the complex, familiar, and costly. They have been carefully and slowly built, almost in a “handicraft” sense. 
     However, with the advent of space-based communication systems, the market for satellites has greatly enlarged. The cost of the payloads, while still considerable, has decreased. Expenses which are tolerable for a few very high value vehicles become unacceptable when the production count will run into the hundreds. 
     The demand for such a large number of arrays threatens to outstrip the capacity of existing manufacturing plants that were sufficient for the previous slow-paced demand. Multiplying plant capacity can permit faster production schedules, provided that additional skilled personnel can be found, and provided that the additional capital is available. Still the arrays would remain at least as costly. 
     Also, the problems of producibility remain. Existing constructions are built very painstakingly, because if one part is imperfectly produced, a large part of the entire array often must be scrapped or reworked at considerable cost. This risk and the unfavorable consequences which inevitably occur, has reduced the yield of these arrays. 
     It is the primary object of this invention to provide a solar panel which can be efficiently manufactured to high standards, and should some part of it be unsuitable, can be quickly and easily repaired or replaced. Thus the entire assembly need no longer be hostage to the acceptability of every part. Instead, all parts will be individually and readily replaceable. 
     The key subassembly which has the greatest vulnerability during integration and test, the highest value, and the longest repair time is the solar cell stringing. This invention proposes to separate the field of solar cells, traditionally arrayed on blanket systems, into separate but identical modules termed SPMs (solar power modules). This will allow mass production at low cost and availability for change out during test as required. Rapid integration is achieved by integrating each SPM into the “blanket” with quick attachment using a minimum number of fastening elements as described herein. The SPMs are hung onto parallel straps of thin (flexible to bending) “spines” spaced to overlay the outside edges of the SPMs. The flexible spines carry all the individual SPMs. The rows of SPMs can be folded to unfold or stow the blanket as an accordion by utilizing the flexibility of the spine straps. 
     With this invention, it appears likely that an array which formerly required a few months to build, can be built in a day. 
     However, manufacturing problems are not the only ones solved by this invention. In order to build a truly lightweight structure, the materials of construction must themselves be lightweight, and will often lack much structural strength while in a gravitational field, or in the fields of force that exist at launch or in transporting it to the site where it will be installed. To overcome this, conventional arrays simply provided more strength with more structure, and more cost. 
     These arrays must be stored in such a way that they can be handled on the ground without extreme care, and which will protect the array from the large launching forces. Then, when the craft is in orbit, the delicate solar blanket must controllably be deployed from the craft and be fully protected during extension to the deployed configuration. This still does not exhaust the problems of most existing arrays. Their tendency is to deploy the structure and lock it physically into a rigid structure. While there are no substantial acceleration forces on it while in orbit, a solar array can be subject to substantial internal forces, for example those which occur during the time while the craft leaves the shadow of the earth and comes into the sunlight. All too often, the different local expansions of material result in a physical snap as relative dimensions of the various parts change with rapidly changing temperature. Such forces can be damaging to the delicate parts of an array, and are a particular nuisance to proper orientation control of the spacecraft. The guide bending or “snapping” of this deployed structure causes a shift of mass centers and the spacecraft oscillates about the intended orientation until the “snap” has been damped out. 
     In the course of simplifying the array of this invention, the applicants have taken a different approach to cause and to assure deployment, which eliminates risk of the snapping action which is experienced in much of the prior art. 
     The consequence of these improvements is to increase the productivity and yield of solar arrays while still providing excellent reliability, at a significantly lower cost. 
     BRIEF DESCRIPTION OF THE INVENTION 
     A solar array according to this invention includes a solar blanket which is folded into adjacent panels that are hinged together in an accordion-folded mode at parallel hinges. A pair of foldable spines is fixed to the panels and runs the length of the array. The hinges are included in the spine. The spines are mounted at one end to a base plate, and also at the other end to a tip plate. The base plate is intended to be attached to the spacecraft structure with a yoke which will provide required standoff for rotational clearance during sun-tracking. 
     A pantograph deployment structure extends between the two plates. It can be retracted for storage and extended to deploy the blanket. A conductive harness is attached to the blanket to collect current from the panels. 
     According to a preferred but optional feature of the invention, the deployable structure lies in a plane parallel to the blanket. It includes a plurality of scissor links, and adjacent to each of the plates, a synchronizing link which compels the scissor arms to deploy in the correct direction. 
     According to another preferred but optional feature of the invention, tensioning springs are contained in at least some of the scissor arms, which tension a cable that tends to pull intersections of opposite scissor links toward one another, thereby to deploy the structure. Deployment is resisted by a lanyard which extends between the plates. When released, the rate at which the tip plate can separate from the base plate and deploy the blanket is limited by limiting the rate of release of lanyard from a damper. Advantageously the lanyard can pass through eyelets on the blanket panels to hold the blanket in its proper position relative to the deployment structure. 
     According to yet another preferred but optional feature of the invention a kicker spring may be placed at the intersections of the scissor linkages, biasing them toward the deploying direction without imparting shear forces and hence without adding drag. 
     According to still another preferred but optional feature of the invention, a plurality of tie-down rods are fixed to the base panel and pass through the tip panel. A releasable fastener holds the stored, accordion-folded, blanket until deployment is desired. Release of the fastener enables the array to deploy at a rate limited by the damper. 
     The above and other features of this invention will be fully understood from the following detailed description and the accompanying drawings, in which: 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIGS. 1-4 are schematic showing of the deployment structure to illustrate its motions between the stored (FIG. 1) and the deployed (FIG. 4) condition; 
     FIG. 5 is a perspective view of the deployment structure in greater detail, in an intermediate configuration; 
     FIG. 6 is a fragmentary view showing kicker spring; 
     FIG. 7 is a fragmentary view showing a cable used to exert a deployment force on the structure; 
     FIG. 8 is a perspective view, partly in schematic notation, showing the deployed blanket with a schematic, deployed structure; 
     FIGS. 9-10 are schematic showings of the synchronizing link from the stowed configuration (FIG. 9) to the deployed configuration (FIG.  11 ); 
     FIG. 12 is a perspective view of the entire array in an intermediate configuration; 
     FIGS. 13-15 are schematic illustrations showing the harness from the displayed condition (FIG. 13) to the stowed condition (FIG.  15 ); 
     FIG. 16 is a fragmentary view showing a negator spring connecting the blanket to the tip plate; and 
     FIG. 17 is a fragmentary cross section showing a reliable feature holding the array in its stored condition. 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Solar array  20  is shown in FIGS. 8 and 12. It includes a base plate  21  and a tip plate  22 . These plates are joined by a pantograph deployment structure  25 . In FIG. 12 the structure is shown partially deployed. In FIG. 8 it is shown deployed. 
     As best shown in FIGS. 1 and 4, the deployment structure has synchronized link  26 , adjacent to the base plate, and synchronized link  27  adjacent to the tip plate. 
     Synchronizer arms  30 ,  31  are respectively mounted to hinges  28 ,  29  on the base plate. Synchronizer arm  32 ,  33  are similarly connected to link  27 . The precise construction of the synchronizer links will be disclosed later. The consequence of this arrangement is that the position of the synchronizer arms is unique for all extensions, and that the deployment will occur nearly linearly. 
     Two pairs of scissor links are provided: pairs  40  and  41 . Their ends are hinged to an adjacent arms so they rotate in the plane of FIGS. 3 and 4. These hinged joints are shown at  42 ,  43 ,  44 ,  45 ,  46  and  47 . 
     Scissor arms  50 ,  51  are pivotally joined at scissor joint  52 . Scissor arms  53  and  54  are pivotally joined at scissor joint  55 . 
     Kicker springs  60 ,  61  are placed in scissor joints  52  and  55 . Scissor joint  52  is shown in FIG.  6 . It includes a pivoted central arm  63  with compression springs  64 ,  65  between it and the adjacent scissor arm, tending to open the angle in which they are placed. This provides an initial force tending to open the outside of the scissor joint when the array is to be deployed. 
     Force for deployment of the deployer mechanism is by springs contained inside the scissor arms that exert a pulling effort on a cable which passes around one scissor joint and is fixed to the opposite scissor joint. These exert a springing continuous bias force tending to draw the ends of the scissor arms together and deploy the array. One such spring  70  is shown in FIG. 7, placed in scissor arm  70 . Cable  71  is bent around a cylinder  72  and extends toward the opposite hinge. Cables  73 ,  74 ,  75 ,  76 , and  77  are shown, which all perform the identical function at their respective hinge joint. 
     This prevailing spring force is resisted by a pair of lanyards  80 ,  81 . These lanyards are wound on a rotational damper  82  (FIG.  12 ). This damper is mounted on the tip plate. The lanyards pass through ports in the plate to extend to and connect to the base plate. When the array is stowed, the lanyards will be wound on the damper spool and the plates will be held to one another. When the tip plate is released from the base plate, the damper will limit the rate at which the lanyard will be released, and the rate of deployment will thereby be regulated. Appropriate selection of the damper and selection of spring rates will establish the rate of deployment. The damper does not exert substantial torque at very low speed, but it effectively limits the rate of extension. 
     To maintain the blanket in a disciplined position during deployment, the lanyards are passed through eyelets  90  on the edge of some of the panels located at the adjacent folds of the blanket. 
     The synchronizing links are schematically shown in FIGS. 9-11. It is hinged to synchronizing arms  30  and  31 . Notice, however, that as a rigid link it extends to contact these arms at equal spacings  120  and  121  from hinges  28  and  29 , but on opposite sides of the hinges. Thus, it acts as a crank for each of these arms. A rotation of one of the synchronizing arms will result in an equal and opposite rotation of the other. Therefore neither synchronizing arm can tilt farther from or nearer to the perpendicular bisector, and extension of the deployment structure along the central axis is assured. 
     In storage, the stowed blanket  100  (FIG. 17) is held between the base plate and the tip plate by a plurality of bolts  101  with releasable nuts  102 . Releasable fasteners are well known in the fastener art, and are available with a number of release schemes. While ordinance- actuated nuts are useful, there is a preferred class in which actuation is the result of change of shape caused by heating some part of the nut. 
     In addition to the hold-down bolts, alignment bolts or pins, and corner braces, may be provided to resist side or shear forces on the stowed array. 
     This completes the description of the deployment structure. This structure provides the support and control for the blanket  100  itself. The blanket is not directly connected to the deployment structure  25 . Instead, whether accordion-folded in the stowed position or extended in the blanket, it is located to the side of the deployment structure, connected to it only through the base plate and the tip plate, to which both the blanket and the structure are attached. 
     Attention is called to the fact that the blanket is not rigidly tensioned between the plates when deployed. Instead it is maintained deployed by a spring system which exerts a steady force while the blanket length changes due to thermal extremes. The pantograph structure is designed for near-zero thermal expansion. Since the blanket loading is constant, the structure will not undergo the snapping action which can occur in a rigid structure when it passes into and out of the earth&#39;s shadow. The unfavorable consequences of this and of any tolerance build-up are thereby obviated. 
     Blanket  100  is comprised of a substrate that comprises a semi-rigid backplane, solar cells on the backplane, and often a coverglass over the cells. The details of these modules are of no importance to the invention, and they will therefore not be described in full detail here. Suffice it to say that they structurally self-sufficient and planar, that the cells have circuitry and circuit connections through which electrical current pass is to be collected and utilized. Such arrangements are well-known. 
     For structural interconnection of the modules, a pair of longitudinal spines  102 ,  103  are attached to each of the modules. These spines extend from end to end of the blanket. End  104  is attached to the base plate. End  105  is attached to the tip plate through a constant force “negator” spring  106  (FIG.  16 ). Each spine is attached to each module at two locations for each module, one preferably being a rigid fixed rivet or other connector, and the other preferably permitting limited sliding between the module and the spine. 
     A hinge portion  110  of each spine is located between each adjacent module to provide for accordion-type folding between them. The spines are intended to be made from relatively thin, lightweight material, thin enough to serve as a hinge without reduction in thickness. If necessary, the spine can be thinned at the folding point to facilitate folding. 
     One or more harnesses  111 , as appropriate, are attached to the edge (or edges) of the blanket, to be connected to the cell circuitry. Preferably the harnesses are ribbon-like flexible conductor assemblies carrying a number of parallel conductors. The details of their circuit connections are of no importance to this invention. 
     It is, however, best practice not to over-stress the harness conductor material with too sharp a fold. For this purpose, as shown in FIG. 13, the harness is provided with extra length and assumes an S-like shape, being attached to the modules at each fold, and at a mid point of the modules. 
     Soft buttons  115  exemplified in FIG. 14, usually made of foam, will be placed on the modules at appropriate locations to hold the modules apart when in the stowed position. They will be placed between the folds to protect the cells, and if desired may also be placed between the backplanes. 
     The negator spring  106  is provided to reduce the effects on the blanket of sharp movements of the deployment structure. A negator spring has an essentially zero spring constant. In the illustrated structure it comprises an appropriate spring bent over a pulley  120  and connected to the respective lanyard and to the tip plate. Its function is to “absorb” any abrupt changes in length without substantial transmission of the exerted force. 
     In storage, the stowed stack  100  (FIG. 17) is held between the base plate and the tip plate by a plurality of bolts  101  with releasable nuts  102 . Release fasteners are well known in the fastener art, and are available with a number of release schemes. While ordinance-actuated nuts are useful, there is a preferred class in which actuation is the result of change of shape caused by heating some part of the nut. Such a nut is the preferred device because it does not apply a physical shock to the assembly. 
     This invention is not to be limited by the embodiments shown in the drawings and described in the description, which are given by way of example and not of limitations, but only in accordance with the scope of the appended claims.