Abstract:
A Deployable Morphing Modular Solar Array (DMMSA) for deploying Deployable Morphing Modular Solar Power Assemblies (DMMSPAs) from a spacecraft is provided. The DMMSA comprises a Root Staging and Deployment Mechanism (RSDM) mounted to a spacecraft. A plurality of petal assemblies are rotatably secured to the RSDM with each petal assembly having at least DMMSPA secured thereon and each DMMSPA having a slight V-configuration. A launch restraint assembly stacks and sandwiches the petal assemblies prior to deployment with the launch restraint assembly pre-loading each petal assembly&#39;s one or more DMMSPA into a substantially flat configuration. Upon release of the launch restraint assembly, the stacked and sandwiched petal assemblies rotate relative to the spacecraft and each petal assemblies DMMSPA elastically morphs from the substantially flat configuration into the slight V-configuration.

Description:
The present application claims benefit of priority of provisional patent application Ser. No. 61/402,520, filed on Aug. 31, 2010, entitled “Solar Array Wing”. 
    
    
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     This invention relates generally to a Deployable Morphing Modular Solar Array (DMMSA) and, more particularly, the invention relates to a subassembly of the solar array, the Deployable Morphing Modular Solar Power Assembly (DMMSPA) that makes the system modular, increases the deployed stiffness of the solar array, improves deployed first mode natural frequency of the system, and reduces overall manufacturing costs. 
     2. Description of the Prior Art 
     The current state-of-the-art (SOA) in solar arrays involves a highly customized design and testing effort for each spacecraft mission that does not use the significant design and production commonality existing among existing systems. The result of this approach is that each solar array is unique, costly, and long lead. In addition, this approach is contrary to what is required to support the commercial, Air Force and other agency needs, i.e., higher performance than current State of the Art (SOA), low cost inventory strategies of common components, rapid response to mission needs, and modular architecture that is semi-customizable and compatible across multiple missions. 
     Additionally, photovoltaic cell technology is evolving rapidly to the point that current solar array structural and mechanical systems do not optimize system level mass and volume performance potential. It is desirable to have a solar array that decreases production costs through modularity, significantly improves power to stowed volume ratio (W/m 3 ) and specific power (W/kg) over conventional SOA systems. In addition, it would be desirable to have a deployable solar array with revolutionary cost and performance improvements that is mechanically simple while meeting the requirements of currently available, as well as future solar cells. 
     The current solar array technology uses primarily panel based solar arrays that are poorly suited to leveraging the advantages of the next generation of Inverted-Meta-Morphic (IMM) multi junction thinned solar cells and their low areal density that are coming on the market in the near future. 
     Finally, the current SOA in ultra-high Performance deployable solar arrays uses membrane mounted solar cells and is designed for large spacecraft applications. There are two configurations, a dish-type fan fold or a blanket-type solar array. Both systems are mechanically complex and do not scale to smaller spacecraft applications. 
     SUMMARY 
     The present invention is a Deployable Modular Morphing Solar Array (DMMSA). The array is notionally simple, it uses a spring powered Root Staging and Deployment Mechanism (RSDM) that fan deploys structural elements similar to daisy petals that each perform a sequential secondary deployment. The stowed petals are folded when the system is stowed for launch on a spacecraft and unfold to a more structurally ideal configuration once deployed. The fan deployment moves the petals into position to be MORPHED-Deployed then locates them in positions ideal for gathering sun light. The petal assemblies are composed of a yoke that attaches to a Morphing Modular Solar Power Assembly, or assemblies (DMMSPA) that unfurl to form the petal assemblies upon beginning to fan deploy from the spacecraft. The DMMSA system is comprised of a Root Staging and Deployment Mechanism (RSDM) mounted to the spacecraft. The RSDM positions the stowed DMMSA 90 degrees from the spacecraft, staging it for fan deployment. Petal assemblies are attached to the RSDM by a yoke structure with each petal assembly having at least one DMMSPA secured thereon. Each DMMSPA elastically morphs to a slight V-configuration once deployed. This elastic flexing of the DMMSPA panel to a V cross section increases the area moment of inertia of the panel by orders of magnitude and hence the petal assemblies deployed natural frequency accordingly. 
     A launch restraint assembly secures at least one folded petal assembly prior to deployment with the launch restraint assembly pre-loading the petal&#39;s DMMSPA(s) into a substantially flat configuration. Upon release of the launch restraint assembly, the DMMSPA&#39;s that form each petal assembly elastically morph from the substantially flat configuration into the aforementioned V-configuration. 
     In addition, the present invention includes a method for deploying the petal assemblies from a spacecraft. The method comprises mounting the RSDM to the spacecraft and securing the yoke of the petal assemblies to the RSDM assembly. Deployment is accomplished by first swinging the un-deployed stack of petals to 90 degrees from the spacecraft then rotating the petal or petals away from the spacecraft in a sequential fan fashion. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a perspective view illustrating the DMMSA, constructed in accordance with the present invention, being in a stowed condition; 
         FIG. 2  is a perspective view illustrating an eight petal embodiment of the DMMSA, constructed in accordance with the present invention, being in a deployed condition; 
         FIG. 3  is a perspective view illustrating a petal assembly configured with two full DMMSPA&#39;s and two flip out solar panels, constructed in accordance with the present invention, with the petal assembly being in the deployed condition; 
         FIGS. 4   a - 4   h  are perspective views illustrating a deployment sequence for the DMMSA, constructed in accordance with the present invention; 
         FIG. 5  is a perspective view illustrating the RSDM of the DMMSA, constructed in accordance with the present invention; 
         FIG. 6  is a perspective view illustrating a graphite and matrix panel that is the structural element of a DMMSPA, constructed in accordance with the present invention, with the solar panel having a V bow; 
         FIG. 7  is a perspective view illustrating a pair of DMMSA&#39;s in 16 petal assembly embodiments, constructed in accordance with the present invention, mounted to a spacecraft on a boom and each being configured in a full circle; 
         FIG. 8  is an elevational end view illustrating the flattened and stowed DMMSPA&#39;s of the DMMSA, constructed in accordance with the present invention; 
         FIG. 9  is a perspective view illustrating DMMSA, constructed in accordance with the present invention, prior to deployment; 
         FIG. 10  is a perspective view illustrating the launch restraint system, constructed in accordance with the present invention; 
         FIGS. 11   a - 11   c  are perspective views illustrating the release sequence for the launch restraint system, constructed in accordance with the present invention; 
         FIGS. 12   a - 12   e  are perspective views illustrating the petal unfolding, constructed in accordance with the present invention; 
         FIG. 13  is a perspective view illustrating a petal latch and petal lanyard, constructed in accordance with the present invention; and 
         FIG. 14  is a perspective view illustrating the petal latch and the petal lanyard, constructed in accordance with the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     As illustrated in  FIGS. 1-14 , the present invention is a deployable, structurally morphing, modular solar array system, indicated generally at  10 , that increases the deployed stiffness of the modular petal assemblies  12 , improving the system&#39;s  10  deployed first mode natural frequency, and reducing overall manufacturing costs and mass. As will be described in further detail below, the DMMSA  10  of the present invention uses pre-loaded and flattened DMMSPA panels  14  arranged in petal assemblies  12  (each petal assembly  12  has at least one DMMSPA  14 ) for surviving the ascent vibration environment. During the deployment sequence the flattening load is released allowing DMMSPA  14  panels that are a part of the petal assembly  12  to flex into a slight V bow. Initially the elastic motion into a V configuration breaks mechanical or electrostatic sticking, that is common once in the outer space environment, and then increases the stiffness of the petal assemblies  12  when they are deployed. 
     The DMMSA  10  of the present invention includes a Root Staging and Deployment Mechanism (RSDM)  16  that provides two functions. First, the RSDM  16  swings the stowed solar array away from the spacecraft into a staged position for fan deployment. Second, after staging is complete, the RSDM  16  deploys each petal assembly  12  sequentially from the stacked configuration which is followed by the elastic self-deployment of each petal assembly  12 . The RSDM  16  includes a spacecraft interface bracket  18  securable to the spacecraft or a stand off boom mounted to the spacecraft and a clevis  20  that is pivotally connected to the bracket  18 . When the DMMSA  10  is stowed and secured to the notional spacecraft, the clevis  20  is initially positioned 90 degrees to the spacecraft interface bracket  18 . Once the system is released for deployment, the clevis  20  rotates to a position parallel to the spacecraft interface bracket  18  and hence the stowed solar array petals  12  approximately ninety (90°) into a staged position that is perpendicular to the mounting surface on the spacecraft for fan deployment. 
     In a preferred embodiment of the RSDM  16 , a torsion spring  28  connects the spacecraft interface bracket  18  to the clevis  20  biasing the clevis  20  to rotate to a position parallel to the spacecraft interface bracket  18  locating the stowed petals  12  to a position perpendicular to the spacecraft mounting plane. In a preferred embodiment this motion is damped by a viscous rotary damper  23  known to a person skilled in the art. Attached to the clevis  20  is at least one petal yoke  22 . The RSDM  16  includes a constant force spring mounted to an output drum  24  and a storage drum  26  to create the torque for deploying the petal assemblies  12  and hence the individual DMMSPAs  14 , as will be described in further detail below. 
     The RSDM  16  of the DMMSA  10  of the present invention deploys the petal assemblies  12  using multi-leaf constant force springs. The constant force springs develop the torque that deploys the first petal assembly  12  which then pulls subsequent petal assemblies  12  sequentially through petal to petal lanyards. Once fully deployed the constant force springs provide sufficient torque to keep the petals  14  of the DMMSA  10  deployed. The torque produced by the RSDM  16  can be fine-tuned by adding or subtracting constant force springs. Actual deployment of the petal assemblies  12  will be described in further detail below. 
     Each petal assembly  12  of the DMMSA  10  of the present invention is attached to the RSDM clevis  20  with a yoke bracket  23 . As mentioned above, in a launch state, the petal assemblies  12  are folded, stacked and held compressed flat so each individual DMMSPA  14  panel is held preloaded and flat. Preloading of the elements of a solar array that support solar cells prevents vibration induced gapping and the spike loads caused by this phenomena. These spike loads can damage solar cells. Conventional solar array systems utilize compressed foam or springs to preload the solar array panels that support the solar cells when stowed adding complexity and mass. The shallow V-shape of the deployed individual DMMSPA panels  14  provides the DMMSA  10  a stable preload in the launch configuration because the DMMSPA panels  14  are elastically deformed to a flat configuration when the DMMSA  10  is stowed. Additionally, the foam often used in conventional systems relaxes during stowage, thus increasing the risk of preload loss and limiting long-term stowage. The use of the elastic deformation of the flattened DMMSPA panels  14  of this invention, when stowed, optimizes mass and cost performance by having fewer parts and is structurally stable. 
     Each individual DMMSPA structural panel  14  of the DMMSA  10  of the present invention is constructed of cyanate ester and carbon fiber (CFRP) with an integral crease  30  in the center. The thickness of the panel  14  is determined by the inertial loads applied to the system during the rocket launch vibration environment. When the petal assemblies  12  and the individual DMMSPA panels  14  are stowed, the center crease  30  in each individual panel  14  allows the petal assemblies  12  to be elastically flattened. In the flat configuration, the petal assemblies  12  stow efficiently and preload the stowed system. Once the individual DMMSPA is  14  no longer under compressive pressure, it returns to its shallow V-shape. This “morphing” is an approximately ten (10°) degree bend in the individual panel  14  but increases the moment of inertia of the section as previously noted. 
     The petal assemblies  12  of the DMMSA  10  of the present invention are stacked and sandwiched under a preload by a launch restraint assembly  32 . In a preferred embodiment, the launch restraint assembly  32  includes the petal assemblies  12  positioned between a vehicle interface spider  34  on the bottom of the stacked petal assemblies  12  and a launch restraint swing spider  36  on the top of the stacked petal assemblies  12  applying a compressive load through multiple stacks of cup-cone elements  38  attached to the vehicle interface spider  34 , all the DMMSPA&#39;s  14 , and the swing spider  36 . When stowed for launch, the DMMSPA&#39;s  14  of each petal assembly  12  and its integral cups and cones  38  located in several places along its axial center line transfer shear loads as well as axial loads determinately securing each DMMSPA  14  to the spacecraft through the launch restraint assembly  32 . In the stowed configuration, the petal assemblies  12  are additionally stabilized with a multitude of rubber snubbers  40 . When stowed and flattened the petal assemblies  12  form a pre-loaded system due to the elastic forces required to flatten the individual DMMSPA&#39;s  14 . 
     In addition, the launch restraint assembly  32  of the DMMSA  10  of the present invention includes a spider link member  42  that is positioned between the vehicle interface spider  34  and the swing spider  36  on a distal end of the petal assemblies  12 . A hinge connection between the spider link  42  and the swing spider  36  allows the swing spider  36  to be rotated in a general direction away from the stacked petal assemblies  12  in order to stage and deploy the petal assemblies  12  and the individual DMMSPA&#39;s  14 . A hold down and release bolt  44  is positioned between the vehicle interface spider  34  and the swing spider  36  on the near end of the petal assemblies  12 . In conjunction with the spider link member  42 , the hold down and release bolt  44  holds the petal assemblies  12  sandwiched between the vehicle interface spider  34  and the swing spider  36 . The launch restraint assembly  32  keeps the folded petal assemblies  12  sandwiched, elastically compressing the individual DMMSPA panels  14 , and maintaining a stable long-term preload on the DMMSPA&#39;s  14  during storage and launch. 
     The sequence for deploying the petal assemblies  12  and hence the individual DMMSPA&#39;s  14  of the DMMSA  10  of the present invention will now be described. As understood by those persons skilled in the art that the deployment sequence described herein is a preferred manner of deployment and other deployment sequences are within the scope of the present invention. 
     First, as described above, the petal assemblies  12  are in the pre-loaded stored condition mounted to the spacecraft by the RSDM  16  and the launch restraint assembly  32 . When the spacecraft reaches a desired position of orbit or travel, the hold down and release bolt  44  is broken or otherwise damaged by known means such as applying power to a heater circuit that breaks the hold down and release bolt  44  thereby releasing the swing spider  36  from the vehicle interface spider  34 . The release is low shock and is not instantaneous, thus making it immune from spurious spikes of current due to electrostatic discharge. The released, un-loaded individual DMMSPA&#39;s  14 , and thus, the petal assemblies  12 , then relax into the V-shape thereby separating the cup-cones and causing the swing spider  36  to pivot away from the petal assemblies  12 . The petal assemblies  12  are now ready to be staged into the deployed condition. 
     In order to move the petal assemblies  12  into the deployed condition, the RSDM  16  rotates the stacked petal assemblies  12  approximately ninety (90°) degrees by torque from the torsion spring between the vehicle interface bracket  18  and the clevis  20  of the RSDM  16  to correctly position the petal assemblies  12  relative to the spacecraft. The petal assemblies  12  are now ready to be deployed with the individual DMMSPA&#39;s  14  in each petal assembly  12 , one at a time, flipping outward and unfolding. The actual amount of flipping and unfolding of the individual petals  12  is dependent on the actual number of individual DMMSPA&#39;s  14  that form each petal assembly  12 . In a preferred embodiment, the staging and fan deployment of the petal assemblies  12  is damped with dampers to limit speed. 
     As the first petal assembly  12  rotates away from the spacecraft, at a predetermined point, such as approximately eleven (11°) degrees, a petal to petal lanyard  25  begins pulling the next petal assembly  12  from the stowed stack of petal assemblies  12 , releasing a petal latch  27  on the first petal assembly  12  that allows the petal  14  to unfold. Initially, the remaining petal assemblies  12  remain stationary through a ball detent located on each yoke bracket  23  in the RSDM  16 . Release of each petal&#39;s petal latch  27  allows the petal assembly to unfold. Once the first petal assembly  12  unfolds and flips, it is fanned away from the remaining stacked petal assemblies  12 . DMMSPA  14 -to-DMMSPA  14  unfolding occurs when the petal latch  27  on each petal assembly  12  is released and petal deployment continues until full deployment. Each adjacent petal assembly  12  is tethered with the petal to petal lanyards  25  to the next adjacent petal assembly causing each successive petal assembly  12  to fan outward with this procedure continuing until all petal assemblies  12  are fanned away from the spacecraft. Now, the petal assemblies  12  create a deployed wing comprised of individual DMMSPA&#39;s  14  for powering the spacecraft and/or the spacecraft&#39;s equipment. It should be noted that the DMMSA  10  of the present invention is simple to reset by folding and rotating the petal assemblies  12  and replacing the hold down and release bolt  44  with a new replacement bolt. 
     The DMMSA  10  of the present invention advances the SOA of deployable photovoltaic power systems. There are many potential benefits and impacts to space missions by using the DMMSA  10 . The benefits and impacts include, but are not limited to: 
     1) Lower costs for small satellites requiring high power; 
     2) Enhancing the capability and utility of satellites; 
     3) Improving the mass fraction for payloads; 
     4) Allocating less volume for solar arrays on the stowed spacecraft; 
     5) Increasing power in the current volume allocation; 
     6) Varying power by using different number of petal assemblies; and 
     7) Rapidly reconfiguring for multiple power needs in support of rapidly deployable space missions. 
     The foregoing exemplary descriptions and the illustrative preferred embodiments of the present invention have been explained in the drawings and described in detail, with varying modifications and alternative embodiments being taught. While the invention has been so shown, described and illustrated, it should be understood by those skilled in the art that equivalent changes in form and detail may be made therein without departing from the true spirit and scope of the invention, and that the scope of the present invention is to be limited only to the claims except as precluded by the prior art. Moreover, the invention as disclosed herein may be suitably practiced in the absence of the specific elements which are disclosed herein.