Abstract:
A space vehicle electromechanical system may employ an architecture that enables convenient and practical testing, reset, and retesting of solar panel and antenna deployment on the ground. A helical antenna winding fixture may facilitate winding and binding of the helical antenna.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application claims the benefit of U.S. provisional patent application Nos. 62/016,548 and 62/016,566, both filed on Jun. 24, 2014. The subject matter of these earlier filed applications is hereby incorporated by reference in its entirety. 
    
    
     STATEMENT OF FEDERAL RIGHTS 
     The United States government has rights in this invention pursuant to Contract No. DE-AC52-06NA25396 between the United States Department of Energy and Los Alamos National Security, LLC for the operation of Los Alamos National Laboratory. 
    
    
     FIELD 
     The present invention generally relates to space vehicles, and more particularly, to a space vehicle electromechanical system and helical antenna winding fixture. 
     BACKGROUND 
     In conventional space vehicles, the electromechanical system cannot be efficiently tested, reset, and tested again to ensure reliable and effective operation. Also, effectively deploying high gain antennas and solar panels in small space vehicles has not previously been possible in a reliable, low cost manner. Accordingly, an improved space vehicle electromechanical system that addresses these issues may be beneficial. 
     SUMMARY 
     Certain embodiments of the present invention may provide solutions to the problems and needs in the art that have not yet been fully identified, appreciated, or solved by conventional space vehicle electromechanical systems. For example, some embodiments of the present invention employ an electromechanical system architecture that enables convenient and practical testing, reset, and retesting of solar panel and antenna deployment on the ground. 
     In an embodiment, an apparatus includes a constrained, deployable helical antenna and a constrained, deployable dipole antenna attached to one end of the helical antenna. Upon deployment, the helical antenna uncoils and the dipole antenna opens. 
     In another embodiment, an apparatus includes a constrained, deployable helical antenna and a ground plane attached to an end of the helical antenna. The apparatus also includes a cable connecting the helical antenna to the ground plane and a coiling and uncoiling cup configured to stow the cable when the helical antenna is stowed. 
     In yet another embodiment, a space vehicle includes a plurality of constrained solar panels and a release mechanism located on a plurality of tip plates that attach to respective solar panels of the plurality of solar panels. When released, the release mechanism is configured to allow the solar panels to deploy using a single release point. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       In order that the advantages of certain embodiments of the invention will be readily understood, a more particular description of the invention briefly described above will be rendered by reference to specific embodiments that are illustrated in the appended drawings. While it should be understood that these drawings depict only typical embodiments of the invention and are not therefore to be considered to be limiting of its scope, the invention will be described and explained with additional specificity and detail through the use of the accompanying drawings, in which: 
         FIG. 1  illustrates an exploded perspective view illustrating a cubesat with an opened chassis, according to an embodiment of the present invention. 
         FIG. 2A  is a perspective view illustrating a cubesat in a stowed configuration, according to an embodiment of the present invention. 
         FIG. 2B  is a perspective view illustrating the cubesat in a deployed configuration, according to an embodiment of the present invention. 
         FIG. 3  is a perspective view illustrating a cubesat deployment process, according to an embodiment of the present invention. 
         FIG. 4  is a vertically oriented side view of a helical antenna, according to an embodiment of the present invention. 
         FIG. 5  is a side view illustrating deployed antennas with design parameters, according to an embodiment of the present invention. 
         FIG. 6  illustrates a helical antenna ground plane printed circuit board (PCB) interface, according to an embodiment of the present invention. 
         FIG. 7  is a perspective view illustrating a crossed dipole antenna, according to an embodiment of the present invention. 
         FIG. 8  is a perspective view illustrating a dipole interface PCB for a crossed dipole antenna, according to an embodiment of the present invention. 
         FIG. 9A  is a perspective view illustrating a helical antenna lacing fixture with an unattached helical antenna, according to an embodiment of the present invention. 
         FIG. 9B  is a perspective view illustrating the helical antenna lacing fixture with an attached helical antenna, according to an embodiment of the present invention. 
         FIG. 10  illustrates a space vehicle power module, according to an embodiment of the present invention. 
         FIG. 11A  is a perspective view illustrating the back of a solar panel, according to an embodiment of the present invention. 
         FIG. 11B  is a perspective view illustrating the front of a solar panel, according to an embodiment of the present invention. 
         FIG. 12  is a perspective view illustrating the bottom of a power module, according to an embodiment of the present invention. 
         FIG. 13  is a top view illustrating four prototype solar panels, according to an embodiment of the present invention. 
         FIG. 14  is a perspective view illustrating a release mechanism, according to an embodiment of the present invention. 
         FIG. 15A  is a perspective view illustrating a space vehicle in a stowed configuration, according to an embodiment of the present invention. 
         FIG. 15B  is a perspective view illustrating the space vehicle in a deployed configuration, according to an embodiment of the present invention. 
         FIG. 16A  is a perspective view illustrating a top of a space vehicle with a release mechanism in a stowed configuration, according to an embodiment of the present invention. 
         FIG. 16B  is a perspective view illustrating a bottom plate of a cauterizing arm, according to an embodiment of the present invention. 
         FIG. 16C  is a perspective view illustrating a cauterizer of the cauterizing arm, according to an embodiment of the present invention. 
         FIG. 16D  is a perspective view illustrating the assembled cauterizing arm, according to an embodiment of the present invention. 
         FIG. 17  is an exploded perspective view illustrating a power module, according to an embodiment of the present invention. 
         FIG. 18  is a perspective view illustrating a LiFePO 4  battery assembly, according to an embodiment of the present invention. 
         FIG. 19A  is a perspective view illustrating the top of a power board, according to an embodiment of the present invention. 
         FIG. 19B  is a perspective view illustrating the bottom of the power board, according to an embodiment of the present invention. 
         FIG. 20A  is a perspective view illustrating a safe arm connector with a power module casing in place, according to an embodiment of the present invention. 
         FIG. 20B  is a perspective view illustrating the safe arm connector without the power module casing in place, according to an embodiment of the present invention. 
         FIG. 21A  is a perspective view illustrating a rail separation switch with a power module casing in place, according to an embodiment of the present invention. 
         FIG. 21B  is a perspective view illustrating the rail separation switch without the power module casing in place, according to an embodiment of the present invention. 
         FIG. 22  is a flowchart illustrating a process for deploying a space vehicle, according to an embodiment of the present invention. 
         FIG. 23A  is a perspective view illustrating a cubesat with a deployed multi-fold solar array, according to an embodiment of the present invention. 
         FIG. 23B  is a closeup perspective view illustrating a bi-fold solar panel of the cubesat, according to an embodiment of the present invention. 
         FIG. 24  is a perspective view illustrating a closed bi-fold solar panel, according to an embodiment of the present invention. 
         FIG. 25  is a perspective view illustrating a cubesat with a closed bi-fold solar panel in a stowed configuration, according to an embodiment of the present invention. 
         FIG. 26A  is a perspective view illustrating a locking pin mechanism with a locking pin in a retracted (stowed) position, according to an embodiment of the present invention. 
         FIG. 26B  is a perspective view illustrating the locking pin mechanism with the locking pin in an extended (deployed) position, according to an embodiment of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE EMBODIMENTS 
     Some embodiments of the present invention pertain to deployable antennas and solar panels, modular power systems, dispensers, and deployment processes.  FIG. 1  illustrates an exploded perspective view of a cubesat satellite  100  with an opened chassis, according to an embodiment of the present invention. In this embodiment, antennas, radio frequency (RF) equipment, and a housing  110  are located on top of cubesat  100 . A power module  120  includes two batteries that store power and provide power to internal components of cubesat  100 . Solar panels  130  convert solar energy into electricity, which is used to charge the batteries of power module  120 . 
       FIG. 2A  is a perspective view illustrating a cubesat satellite  200  in a stowed configuration, according to an embodiment of the present invention. Constrained helical and dipole antennas  210  are located on top of cubesat  200 . The deployment of antennas  210  is discussed in further detail below. 
     Constrained solar panels  220  are stowed folded upwards on cubesat  200 , forming a box shape. When deployed, solar panels  220  fold downward to form an “x” shape. See  FIG. 2B . However, it should be appreciated that the chassis of the cubesat of other embodiments may have any desired shape, and any number and configuration of solar panels may be used. Furthermore, each solar panel may be configured to fold in any desired direction. 
     Whereas most cubesats use a remove-before-flight pin, a safe/arm connector  222  is used to disconnect the battery from the main space vehicle electronics in some embodiments, ensuring that the space vehicle is powered down while stored on the dispenser or while sitting in storage. Safe/arm connector  222  may also provide communications and diagnostics for the space vehicle, allowing the space vehicle to be programmed, configured, and tested while in the fully assembled flight configuration, as well as allowing the batteries to be charged. A rail separation switch  224  is triggered when cubesat  200  is released from a dispenser or other storage mechanism once reaching the desired release location. Rail separation switch  224  may be used to trigger various actions of the deployment process, converting cubesat  200  from the stowed configuration of  FIG. 2A  to the deployed configuration of  FIG. 2B , for example. Furthermore, the battery or batteries may be disconnected by rail separation switch  224  until deployment, at which point the internal components of cubesat  200  power on. In some embodiments, connection of the battery may initiate a timer (e.g., 30 minutes) for releasing deployables, such as solar panels  220  and antennas  210 . 
       FIG. 2B  is a perspective view illustrating cubesat satellite  200  in a deployed configuration, according to an embodiment of the present invention. Antennas  210  extend vertically, uncompressing helical antenna  212  and unfolding crossed dipole antenna  214 . Helical antenna  212  has relatively high gain for data uplink and a narrower beam width than crossed dipole antenna  214 , which increases pointing accuracy. Crossed dipole antenna  214  has lower gain than helical antenna  212  and is omnidirectional. In other words, crossed dipole antenna  214  enables communication with cubesat  200  in any orientation. This generally cannot be achieved with high gain antenna  212  without much larger ground station antennas. 
     Conventionally, helical and dipole antennas are separate from one another, and the dipole antenna is proximate to the satellite chassis. Having the dipole antenna located relatively close to metal in the chassis reduces its capabilities. By attaching crossed dipole antenna  214  to the end of helical antenna  212 , improved performance may be realized. 
     Solar panels  230  deploy into an “x” configuration in this embodiment. However, other shapes and configurations are envisioned within the scope of other embodiments. When power is low, or more efficient charging is otherwise desired, cubesat  200  may be positioned such that solar panels  230  face the sun. Positioning of cubesat  200  may be achieved by any desired means, such as wheels, movable masses, ion thrusters, rocket engines, any other suitable positioning system, or any combination thereof. A camera  240  provides imaging capabilities. 
     A coiling and uncoiling cup  250  enables stowing and uncoiling of an RF cable  216  quickly and reliably. The helix of helical antenna  212  centers around coiling and uncoiling cup  250 , and an RF cable  216  passes through it. Coiling and uncoiling cup  250  may be fabricated using an additive manufacturing process, providing a lower cost part that generally cannot be built using normal machining processes. Coiling and uncoiling cup  250  provides a mechanism for stowing RF cable  216  before deployment, holding helical antenna  212  in place during mechanical vibration, ensuring a smooth release of helical antenna  212  without tangling supporting tethers  218 , and ensuring a smooth release of RF cable  216  during deployment, preventing tangling of RF cable  216 . 
     To improve the performance of helical antenna  212 , a deployable ground plane  260  is used in this embodiment. Ground plane  260  is formed by the top of the structure and by four deployable panels (i.e., “flappers”)  262  that give the circular shape to ground plane  260 . Flappers  262  may be released using the same, single-point release mechanism. When deployed, ground plane  260  is full sized, providing optimal antenna performance in some embodiments. 
     Dispenser and Deployment 
       FIG. 3  is a perspective view illustrating a cubesat deployment process  300 , according to an embodiment of the present invention. Cubesat  320  is launched into space via a rocket or other space vehicle while stored in a dispenser  310 . Dispenser  310  may house multiple satellites in some embodiments, and may be configured to release satellites at different points and/or in different directions. 
     Once released from dispenser  310 , a timer starts for cubesat  320  to deploy. Once the timer expires, cubesat  320  deploys its solar panels and antennas. Cubesat  320  may then orient itself in a desired orientation and begin operation in accordance with its mission. 
     Antennas 
       FIG. 4  is a vertically oriented side view of a helical antenna  400 , according to an embodiment of the present invention. Helical antenna  400  includes a coiled helix  410  that is held in place in a deployed state by three Kevlar™ threads  420 ,  422 ,  424 . However, in other embodiments, other materials and/or numbers of threads may be used. Furthermore, in certain embodiments, the threads may have a different pattern, such as a mesh. Crossed dipole antenna  440  is located at the end of helical antenna  400 , which reduces shadowing from a metal space vehicle body in some embodiments. A cable  430  provides a connection between the body of the space vehicle and crossed dipole antenna  440 . An internal circuit board (not shown) may transmit and process data from crossed dipole antenna  440  and/or helical antenna  400 . 
     Coiled helix  410  is a ten turn helix in this embodiment, with a pitch of 12.5° and an overall length of 13.08 inches. The circumference is 1.88 inches, which is λ/π at two gigahertz (GHz). λ represents the wavelength. In this embodiment, the circular polarization of helical antenna  400  is right handed. The half power beam width is 35° and the beam width at −1 dB is 10°. However, the number of turns, pitch, size, polarization direction, and beam width may be altered in other embodiments according to desired design parameters. 
       FIG. 5  is a side view illustrating deployed antennas  500  with design parameters, according to an embodiment of the present invention. In some embodiments, a design goal is to have the helical antenna rigid when deployed horizontally in Earth&#39;s gravity. This may be accomplished by ensuring that the total torque (T Total ) is greater than zero. The first turn of the helix generally requires the most restoring torque. In some embodiments, the deployed antenna is almost completely rigid when horizontal right on the edge of the free length. 
       FIG. 6  illustrates a helical antenna and ground plane printed circuit board (PCB) interface  600 , according to an embodiment of the present invention. In this embodiment, helical spring  610  is secured to ground plane PCB  630  by six double Teflon Kevlar™ cords  620 . However, any number of cords may be used as a matter of design choice. Furthermore, the cords may be made from any suitable material that allows helical spring  610  to deploy but holds the antenna substantially rigid, such as certain plastics, carbon fiber, or wire made from metal that does not interfere with the operation of the antenna. Helical spring  610  is over-pitched to ensure that cords  620  are under tension when the antenna is deployed. 
     Cords  620  are attached to helical spring  610  via beads of glue  622  in this embodiment. However, any suitable connector may be used, such as plastic beads physically attached to both helical spring  610  and cords  620 . Guides  640  attached to ground plane PCB  630  hold coils of helical spring  610  when it is stowed and assist with deployment of helical spring  610 . 
     Cords  620  are secured to ground plane PCB  630  via anchors  628 . Lacing from respective cords  620  is knotted, threaded through hole  625 , and bonded into insert  626 . Anchor  628  threads into insert  626 . 
     Antenna feed structure  612  acts as both a mechanism to attach helical spring  610  to ground plane PCB  630  and an impedance matching circuit to match the antenna to a 50 ohm cable, for example. Antenna feed structure  612  may be a quarter-wave microstrip trace, and as such, may be simple, low cost, and enable a separate connection point for an RF cable  624 . RF cable  624  is attached to ground plane PCB  630  using a connector (not shown), providing an external connector for radio and antenna testing while the space vehicle is in the fully assembled flight configuration. 
       FIG. 7  is a perspective view illustrating a crossed dipole antenna  700 , according to an embodiment of the present invention. A crossed dipole  710  is secured to a dipole interface PCB  720  via screws  712 . A prototype of a dipole interface PCB  800  is shown in  FIG. 8 . Dipole interface PCB  720  is secured to helical spring  740  via brace  730 , which holds crossed dipole antenna  700  in place at the end of helical spring  740 . A cable  750  interfaces with dipole interface PCB  720  and provides data from dipole interface PCB  720  to internal electronics of the space vehicle. In this embodiment, crossed dipole antenna  700  has a gain centered at 915 MHz and is LHCP polarized. In some embodiments, to enable more efficient stowing of cable  750 , a small diameter, flexible coaxial RF cable may be used, which generally has higher losses than larger diameter coaxial cables. 
     Helical Antenna Lacing Fixture 
       FIG. 9A  is a perspective view illustrating a helical antenna lacing fixture  900  an unattached helical antenna  930 , according to an embodiment of the present invention. Helical antenna lacing fixture  900  includes a rod  910  with a groove  912  to hold cords and holes  914  for screws  916  (see  FIG. 10B ) to hold helical wire  932  in place. A vertical plate  924  clamps onto and holds rod  910  in a horizontal position. However, in other embodiments, rod  910  may be positioned in other orientations. Vertical plate  924  fits into a base  920 , which includes a screw  922  that clamps vertical plate  924  in place. 
       FIG. 9B  is a perspective view illustrating helical antenna lacing fixture  900  with an attached helical antenna  930 , according to an embodiment of the present invention. Helical spring  932  is held in place on rod  910  via screws  916 , which are positioned so as to achieve the desired winding of helical spring  932 . Lacing of cords  938  starts at the top of helical antenna  930 , and cords  938  are knotted and bonded at each coil of helical spring  932 . Ground plane PCB  934  is attached to the end of rod  910  via tensioner  940 . Tensioner  940  also provides a constant preload to each cord  938  before bonding of each knot. Cords  938  are attached to ground plane PCB  934  in a similar manner to that shown in  FIG. 6 . 
     Power Module 
       FIG. 10  illustrates a space vehicle power module  1000 , according to an embodiment of the present invention. Power module  1000  plugs into the main body of a space vehicle. This modular design simplifies building, testing, and assembly of power module  1000  and the remainder of the space vehicle. 
     Power module  1000  includes four double sided solar panels  1010  that convert solar energy into electricity. A power board  1020  interfaces with solar panels  1010  and batteries  1030 , providing electricity to charge batteries  1030 . Power board  1020  also channels power from batteries  1030  to other space vehicle components. 
       FIG. 11A  is a perspective view illustrating the back of a solar panel  1100 , according to an embodiment of the present invention. Photovoltaic solar cells  1110  convert solar energy into electricity. A thermistor (not shown) is embedded inside solar panel  1100  to permit maximum power point tracking. The thermistor connector exits at point  1120 . 
       FIG. 11B  is a perspective view illustrating the front of a solar panel  1100 , according to an embodiment of the present invention. A hinge stop/indicator  1130  contacts a small switch in the space vehicle to indicate successful deployment. 
       FIG. 12  is a perspective view illustrating the bottom of a power module  1200 , according to an embodiment of the present invention. Maximum power point tracking thermistor connections  1220  from solar panel  1210  are connected to a thermistor (not shown) located inside the solar panel. A hinge stop/indicator  1230  is connected to the body of power module  1200  via a rod  1232  such that solar panel  1210  can pivot and deploy. 
       FIG. 13  is a top view illustrating four prototype solar panels  1300 , according to an embodiment of the present invention. Thermal sensors (not shown) are bonded into all blank panels. A Kapton™ layer was added to solar panels  1300  for temperature stability in this embodiment. 
       FIG. 14  is a perspective view illustrating a release mechanism  1400 , according to an embodiment of the present invention. In some embodiments, release mechanism  1400  is constructed from aluminum and glass filled Noryl™. Release mechanism  1400  releases and deploys solar panels  1420  and antenna assembly  1430 . Tip plates  1410  attach to solar panels  1420 , holding solar panels  1420 , ground plane extension flaps (not shown), and antenna assembly  1430  in place when they are stowed. 
     To deploy solar panels  1420  and antenna assembly  1430 , cauterizer tips  1412  are heated to cut a nylon line  1414 . The stowed force of antenna assembly  1430  deploys the antennas, solar panels  1420 , and ground plane extension flappers (see element  222  in  FIG. 2B , for example). Prior to launch, solar panels  1420 , antenna assembly  1430 , and release mechanism  1400  can quickly be reset, allowing for testing and retesting without disassembling the space vehicle. 
     The antennas may be designed such that RF signals are available for testing via an external connector (not shown) when the space vehicle is fully assembled. This enables testing of a space vehicle in a fully assembled configuration—for example to measure power output and frequencies, to validate radio functionality, and to validate antenna functionality. This is not present in conventional cubesat designs. 
       FIGS. 15A and 15B  illustrate a space vehicle  1500  with tip plates  1510  in a stowed and deployed configuration, according to an embodiment of the present invention. Tip plates  1510  are attached to the solar panels and remain attached following deployment. Tip plates  1510  also contain electronics, such as a magnetometer, used by the ADCS. 
       FIG. 16A  is a perspective view illustrating a top of a space vehicle with a release mechanism  1600  in a stowed configuration, according to an embodiment of the present invention. Release mechanism  1600  includes a pair of cauterizing arms  1610  configured to cut a nylon line and a pair of line holding arms  1620  configured to hold the nylon line in place until cut. Cauterizing arms  1610  and wire holding arms  1620  also hold the space vehicle in a stowed position until deployment. 
       FIG. 16B  is a perspective view illustrating a bottom plate  1630  of cauterizing arm  1610 , according to an embodiment of the present invention. Bottom plate  1630  includes a recess  1632 .  FIG. 16C  is a perspective view illustrating a cauterizer  1640  of cauterizing arm  1610 , according to an embodiment of the present invention. Cauterizer  1640  includes a replaceable cauterizer tip assembly  1642  including an insulator to insulate cauterizer tip assembly  1642  from the heat generated by cauterizer tip  1644 . Cauterizer tip assembly  1642  is placed within recess  1632  and plugs into a connector  1646 . Electrical current is supplied by the main space vehicle via connector  1646  to heat cauterizer tip  1644 .  FIG. 16D  is a perspective view illustrating assembled cauterizing arm  1610 , according to an embodiment of the present invention. A top panel  1650  covers cauterizer tip assembly  1642 . 
       FIG. 17  is an exploded perspective view illustrating a power module  1700 , according to an embodiment of the present invention. A pair of batteries provide primary power storage for the space vehicle. A power rail  1704  provides connections from batteries  1702  to the backplane (not shown). Backplane power sense and digital control lines  1706  provide the main space vehicle with access to temperature, voltage, and current sensors located on the power board and switches for sensing solar panel deployment. 
     A −Z sun angle sensor  1708  provides information about the position of the sun relative to the space vehicle. Solar panel hinges  1710  connect with hinges of a solar panel via a rod (not shown). Solar panel deployment detection switches  1712  detect whether the solar panels (not shown) are in a deployed state. Whereas most cubesats use a remove-before-flight pin, in this embodiment, a safe/arm connector  1714  is used to disconnect the battery from the main satellite electronics, ensuring the satellite is powered down while stored on the dispenser or while sitting in storage. The safe/arm switch also provides power, communications, and diagnostics for the space vehicle, allowing the space vehicle to be programmed, configured, and tested while in the fully assembled flight configuration, and also facilitates battery charging. 
     Solar panel cabling  1716  provides a path for the flow of electricity from the solar panels to batteries  1702 . Separation power switches  1718  are switched on when the space vehicle is released from a dispenser or other vehicle or container. Rail separation switch plungers  1720  engage separation power switches  1718  when the space vehicle is loaded into the dispenser due to being in contact with rails inside the dispenser, and are released when the space vehicle is released. 
       FIG. 18  is a perspective view illustrating a battery assembly  1800 , according to an embodiment of the present invention. A housing  1810  secures batteries  1820  in place within a power module of a space vehicle, fabricated using an additive machining process. A pair of high current welded tabs  1830  connect batteries  1820  to a power board, such as power board  1900  of  FIGS. 19A and 19B . 
       FIG. 19A  is a perspective view illustrating the top of power board  1900 , according to an embodiment of the present invention. Power board  1900  provides maximum power point tracking with temperature for the solar panels of the space vehicle, charges the batteries, and provides power to the backplane. In some embodiments, power board  1900  includes a 5V rail, a 3.3V rail, a 1.5V rail, and a high current direct connection to the batteries. A safe arm connector  1910  is mounted directly to power board  1900 , as shown in more detail in  FIGS. 20A and 20A . A power sense and digital interface  1920  is used to connect the space vehicle command and data handling (C&amp;DH) processor to the power board to access sensors and switches placed on the power board. High current power rail connections  1930  provide power to the main space vehicle assembly. A −Z sun sensor feed-through  1940  provides power to the sun sensor. 
       FIG. 19B  is a perspective view illustrating the bottom of power board  1900 , according to an embodiment of the present invention. Solar panel deployed indicator switches  1950  are activated for each solar panel when the respective solar panel deploys. A separation switch  1960  indicates that the space vehicle has been released from its dispenser or other deployment vehicle and connects/disconnects the batteries with the main power system. Two or more switches may be used in some embodiments to provide redundancy. 
       FIGS. 20A and 20B  are perspective views illustrating a safe arm connector  2000  with and without a power module casing  2010  in place, respectively, according to an embodiment of the present invention. Safe arm connector  2000  mounts directly to a power board  2020  and provides access to universal asynchronous receiver/transmitter (UART) to C&amp;DH, the battery charging port, and an external power supply port for 5V. 
       FIGS. 21A and 21B  are perspective views illustrating a rail separation switch  2100  with and without a power module casing  2110  in place, according to an embodiment of the present invention. In this embodiment, separation switch is a rail-based switch, ensuring that separation switch  2100  stays closed during vibration. When the space vehicle is deployed from a dispenser, for example, an internal rail of the dispenser is no longer in place, allowing separation switch  2100  to pop out. This may power on the space vehicle and start a timer (e.g., 30 minutes) for deployment of the solar panels and antenna. 
       FIG. 22  is a flowchart  2200  illustrating a process for deploying a space vehicle, according to an embodiment of the present invention. The process begins with releasing the space vehicle from a dispenser at  2210 . Once released, a separator switch of the space vehicle is tripped and electronics power on at  2220 . This may also start a timer for deployment of the antennas and solar panels. 
     The solar panels and antenna are then deployed at  2230 . This may be accomplished by the same release mechanism or different release mechanisms. In some embodiments, a cauterizing wire may be heated and a nylon line may be cut. The space vehicle may then be oriented using wheels or any other desired orientation mechanism at  2240 . The space vehicle then begins its mission at  2250 . 
     In some embodiments, the solar panels have a multi-fold configuration where the panels fold outward from the space vehicle and then fold at least one more time, increasing the overall surface area of the solar panel array. An embodiment of a bi-fold solar panel and locking pin are discussed in more detail below with respect to  FIGS. 23A, 23B, 24A, and 24B . The bi-fold panel doubles the solar cell area of the space vehicle solar panel in this embodiment, doubling the amount of power that is generated. However, one panel may be smaller than the other in some embodiments. Furthermore, in certain embodiments, the solar panels may fold out two times, three times, etc., and have three solar panels, four solar panels, etc. Any potential number of solar panels, panel shapes, and folds is encompassed within embodiments of the present invention. The number of solar panels and folds may be implemented based on power requirements, size requirements, cost, and complexity tolerance. 
     In some embodiments, solar cells are recessed into the solar panels. The solar panels may be aluminum, for instance. This may enable each solar panel pair in bi-fold embodiments fit within a cubesat dispenser when stowed where the space vehicle is a cubesat. Each solar panel may contain a flex-circuit underneath the solar cells. This flex circuit may route power from the cells to the space vehicle, and also route power, digital signals, and RF signals between the space vehicle and electronics mounted on the solar panel in some embodiments, such as deployment “hot tips,” GPS antennas, low noise amplifiers, magnetometers, etc. Using this scheme, the solar panels may be connected to the space vehicle via a simple connector, which greatly simplifies space vehicle fabrication and assembly while increasing reliability. The solar panels may also contain a deployment switch that indicates to the satellite that the panels deployed correctly. When stowed, the solar panels may form a rigid box structure, enabling the solar panels to be robust to mechanical shock and vibration while keeping the individual solar panels thin and light when deployed. 
     The solar panels may be spring deployed in some embodiments, and may be deployed in 1 g (i.e., Earth&#39;s gravity) to enable testing on the ground, such as in a thermal vacuum chamber. The solar panels may feature an innovative locking pin mechanism. This pin may perform several functions. First, the locking pin mechanism may lock the deployed solar panel in place, and not allow it to bounce back after being fully deployed. Second, the locking pin mechanism may hold the outer solar panel (i.e., the solar panel that folds out from the inner solar panel) in place during deployment, until the inner solar panel (i.e., the solar panel that is attached to the space vehicle) has fully deployed. Third, the locking pin mechanism may provide a release mechanism for the outer solar panel, deploying the outer solar panel at the appropriate time. Taken together, this design ensures that: (1) the solar panels and deployment mechanism are robust and not damaged due to random vibrations; (2) the solar panel deployment is smooth, and will not hang up against the space vehicle or other components during deployment; and (3) deployment is consistent and the solar panels are locked in place every time the solar panels are deployed. 
       FIG. 23A  is a perspective view illustrating a cubesat  2300  with a deployed multi-fold solar array having four bi-fold solar panels  2310 , according to an embodiment of the present invention. In a stowed configuration, outer solar panels  2312  are folded against inner solar panels  2314 , which are then folded against space vehicle  2300  via locking hinges  2318 . Solar panels  2310  form a rigid box frame when stowed, providing ruggedness during random vibration. Solar cells  2316  are recessed to enable stowed solar panels  2310  to fit within a cubesat dispenser. 
     During deployment of solar panel  2310 , inner solar panel  2312  first folds all the way down and locks in place via locking hinge  2318  before deploying outer solar panel  2314 . In some embodiments, solar panel  2310  is deployable in 1 g, enabling testing on the ground. Each solar panel may contain a flex-circuit (not shown) that routes power, digital and analog signals, and RF to components at the end of solar panels  2310 , such as the deployment mechanism “hot tips,” a GPS antenna, and electronics including low noise amplifiers and a magnetometer. Finally, solar panels  2310  integrate a deploy switch (not shown) to indicate to space vehicle  2300  that solar panels  2310  deployed correctly. 
       FIG. 23B  is a closeup perspective view illustrating bi-fold solar panel  2310  of cubesat  2300 , according to an embodiment of the present invention. When stowed, the outer edge of solar panel  2310  rests against a raised boss  2320  on inner solar panel  2312 , while tabs  2322  mate with tab indents  2324 , forming a solid box structure. This allows the use of thinner panels made from material such as aluminum. Although the deployed panels would likely not survive random vibration, when stowed in this manner, the resulting box structure is very strong and robust. Also, when stowed, locking pin  2326  attaches to pin receptacle  2328 . Locking pin  2326  may be spring loaded such that when the panel fully deploys, the pin retracts, freeing outer solar panel  2314  to deploy. This both locks solar panel  2310  in place and ensures that outer solar panel  2314  does not deploy until inner solar panel  2312  has fully deployed. Without this feature, outer solar panel  2314  may scrape or get hung up on the space vehicle surface or antenna elements during deployment, causing a deployment failure.  FIG. 24  is a perspective view illustrating a closed bi-fold solar panel  2400 , according to an embodiment of the present invention.  FIG. 25  is a perspective view illustrating a cubesat  2500  with a closed bi-fold solar panel  2510  in a stowed configuration, according to an embodiment of the present invention. 
       FIGS. 26A and 26B  are perspective views illustrating a locking pin mechanism  2600  including a locking pin  2610  in a retracted (stowed) position and an extended (deployed) position, respectively, according to an embodiment of the present invention. When stowed, locking pin  2610  holds the outer solar panel in place, and rests against the wall of a pin lock  2620 . Locking pin  2610  is spring loaded in this embodiment via a spring (not shown). When the inner solar panel is fully deployed, locking pin  2610  moves into a recessed hole  2630 , locking the inner solar panel in place and releasing the outer solar panel. 
     It will be readily understood that the components of various embodiments of the present invention, as generally described and illustrated in the figures herein, may be arranged and designed in a wide variety of different configurations. Thus, the detailed description of the embodiments of the present invention, as represented in the attached figures, is not intended to limit the scope of the invention as claimed, but is merely representative of selected embodiments of the invention. 
     The features, structures, or characteristics of the invention described throughout this specification may be combined in any suitable manner in one or more embodiments. For example, reference throughout this specification to “certain embodiments,” “some embodiments,” or similar language means that a particular feature, structure, or characteristic described in connection with the embodiment is included in at least one embodiment of the present invention. Thus, appearances of the phrases “in certain embodiments,” “in some embodiment,” “in other embodiments,” or similar language throughout this specification do not necessarily all refer to the same group of embodiments and the described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments. 
     It should be noted that reference throughout this specification to features, advantages, or similar language does not imply that all of the features and advantages that may be realized with the present invention should be or are in any single embodiment of the invention. Rather, language referring to the features and advantages is understood to mean that a specific feature, advantage, or characteristic described in connection with an embodiment is included in at least one embodiment of the present invention. Thus, discussion of the features and advantages, and similar language, throughout this specification may, but do not necessarily, refer to the same embodiment. 
     Furthermore, the described features, advantages, and characteristics of the invention may be combined in any suitable manner in one or more embodiments. One skilled in the relevant art will recognize that the invention can be practiced without one or more of the specific features or advantages of a particular embodiment. In other instances, additional features and advantages may be recognized in certain embodiments that may not be present in all embodiments of the invention. 
     One having ordinary skill in the art will readily understand that the invention as discussed above may be practiced with steps in a different order, and/or with hardware elements in configurations which are different than those which are disclosed. Therefore, although the invention has been described based upon these preferred embodiments, it would be apparent to those of skill in the art that certain modifications, variations, and alternative constructions would be apparent, while remaining within the spirit and scope of the invention. In order to determine the metes and bounds of the invention, therefore, reference should be made to the appended claims.