Deployable and tracked solar array mechanism for nano-satellites

The present application claims benefit of priority of provisional patent application Ser. No. 61/276,197, filed on Sep. 9, 2009, entitled “CubeSat Tracked Solar Array”.

1. Field of the Invention

This invention relates generally to a deployable and tracked solar array mechanism for nano-satellites and, more particularly, the invention relates to a deployable and tracked solar array mechanism for nano-satellites for a wrap-folded solar wing formed with a frame that restrains the lengthwise edges of a center panel of the solar wing on each side prior to deployment.

2. Description of the Prior Art

CubeSats are small spacecraft often 10 cm×10 cm×30 cm. Existing CubeSat deployable solar array panels hinge from all four 30 cm bus faces in a maximum power configuration. While this yields the maximum area of cells, the structural architecture is not well suited to sun tracking and hence its Average Orbital Power (AOP) is severely limited due to non-optimal pointing during the majority of an orbit. The peak power of the existing State Of The Art (SOTA) CubeSat systems is 21 watts. When sun off pointing is factored in to obtain an Average Orbital Power (AOP), the power in a nadir pointing mission drops to 5 to 7 watts AOP.

In addition, restraining the solar array panels during liftoff, flight, and positioning of the spacecraft and deploying the solar array panels when ready presents numerous problems. It is of utmost importance to protect the solar array panels and to reliably release the solar array panels when desired. Due to the fact that the spacecraft is orbiting the Earth, it is not possible to “fix” any solar array panels or release mechanisms after the spacecraft is in orbit.

The present invention is a deployable and tracked solar array mechanism for restraining and releasing deployable solar panel arrays on a spacecraft. The mechanism comprises a solar array drive mountable to the top side surface (10 cm×10 cm) of the spacecraft. A solar panel array panel is pivotally attached to the solar array drive and positionable against the 30 cm×10 cm faces of the spacecraft. At least one notch is formed in at least one of the side edges of the solar panel array. A frame is slidably mounted to the spacecraft around the solar panel array. At least one tab extends from the frame over the solar panel array and releasably restrains the solar panel array. A spring mechanism urges the frame in a direction generally away from the solar array drive assembly. A release mechanism holds the frame against the force of the spring means wherein upon initiation of deployment, the release mechanism disintegrates allowing the spring means to slidably urge the frame in a general direction away from the solar array drive assembly thereby aligning the notches (at least one) with the frame tab thus releasing the solar panel array from being restrained against the spacecraft.

In addition, the present invention includes a method for restraining and releasing deployable solar panel arrays on a spacecraft. The method comprises mounting a solar array drive assembly to the top side surface of the spacecraft, providing a solar panel array having a center panel, pivotally attaching a first panel to one side edge of the center panel, pivotally attaching a second panel to the other side edge of the center panel opposite the first panel, pivotally attaching the center panel to the solar array drive assembly, positioning, the solar panel array against the spacecraft, forming at least one notch in at least one of the side edges of the solar panel array, slidably mounting a frame to the spacecraft around the solar panel array, extending at least one tab from the frame over the solar panel array, releasably restraining, the solar panel array with the at least one tab, urging the frame in a direction generally away from the solar array drive assembly, holding the frame against the force of movement, initiating deployment, releasing the hold on the frame, slidably urging the frame in a general direction away from the solar array drive assembly, aligning the at least one notch with the at least one tab, and releasing the solar panel array from against the spacecraft.

The present invention further includes a deployable and tracked solar array mechanism for restraining and releasing deployable solar panel arrays on a spacecraft. The spacecraft has a first side surface, a second side surface, a third side surface opposite the first side surface, a fourth side surface opposite the second side surface, a top surface, and a bottom surface. The mechanism comprises a solar array drive assembly mountable to the top side surface of the spacecraft. A first solar panel array and a second solar panel array are provided with each solar panel array having a center panel, a first panel pivotally attached to one side edge of the center panel, and a second panel pivotally attached to the other side edge of the center panel opposite the first panel. The center panels are pivotally attached to opposite sides of the solar array drive assembly. The first solar panel array positionable against the first side surface and the second solar panel array positionable against the third side surface. A plurality of notches are formed in at least one of the side edges of both the solar panel arrays. A first frame is slidably mounted to the first side surface of the spacecraft around the first solar array panel and a second frame is slidably mounted to the third side surface of the spacecraft around the second solar panel array. A plurality of first tabs extend from the first frame over the first solar panel array with the number of first tabs corresponding to the number of notches on the first solar panel array and the first tabs releasably restraining the first solar panel array. A plurality of second tabs extend from the second frame over the second solar panel array with the number of second tabs corresponding to the number of notches on the second solar panel array and the second tabs releasably restraining the second solar panel array. First spring means urge the first frame in a direction generally away from the solar array drive assembly and second spring means urge the second frame in a direction generally away from the solar array drive assembly. First release means holds the first frame against the force of the first spring means and second release means holds the second frame against the force of the second spring means. Upon initiation of deployment, both the release means disintegrate allowing the both spring means to slidably urge the frames in a general direction away from the solar array drive assembly thereby aligning the notches with the tabs and releasing the solar panel arrays from against the spacecraft.

FIG. 1 is an exploded perspective view illustrating a deployable and tracked solar array mechanism for nano-satellites, constructed in accordance with the present invention, for restraining and releasing deployable solar panel arrays on a spacecraft;

FIG. 1A is a cross-sectional perspective view of the frame, a pair of holding elements attached to the frame, and a pair of tabs;

FIG. 2 is a perspective view illustrating the deployable and tracked solar array mechanism for nano-satellites, constructed in accordance with the present invention, with a solar array drive assembly and a root hinge assembly for deploying the solar panel arrays away from the spacecraft;

FIG. 3 is a perspective view illustrating the deployable and tracked solar array mechanism for nano-satellites, constructed in accordance with the present invention, with the mechanism mounted to the spacecraft;

FIG. 4 is another perspective view illustrating the deployable and tracked solar array, mechanism for nano-satellites, constructed in accordance with the present invention, with direction of movement upon melting of meltable release pin;

FIG. 5 is an exploded perspective view illustrating the deployable and tracked solar array mechanism for nano-satellites, constructed in accordance with the present invention, with the solar panel arrays and the spacecraft;

FIG. 6 is a perspective view illustrating the deployable and tracked solar array mechanism for nano-satellites, constructed in accordance with the present invention, with the mechanism mounted to the spacecraft;

FIG. 7 illustrates a deployment sequence of the solar panel arrays for use with the deployable and tracked solar array mechanism for nano-satellites, constructed in accordance with the present invention;

FIG. 8 is a perspective view illustrating the solar panel array for nano-satellites, constructed in accordance with the present invention, with small folded braces deployed on the back sides (non-cell side) of the solar panel array; and

FIG. 9 is a close-up perspective view illustrating a portion of the deployable and tracked solar array mechanism for nano-satellites 9 shown in FIG. 3 , constructed in accordance with the present invention.

FIG. 10 is a close-up, perspective view illustrating a folded solar panel array positioned adjacent to the first side of a nano-satellite, root hinge assembly, spring mechanisms, a portion of the frame that has a pair of tabs and a pair of slots that allow the frame to be slidably mounted to the first side of the nano-satellite, and the tabs of the frame aligned with pairs of notches in the folded solar panel array to allow deployment of the solar panel array (as would occur after the release device has been released to allow the frame to be displaced so the tabs and notches move from an unaligned state to an aligned state).

As illustrated in FIGS. 1 , 1 A, and 2 - 10 , the present invention is a deployable and tracked solar array mechanism, indicated generally at 10 , for restraining and releasing deployable solar panel arrays 12 on a spacecraft 14 . The spacecraft 14 used with the mechanism 10 of the present invention is preferably a CubeSat spacecraft 14 although using the mechanism 10 on any type of spacecraft 14 is within the scope of the present invention. In addition, preferably, the solar panel arrays 12 are preferably constructed from a thin graphite-based composite panel stiffness and strength although using other types of solar panel arrays 12 is within the scope of the present invention.

The CubeSat spacecraft 14 has a four side surfaces, a top surface, and a bottom surface. Preferably, each side surface is substantially parallel to each opposite side surface and the top surface is substantially parallel to the bottom surface. Mounted on the top surface is a solar array drive assembly 16 . The solar array drive assembly 16 is preferably a simple, self contained, ultra-thin, low-power, stackable single actuator drive system designed to operate on approximately 500 mW of power while articulating the arrays, and less than 1 mW while the actuators and system are inactive. The solar array drive assembly 16 specifically fits in previously wasted payload space, on a typical CubeSat bus, consisting of an approximate 6.5 mm “slice” of the CubeSat bus.

While a specific type of solar array drive assembly 16 has been described for use with the deployable and tracked solar array mechanism 10 of the present invention, it is within the scope of the present invention to utilize any type of solar array drive assembly 16 so long as the solar array drive assembly 16 performs key system functions, rotating the solar panel arrays 12 to keep them optimally oriented with respect to the Sun, and providing a path for power transfer from the solar panel arrays 12 to the CubeSat spacecraft 14 .

The solar panel arrays 12 of the deployable and tracked solar array mechanism 10 of the present invention include a first solar panel array 18 and a second solar panel array 20 . Each of the first solar panel array 18 and the second solar panel array 20 has a center panel 22 , a first panel 24 pivotally attached to one side edge of the center panel 22 with a pivotal attachment 23 , and a second panel 26 pivotally attached to the other side edge of the center panel 22 opposite the first panel 24 with a pivotal attachment 23 . The center panels 22 of the first solar panel array 18 and the second solar panel array 20 are each pivotally secured to the solar array drive assembly 16 with a root hinge assembly 28 , each of which includes one of mounting fasteners 37 A, 37 B. The two root hinge assemblies 28 operate to deploy the solar panel arrays 12 , as described above. Furthermore, each of the edges of the center panels 22 has a series of spaced notches 30 for allowing deployment of the solar panel arrays 12 with the mechanism 10 , as will be described in further detail below.

Prior to deployment, the first solar panel array 18 is wrap-folded against the first side surface of the spacecraft 14 and the second solar panel array 20 is wrap-folded against the third side surface of the spacecraft 14 with the center panels 22 of each of the solar panel arrays 12 facing in a generally outward direction. By positioning the center panels 22 in a generally outward direction, complete string of cells for trickle charging is provided without having to deploy the solar panel arrays 12 .

While the solar panel arrays 12 , i.e., the first solar panel array 18 and the second solar panel array 20 , of the deployable and tracked solar array mechanism 10 of the present invention have been described as being tri-fold with three panels, i.e., a center panel 22 , a first panel 24 , and a second panel 26 , it is within the scope of the present invention for the solar panel arrays 12 to have multiple panels, i.e., more than three, flipping away from each other during deployment. Regardless of the number of panels, when the panels are stored, at least a portion of the center panel 22 faces outward and, when deployed, all of the panels face in a desired position to achieve the tracking and power required for the space mission. Furthermore, regardless of the number of panels, deployment springs 31 are positioned between each of the panels. Preferably, the springs are redundant stainless steel torsion rod springs that unfurl the panels of the solar panel arrays 12 during deployment, as will be described in further detail below.

The deployable and tracked solar array mechanism 10 of the present invention maintains the solar panel arrays 12 , i.e., the first solar panel array 18 and the second solar panel array 20 , against the spacecraft 14 until deployment is desired. The mechanism 10 further comprises a first frame 32 mounted to the first side surface of the spacecraft 14 and a second frame 34 mounted to the third side surface of the spacecraft 14 . Each frame 32 , 34 has an end member, a first side member substantially perpendicular and connected to one end of the end member, and a second side member connected to the other end of the end member and substantially parallel to the first side member. A plurality of slots 36 are formed in each of the first side member and second side member with a bolt 38 or other fastening mechanism extending through the slot 36 into the spacecraft 14 . The interaction of the slots 36 sliding along the bolts 38 allows deployment of the solar panel arrays 12 .

Each of the first side member and the second side member of each frame 32 , 34 of the deployable and tracked solar array mechanism 10 of the present invention has a series of holding devices 40 spaced therealong. Each holding device 40 extends in a generally outward direction away from the spacecraft 14 and has a frame tab 42 bending inward at an approximately ninety) (90° degree angle and substantially parallel to the spacecraft 14 . Each frame tab 42 of each holding device 40 contacts at least a portion of the edges of the center panels 22 of the solar panel arrays 12 adjacent the notches 30 for holding the solar panel arrays 12 against the spacecraft prior 14 to deployment, as will be described in further detail below.

In addition, the deployable and tracked solar array mechanism 10 of the present invention has a pair of spring mechanisms 44 acting against the first side member and the second side member of each frame 32 , 34 . Preferably, each spring mechanism 44 is a redundant stainless steel compression spring although using other types of spring mechanisms 44 is within the scope of the present invention.

The deployable and tracked solar array mechanism 10 of the present invention further includes a meltable release pin 46 interacting between the end member of each frame 32 , 34 and the spacecraft 14 . The release pin 46 maintains the position of each frame 32 , 34 relative to the spacecraft 14 against the force of the spring mechanisms 44 . When the deployment event is initiated by a command to burn the release pin 46 , the disintegration of the release pin 46 allows the spring mechanisms 44 to urge each of the frames 32 , 34 in a slidable direction, indicated by arrow 35 , opposite the solar array drive assembly 16 along the sides of the spacecraft 14 . This type of release event is preferably not instantaneous making the release event low shock and immune from spurious spikes of current due to Electro Static Discharge (ESD). Furthermore, the release pin 46 is lightweight, typically weighing less than one (1) gram, is easy to reset by replacing the release pin 46 .

Upon complete movement of the frames 32 . 34 of the deployable and tracked solar array mechanism 10 of the present invention, to the movement capacity of the spring mechanisms 44 , the notches 30 in the edges of the center panels 22 of the stowed solar panel arrays 12 align with the holding devices 40 . Being free from constraint by movement of the frames 32 , 34 , the spring-powered root hinge 28 interacting with the solar array drive system 16 rotates the solar panel arrays 12 in a general direction away from the first side surface and the third side surface of the spacecraft 14 , respectively.

As the solar panel arrays 12 of the deployable and tracked solar array mechanism 10 of the present invention reach approximately eighty (80°) degrees of deployment rotation from the spacecraft 14 , the solar panel arrays 12 begin to unfurl by spring energy. Simultaneously, small folded braces 47 deploy on the back sides (non-cell side) of the solar panel arrays 12 providing the solar panel arrays 12 with deployed stiffness and position accuracy. The preferred deployment sequence for the solar panel arrays 12 is illustrated in FIG. 7 .

The deployable and tracked solar array mechanism 10 of the present invention securely and releasably restrains the solar panel arrays 12 during liftoff, flight, and positioning of the spacecraft 14 . In addition, the mechanism 10 reliably releases the solar panel arrays 12 for deployment during orbit.

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.