Batteries: thermoelectric and photoelectric – Photoelectric – Panel or array
Reexamination Certificate
2000-03-06
2001-09-04
Diamond, Alan (Department: 1753)
Batteries: thermoelectric and photoelectric
Photoelectric
Panel or array
C136S245000, C136S292000, C244S173300, C220S212000, C220S252000, C220S259300, C220S380000, C135S099000, C135S087000
Reexamination Certificate
active
06284966
ABSTRACT:
FIELD OF THE INVENTION
The invention relates to the field of the construction and deployment of a satellite. More particularly, the present invention relates to the construction and deployment of a satellite in the form of a Sphere using a plurality of flat panels.
BACKGROUND OF THE INVENTION
The development of microsatellites and nanosatellites for low earth orbits requires the collection of sufficient power for onboard payload instruments that are low in weight and low in volume. Because the overall surface area of a microsatellite or nanosatellite is small, body-mounted solar cells may be incapable of providing enough power to the payload instruments. A power choke problem is caused by inherent low solar panel efficient and inherent low solar energy per area resulting in the need for very large collection surfaces. Deployment of traditional, rigid, solar arrays necessitates larger satellite volumes and weights. Due to the solar radiation incident, large satellites also require additional navigation apparatus for pointing at a fixed attitude for maximum collection of solar energy.
Cylindrical spinning satellites have long been deployed to provide various functions such as communications and imaging. The cylindrical satellite has a plurality of flat thin elongated solar cell panels running the length of the satellite forming a cylinder. The plurality of solar cell panels is equiangularly radially positioned around the diameter of the cylinder. The cylindrical spinning satellite is deployed at a specific attitude relative to the earth and sun so that the satellite collects maximum solar energy to power onboard payload instruments. As the satellite collects solar radiation, the energy is expended by the instruments producing thermal radiation. The instruments must not over heat. The satellite design provides for sufficient thermal radiation so that the payload instruments are maintained within thermal limits. The payload instruments radiate thermal energy within the satellite cylindrical cavity. Due to a lack of symmetry, cylindrical spinning satellites require specific attitudes relative to the sun to reduce the variance of the solar radiation and hence to reduce the temperature variations of equipment within the cylinder. Hence, the cylindrical spinning satellite disadvantageously requires precision attitude guidance and deployment. The cylindrical spinning satellites typically have relatively large solar panels in elongated shapes requiring large stowage capacity within a deployment launch vehicle. The payload within the cavity is surrounded by a large and unwieldy solar panel rendering deployment difficult as well as disadvantageously requiring large launch cavities.
Spherical shapes have been proposed for nanosatellites, and geodetic shapes have been used for fabricating roughly spherical structures for many years. However, deployable spherical structures have not been designed for space satellites, and deployment methods for terrestrial geodetic shapes have not been made to enable easy deployment. These and other disadvantages are solved or reduced using the invention.
SUMMARY OF THE INVENTION
An object of the invention is to provide a satellite that is solar radiation insensitive to attitude position.
Another object of the invention is to provide a thermal environment that is independent of attitude for a Payload that is within the sphere.
Still another object of the invention is to provide a spherical enclosure for enclosing a payload using a plurality of flat polygonal panels.
Yet another object of the invention is to provide a method of deploying a satellite in the shape of a sphere formed from a plurality of flat polygonal panels.
Still another object of the invention is to provide a method for deploying a housing enclosure having an arbitrary outer curvature formed from a plurality of flat polygonal panels.
The invention in a first aspect is an apparatus that is primarily directed to a power sphere that can be used as a spherical satellite having an external shape in the form of a sphere approximated by a plurality of flat polygonal panels enclosing a payload. The panels are preferably solar panels for collecting solar energy and also function as passive radiators. In a second aspect, the invention is a method that is primarily directed to stowing and then deploying the flat polygonal panels to form the spherical enclosure after deployment. The enclosure housing approximates a sphere with a spherical curved exterior surface. Such a housing enclosure could be used as a spherical array of flat polygonal solar panels in the case of a satellite, but could be extended to various types of housing enclosures, such as, a geodesic camping tent used for human recreation in parks, wilderness and snow covered areas.
The spherical shape renders the satellite attitude-insensitive to solar radiation, in that, at any arbitrary attitude, the amount of external area exposure to solar radiation is a constant. While in view of the sun at any arbitrary sun angle, the same solar power is collected, regardless of the satellite attitude. A payload may be disposed within the sphere. The payload may radiate thermal energy in all directions passively transmitted through the panels. The spherical satellite payload enclosure enables solar energy collection and power generation while enabling passive thermal control. Due to the symmetry of the sphere, the solar radiation exposure and passive thermal radiation remains constant, irrespective of the attitude of the satellite relative to the sun, thereby stabilizing temperatures. The enclosure is particularly useful in nanosatellite and microsatellite designs.
The spherical design using the plurality of flat polygonal panels provides for improved power collection efficiency and improved stowage capacity efficiency of the satellite. The implementation of the solar array formed by flat polygonal panels eliminates the need for solar array tracking and pointing mechanisms while reducing mass and complexity with increased power efficiency. The collection of solar power is only a function of the cross-sectional area of the solar array presented to the sun. Transient temperature variations of spacecraft equipment located within the spherical enclosure are moderated because the thermal barrier provided by the enclosing solar array effectively provides a layer of isolation. That is, the solar array serves a secondary purpose of providing passive transient thermal control, further increasing thermal maintenance efficiency. Additionally, in a stowed launch configuration, for a given desired solar array area, the deployable solar array panels are stacked in the preferred configuration in a more compact form than a conventionally designed satellite having externally attached solar panels of the same area, thereby reducing required stowage capacity and thereby reducing overall launch costs.
The preferred implementation is based upon a satellite or spacecraft with two hemispherical sets of deployable solar arrays, one at each end of a deploying strut, such that when deployed, each set of solar array panels forms a geodetic hemisphere. The two hemispheres, when completely unfurled, form a complete geodetic sphere enclosing the spacecraft payload. The complete enclosure configuration approximates a sphere using the flat polygonal panels enclosing the spacecraft payload. The spacecraft payload can be of any shape smaller than the enclosing solar array sphere.
The preferred deployable solar arrays may either be attached to the spacecraft payload with telescoping supporting struts or attached by hard mounted struts coupling the deployable solar arrays to the spacecraft payload. Telescoping supporting struts allow for a small size satellite relative to the overall spherical volume of the solar array. The hard mounted struts can be used in cases where the satellite axis is rigidly designed to span the diameter of the deployed spherical solar array.
The spherical solar array is stowed in as small a volume as practicable during the launch phase of the satellite using flat
Gilmore David G.
Hinkley David A.
Osborn Jon V.
Robinson Ernest Y.
Simburger Edward J.
Diamond Alan
Reid Derrick Michael
The Aerospace Corporation
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