Patent Application
20060227063
Not Applicable
Not Applicable
The present invention relates in general to a reflector antenna, and more particularly, to a dual-shell satellite reflector assembly with an improved rear structure designed to increase stiffness, reduce distortion during the lifetime of the reflector, and improve structural attachment of holddown fittings.
Satellite antenna systems are used to receive and transmit signals to and from the satellite. The transmit antenna is typically mounted on one side of a spacecraft to transmit signals from the spacecraft to receivers on the earth. During launch, the transmit antenna is secured to the spacecraft in the stowed configuration via four individual attach points and a hinge/gimbal interface. After the spacecraft has reached its intended orbit, the attach points are released and the antenna is deployed from the hinge mechanism at the hinge/gimbal interface.
A dual-shell reflector is typically used when a linearly polarized output is required. Particularly, such type of reflector is necessary to provide coverage on orthogonal polarizations.
The principal components of a typical dual-shell reflector include a front shell and a rear shell separated from each other by a circumferential or “intercostal” ring, and a backing structure. All the components are bonded together with a structural adhesive to obtain a unitized structure. The front shell is typically comprised of Kevlar™ skins co-cured to Kevlar™ or Korex™ core, which provides an RF-transparent surface. Grids are formed on the surface of the front shell to control polarization. The grids are spaced from each other with a predetermined distance to provide optimal electrical performance. An electrical energy within a first frequency range is fed to the surface of the front shell and reflected off the grids. An electrical energy within a second frequency range is also fed to the surface of the front shell. The electrical energy within the second frequency range transmits through the front shell and is reflected off the surface of the rear shell. The rear shell is constructed from carbon fiber skins co-cured to Kevlar™ or Korex™ core. The fiber carbon sandwich of the rear shell ensures electrical isotropic performance and maintains polarization purity. Similar to the front shell, the intercostal ring is a sandwich construction comprised of Kevlar™ skins co-cured to Kevlar™ or Korex™ core to allow specific frequencies to be received by and reflected from the rear shell. The backing structure is fabricated from carbon fiber skins co-cured to Kevlar™ or Korex™ core bonded together using a structural adhesive. The high specific stiffness of the carbon fiber backing structure acts to reduce distortion during launch and minimize the thermoelastic distortions of the structure during the lifetime of the satellite.
The typical satellite structure requires four hinge release mechanism (HRM) to hold the antenna in the stowed configuration until it reaches its intended orbit. One conventional design of the backing structure uses carbon fiber cylindrical tubes to encapsulate metallic fittings for tie-down of the antenna. The metallic fittings are bonded to the tubes. The sandwich panels or ribs of the conventional backing structures are bonded to the tubes to create a box-type of frame.
The conventional designs of the backing structure have several drawbacks. A common frequency requirement for antenna structures at launch condition is 55 Hz, while an analytical prediction shows that the launch frequency of the conventional design ranges approximately 45 Hz to 60 Hz. The material difference between the tie-down tube and the metallic fitting results in risk of disengagement at thermal excursions. Further, these designs do not efficiently distribute the loads between each of the hinge release mechanisms and the hinge/gimbal locations.
It is therefore a substantial need to provide a reflector backing structure that overcomes the drawbacks.
A backing structure of an antenna reflector for satellite communication is provided to increase the stiffness and the frequency of the antenna reflector. The antenna reflector includes a deployable panel attached to a spacecraft at hinge/gimbal interfaces and a plurality of hinge-release points, and the backing structure comprises a plurality of first ribs and second ribs protruding from the deployable panel and extending substantially across the center of mass of the deployable panel. The first ribs connect the hinge-release mechanism points across a center of mass of the deployable panel, and the second ribs connect the hinge/gimbal interface points to the hinge-release mechanism points.
Preferably, the first and second ribs are fabricated from sandwich panels of graphite skin and honeycomb core. In one embodiment, the backing structure preferably includes two hinge-release mechanism points distal to the hinge/gimbal interface point and two hinge-release mechanism points proximal to the hinge/gimbal interface point. Therefore, there are two first ribs extending between the proximal hinge-release mechanism points and the distal hinge-release mechanism points at different sides of the center of mass, and two second ribs extending between the hinge-gimbal interface point and the distal hinge-release mechanism points. Preferably, a third rib is formed to extend laterally to connect two of the hinge-release mechanism points to form a star-like rib structure on the reflector panel.
Preferably, each hinge-release mechanism point comprises a metal fitting and a nut for retaining the metal fitting, and each metal fitting includes a plurality of threads engagable with the nut. The nut prevents disengagement of the fitting at thermal excursions. The antenna reflector comprises a dual-shell deployable panel interconnected by a circumferential ring or a single-shell deployable panel.
An antenna reflector deployably attached to a spacecraft is also provided to overcome the drawbacks of the conventional design. The antenna reflector comprises a reflector panel, a hinge-gimbal interface, a plurality of hinge-release mechanisms and a star-like rib backing structure. The hinge-gimbal interface deployably connects the reflector panel to the spacecraft. The reflector panel is also releasably connected to the spacecraft by the hinge-release mechanisms. The star-like ribs extend between the hinge-gimbal interface and the hinge-release mechanisms to provide a plurality of direct load transfer paths to a center of mass of the reflector panel.
In one embodiment, each of the hinge-release mechanisms comprises one threaded metallic fitting and one nut to retain the threaded metallic fitting. The star-like rib is preferably fabricated from sandwich panels of graphite skin and honeycomb core. The star-like rib includes two longitudinal ribs extending across the center of mass and connecting the hinge-gimbal interface to two hinge-release mechanisms, and two diagonal ribs each extending across the center of mass and connecting two hinge-release mechanisms at two diagonal positions of the reflector panel. One lateral rib may also be formed to connect two hinge-release mechanisms proximal to the hinge/gimbal interface. The antenna reflector can be either a single-panel structure or a dual-panel structure that includes a front panel and a rear panel.
A method of increasing stiffness of an antenna reflector is also provided. The antenna reflector includes a reflector panel deployably connected to a hinge/gimbal interface and releasably connected to a plurality of hinge-release mechanisms of a spacecraft. A plurality of protruding ribs is formed on the reflector panel to extending across a center of mass of the reflector panel between the hinge-mechanisms. At least two protruding ribs are also formed to extend across the center of mass between the hinge/gimbal interface and the hinge-release mechanisms. A lateral protruding rib may also be formed to connect two hinge-release mechanisms proximal to the hinge/gimbal interface.
These as well as other features of the present invention will become more apparent upon reference to the drawings wherein:
Referring now to the drawings wherein the showings are for purpose of illustrating preferred embodiments of the present invention only, and not for purposes of limiting the same,
As shown in
However, as the protruding ribs 16 and 18 are formed to interconnect the hinge-release mechanism points 12a at the same sides of the center of mass, only the protruding ribs 14 that connect the hinge/gimbal interface point 12b and the distal hinge-release mechanism points 12a extend closely to the center of mass 11. That is, only the protruding ribs 14 provide direct load paths from the center of mass. The indirect load path established by the protruding ribs 16 and 18 still transfer the load from the center of mass to the hinge mechanism points 12a through the rear shell 20b, which is mostly unsupported at the center of mass. Thereby, a lower frequency of the reflector structure is resulted.
To prevent from transferring load from the rear shell 20b, so as to avoid lowering the frequency of the reflector structure, a star-like rib structure is formed to provide more load paths extending near or across the center of mass. As shown in
The above description is given by way of example, and not limitation. Given the above disclosure, one skilled in the art could devise variations that are within the scope and spirit of the invention. Further, the various features of this invention can be used along, or in varying combinations with each other and are not intended to be limited to the specific combination described herein. Thus, the invention is not to be limited by the illustrated embodiments but is to be defined by the following claims when read in the broadest reasonable manner to preserve the validity of the claims.
