Spacecraft antennas and beam steering methods for satellite communciation system

• US 5,642,122 A
• Filed: 05/11/1994
• Issued: 06/24/1997
• Est. Priority Date: 11/08/1991
• Status: Expired due to Term

• Current Assignee: Wengen Wireless LLC (Intellectual Ventures LLC)
• Original Assignee: Teledesic, LLC (Michael J. Larson, MBA)
• Inventors: Lockie, Douglas Gene, Thomson, Mark
• Primary Examiner(s): Wimer, Michael C.
• Application Number:
  US08/241,103
• Time in Patent Office: 1,140 Days
• Field of Search: 343/853, 343/915, 343/881, 343/DIG. 2, 343/878
• US Class Current: 343/881
• CPC Class Codes: B64G 1/22 Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
• CPC Class Codes: H01Q 1/288 Satellite antennas
• CPC Class Codes: H01Q 25/00 Antennas or antenna systems providing at least two radiating patterns arrangements for changing or varying the orientation or the shape of the directional pattern H01Q3/00
• CPC Class Codes: H04B 7/2041 Spot beam multiple access
• CPC Class Codes: Y10S 343/02 Satellite-mounted antenna

Abstract

An advanced active element phased array satellite antenna is disclosed. Incorporating these novel antenna systems on a constellation of low Earth orbit spacecraft, allows phone customers worldwide to communicate through a system whose switching intelligence resides on orbit, bypassing traditional land-based networks, and offering a revolutionary expansion of communications potential. The present invention utilizes electronic beam steering is utilized to provide extremely high gain signals. In one preferred embodiment, a satellite (S) includes an Earth-facing array (10) of hexagonal antenna facets (12), mated together along their sides to form a slightly flattened, hemispherical shell. The antenna array (10) is connected to two rectilinear, unfurled, solar panels (P). The antennas (10) transmit and receive signals from terrestrial units located within the footprints (14) of the beams (11). Another embodiment (34) uses deployable, folding panels (41) which maximize panel surface area for a given weight and launch vehicle container volume. The panels (41) are stacked upon a central plate (39) along the depth axis of the container in accordion fold layers and are separately deployed radially from the center (C) of the central plate (39). The contour of the deployed panels permit the antenna beam footprints (14) to cover the desired area.

Claims

1. A deployable, folding satellite antenna panel apparatus capable of being carried aboard a launch vehicle stowage container, said stowage container having a diameter (D) and a depth (H), said deployable, folding satellite antenna panel apparatus comprising:

a central plate (39) having a polygon shape having n sides (SD), n being an integer divisible by 2, and a center (C);

said central plate (39) having a first adjacent side (S1) and a second adjacent side (S2) with reference to each one of said n sides (SD);

said central plate (39) having an obverse side (O) and a reverse side (R) and having a planform, peripheral outline determined by inscribing said polygon shape within said stowage container diameter (D); and

a plurality of articulated arms (40) having a plurality of non-reflecting, structural support panels (41) for actively transmitting and receiving radio signals;

said non-reflecting, structural support panels (41) being stowed in layers in an accordion fold, upon said central plate (39), each one of said plurality of non-reflecting, structural support panels (41) having a hinge (43) along an edge by which each of said plurality of non-reflecting, structural support panels (41) is joined to another and to said central plate (39), each of said plurality of non-reflecting, structural support panels (41) having a thickness (t) and having a plurality of devices disposed thereon, including a plurality of discrete antennas (32);

said articulated arms (40) being the same in number as the number of said n sides (SD);

each one of said articulated arms (40) being separately deployable in a radial direction from said center (C);

said plurality of articulated arms (40) when positioned by rotation about said hinge (43) into said accordion fold, being stowed on both said obverse side (O) and said reverse side (R) of said central plate (39) in an absolute minimum axial distance (d) determined only by the aggregate of said thickness (t), which maximizes surface area of said plurality of panels (41) for a given satellite weight and said launch vehicle stowage container diameter (D) and depth (H).

2. The apparatus as claimed in claim 1 in which each said hinge (43) registers each of said plurality of non-reflecting, structural support panels (41) in a line of sight region in respect of the surface of the Earth (E) that is defined by a predetermined mask angle, when said articulated arms (40) are deployed.

3. The apparatus as claimed in claim 1 in which said plurality of articulated arms (40) assume an oblate, hemispherical configuration when fully deployed.

4. The apparatus as claimed in claim 1 in which each one of said plurality of articulated arms (40) further comprises:

a first non-reflecting, structural support panel (41a);

said first non-reflecting, structural support panel having a first hinge line (h1) coextensive with one of said n sides (SD) of said central plate (39) and a second hinge line (h2) parallel to said first hinge line (h1);

said first non-reflecting, structural support panel (41a) having a planform which is an area included between said first hinge line (h1), said second hinge line (h2), a first perpendicular (P1) erected to said first hinge line (h1) at an intersection with said first adjacent side (S1), a second perpendicular (P2) erected to said first hinge line (h1) at an intersection with said second adjacent side (S2), a first perpendicular bisector (B1) of said first adjacent side (S1), and a second perpendicular bisector (B2) of said second adjacent side (S2);

said first perpendicular bisector (B1) and said second perpendicular bisector (B2) each extended until it intersects said center (C);

said second hinge line (h2) intersecting each said perpendicular bisector (B1, B2) when said one of said plurality of articulated arms (40) is in a stowed position;

a second non-reflecting, structural support panel (41b) hinged to said first non-reflecting, structural sunport panel at said second hinge line (h2);

said second non-reflecting, structural support panel (41b) having a planform which is an area included between said first hinge line (h1), said second hinge line (h2) when in its stowed position, said first adjacent side (S1), said second adjacent side (S2), said first perpendicular bisector (B1) and said second perpendicular bisector (B2); and

a third non-reflecting, structural support panel (41c) hinged to said second non-reflecting, structural support panel at a third hinge line (h3), said third hinge line (h3) overlying said first hinge line (h1) when said plurality of non-reflecting, structural support panels (41) are stowed in said layers;

said third non-reflecting, structural support panel (41c) having a planform area which is an area included between said third hinge line (h3), said first adjacent side (S1), said second adjacent side (S2), said first perpendicular bisector (B1) and said second perpendicular bisector (B2).

5. The satellite antenna apparatus as claimed in claim 3 in which said oblate, hemispherical configuration is convex in respect of a direction along the Zenith (Z) toward the Earth (E).

6. The satellite antenna apparatus as claimed in claim 3 in which said oblate, hemispherical configuration is concave in respect of a direction along the Zenith (Z) toward the Earth (E).

7. The apparatus as claimed in claim 3 in which said plurality of articulated arms (40) include non-reflecting, structural support panels (41d) for communicating directly between satellites (S).

8. The apparatus as claimed in claim 4 in which each one of said articulated arms (40) further comprises:

a plurality of additional non-reflecting, structural support panels (41i) numbered fourth through ith;

said ith panel (41i) being hinged to an i-1 panel, which is a preceding panel, at an ith hinge line;

each one of said plurality of additional non-reflecting, structural support panels (41i), up to and including said ith panel (41i), being hinged to a said preceding panel at an ith hinge line (hi) in a similar manner as said second panel (41b) is hinged to said first panel (41a) at said second hinge line (h2) and said third panel (41c) is hinged to said second panel (41b) at said third hinge line (h3), respectively;

said fourth non-reflecting, structural support panel (41d) and each succeeding numbered panel, through a panel numbered i-1, having a planform including an area no greater than said planform area of said second panel (41b);

said ith panel (41i) having a planform which is an area included between said ith hinge line (hi), said first adjacent side (S1), said second adjacent side (S2), said first perpendicular bisector (B1) and said second perpendicular bisector (B2).

9. The apparatus as claimed in claim 7 in which said plurality of discrete antennas (32) includes a plurality of fixed and electronically steerable antennas.

10. A method of constructing deployable, folding satellite antenna panels, capable of being carried aboard a launch vehicle stowage container, said stowage container having a diameter (D) and a depth (H), comprising the steps of:

providing a central plate (39), said central plate (39) having a polygon shape having n sides (SD), n being an integer divisible by two, and a center (C);

said central plate (39) having a first adjacent side (S1) and a second adjacent side (S2) with reference to each one of said n sides (SD);

said central plate (39) having an obverse side (O) and a reverse side (R) and having a planform, peripheral outline determined by inscribing said polygon shape within said stowage container diameter (D);

providing a plurality of articulated arms (40) joined to said central plate (39), the number of said articulated arms (40) being the same as the number of said n sides (SD) of said central plate (39);

each of said plurality of articulated arms (40) having a plurality of non-reflecting, structural support panels (41) for actively transmitting and receiving radio signals, each of said non-reflecting, structural support panels (41) having a thickness (t) and a hinge (43) along an edge by which each of said plurality of non-reflecting, structural support panels (41) is joined to another and to said central plate (39);

each of said non-reflecting, structural support panels (41) having disposed thereon a plurality of discrete antennas (32);

stowing upon both said obverse side (O) and said reverse side (R) of said central plate (39) in an absolute minimum axial distance (d) determined only by the aggregate of said thickness (t), said plurality of articulated arms (40), in layers in an accordion fold; and

separately deploying each one of said articulated arms (40) in a radial direction from said center (C);

said method of stowing maximizing surface area of said plurality of panels (41) for a given satellite weight and said launch vehicle stowage container diameter (D) and depth (H).