EP0951760A1

EP0951760A1 - Method for controlling the transmission of a beam of radiated energy in a cellular satellite system

Info

Publication number: EP0951760A1
Application number: EP97923397A
Authority: EP
Inventors: Edward F. Tuck
Original assignee: Individual
Current assignee: Individual
Priority date: 1997-01-17
Filing date: 1997-01-17
Publication date: 1999-10-27
Also published as: WO1998032245A1; AU2921097A

Abstract

Satellite footprints (16a, 16b, 16c) are partitioned into elongate spanning cells (18) and multiple elongate segments (22). The elongate spanning cells (18) resemble long strips that extend along an entire footprint (16a, 16b, 16c). The alignment of the elongate spanning cells (18) along a footprint (16a, 16b, 16c) is determined by a correction angle Ζ, which compensates for the changing relative motion of the Earth's surface at various latitudes.

Description

METHOD FOR CONTROLLING THE TRANSMISSION OF A BEAM OF RADIATED ENERGY IN A CELLULAR SATELLITE SYSTEM

TECHNICAL FIELD

The present invention relates to the field of satellites. More particularly, this invention provides methods of conveying and controlling a beam of radiated energy to or from a satellite orbiting the Earth in an orbit which is not an Equatorial orbit at geosynchronous altitude. The present invention offers optimized frequency utilization and enhanced availability, capacity and reliability of remote sensing and communication services.

5 BACKGROUND ART

Over the past few years, several satellite communication systems have been proposed. These systems are generally designed to enhance existing terrestrial cellular networks. Cellular telephone networks utilize fixed transmitting and receiving stations located in adjacent cells. Some of the satellite communication systems that will be launched later in this decade will rely on satellites operating in low Earth orbits. In cellular networks, the

10 central transmitting and receiving antennas are stationary, and subscribers may communicate with one another while moving from cell to cell. A low Earth orbit network also allows subscribers to move, but the transmitting and receiving antennas aboard each satellite in the network are constantly in motion. Unlike some earlier communications systems which employed satellites in geosynchronous orbits, the spacecraft in low Earth orbit will move rapidly across the sky over any given place on the ground. Since the position of a low Earth orbit

15 (LEO) satellite is not fixed with respect to a location on the ground, some LEO networks require complicated schemes for steering radio beams to users on the ground. Since each LEO satellite is only visible for a few minutes from the user's point of view, any communication through the network that lasts more than a few minutes must be handled by more than one satellite. The complex switching that is required to "hand-off continuous communications service from one satellite to another imposes severe processing and power burdens on the LEO 0 satellites.

In previous satellite communication schemes, spacecraft which are not held stationary over one particular location on the Earth in geosynchronous orbits fly over large regions of the Earth very rapidly. The radio beams generated by these fast moving spacecraft sweep across vast regions of the Earth's surface at the same rate of speed. If these beams were visible to the eye, they would paint bright circular and elliptical patches of light on 5 the ground beneath the satellite which emitted them. In a system that employs satellite-fixed cells, the "footprint" of the radio beams propagated by the spacecraft defines the zone on the ground called a "cell" which is illuminated by the spacecraft. This satellite-fixed cell moves constantly as the spacecraft orbits around the globe.

United States Patent No. 5,268,694 by Jan et al. discloses a communication system employing low Earth orbiting satellites which project footprints on the Earth by means of antennas radiating and receiving energy. Each footprint is divided into cells. The cell shape is shown as hexagonal and described as being more circular or elliptical dependent upon the beam control. In U.S. Patent No. 5,422,647, Hirshfield et al. describe a communications satellite payload providing for communication between user devices, such as hand-held cellular telephones and terrestrial gateways. The payload supports a plurality of independent antenna beams. Jan et al. 5 show many discrete cells of small diameter which are organized in the footprint according to the best fit. In this manner they are not aligned with the satellite's track. Hirshfield et al. show somewhat larger and more elliptical cells than Jan et al. Hirshfield cells are organized into a best-fit pattern and are not maintained in any particular reference to the changing relative velocity of the satellite over the Earth as it travels through its orbit.

In International Publication WO 93/07683, Rouffet et al. teach a system for communication with

10 terminals via low-orbiting satellites, in which elongated elliptical beams of each coverage area are illuminated in accordance with a beam-hopping procedure. A corresponding English language version of this Patent Application may be found in Australian Patent Application AU9226029. The content of this earlier disclosure corresponds to the preamble of the independent claims.

Rouffet et al. specifically call for beams aligned with the direction of a satellites travel in orbit (direction

15 15) a vector in space. However, Rouffet et al. do not suggest that the antenna beams be continuously maintained aligned with a vector which is the satellite's track on the Earth's surface. The system described by Rouffet et al. is also constrained by the use of a single frequency band for communication. It is therefore necessary for the system to operate the beams alternatively to prevent interference. Rouffet et al. disclose sub-division of the beams to implement a time division multiple access technique for preventing interference between signals and for no 0 other purpose. However, their elongated cell is not disclosed as being divided into a plurality of linear segments.

One recent attempt to overcome the problem of providing a high capacity LEO communications system which ameliorates the burdens of complex switching due to frequent hand-offs between satellites is described in

U.S. Patent No. 5,408,237 entitled Earth-Fixed Cell Beam Management for Satellite Communication System.

This Patent describes methods and apparatus which pertain to the allocation of radio beams which are generated

25 by a constellation of satellites orbiting below geosynchronous altitude. These beams are precisely controlled so that they illuminate "Earth-fixed cells," as opposed to "satellite-fixed cells."

In sharp contrast to satellite fixed cells, an "Earth-fixed cell" is a stationary region mapped to an "Earth- fixed grid" that has permanent fixed boundaries, just like a city or a state. Although the rapidly moving satellites still shine their radio beams over the ground in rapidly moving footprints, the locations of the footprints at any

30 given time do not determine the location of the unchanging Earth-fixed cells. The great advantage provided by using cells having boundaries that are fixed with respect to an Earth-fixed grid is realized when a subscriber being served by one satellite must switch to another beam in the same satellite or to a second satellite because the first is moving out of range below the local horizon. With satellite-fixed cells, this "hand-off involves the assignment to the terminal of a new communication channel within the new beam or new satellite.

35 This assignment process takes time and consumes processing capacity at both the terminal and the satellite. It is also subject to blocking, call interruption, and call dropping if there is not an idle communication channel in the next serving beam or satellite. The Earth-fixed cell method avoids these problems by allocating communication channels (frequency, code and or time slot) on an Earth-fixed cell basis rather than on a satellite- fixed cell basis. Regardless of which satellite/beam is currently serving a particular cell, the terminal maintains the same channel assignment, thus substantially eliminating the "hand-off problem.

The Earth-fixed cell method uses software that provides position and attitude information about each satellite in the constellation. Position data from this software enables each satellite to map or bound the surface into areas representing an unchanging "Earth-fixed grid." Each satellite is capable of steering, transmitting and receiving beams conveying packets of information to the Earth-fixed grid. The beams are continually adjusted to compensate the effects of satellite motion, attitude changes, and the rotation of the Earth. These goals are accomplished by partitioning satellite footprints into elongate, preferably linear, spanning cells and optionally also into multiple elongate, preferably linear, segments. Unfortunately, many satellite communication and remote sensing systems that will be launched in the near future will rely on the more conventional method of utilizing satellite-fixed cells. Without the great benefits of an Earth-fixed cell method, a large portion of the microprocessor capabilities and power capacities of these systems will be diverted to internal switching tasks.

This problem of designing satellite systems which employ satellite-fixed cells but which avoid the deleterious effects of complex switching due to frequent hand-offs has presented a major challenge to the satellite business. The development of a high capacity satellite system which is capable of using satellite-fixed cells but which also minimizes hand-off overhead would constitute a major technological advance and would satisfy a long felt need within the communications and remote sensing industries.

DISCLOSURE OF THE INVENTION

The preferred embodiment of the Cellular Satellite System pertains to the management and allocation of radio beams which are generated by a satellite which moves with respect to the Earth's surface. The invention pertains to any satellite which is not in an Equatorial orbit at geosynchronous altitude. The present invention reduces hand-off overhead in systems that utilize satellite-fixed cells, augments frequency re-use, and enhances the communications and sensing capacity of the satellite. These goals are accomplished by partitioning satellite footprints into elongate spanning cells and multiple elongate segments. The elongate spanning cells are areas that contain the satellite's track along the ground. They resemble long strips that extend along the entire footprint. Multiple elongate segments are smaller contiguous areas that lie within the elongate spanning cells. While the elongate spanning cells are generally located so that their longest axis is collinear with the apparent direction of travel of the satellite over the Earth's surface at the nadir, the longitudinal axis of the elongate spanning cells is actually slightly offset from the projection of apparent direction of travel of the satellite by a correction angle φ. This correction angle compensates for the changing relative motion of the Earth's surface at various latitudes.

As a result of creating elongate spanning cells and continually maintaining the cells in alignment with the satellite's track over the Earth's surface, terminals in the satellite's track remain within the spanning cells for approximately the entire time the satellite footprint passes over the terminal. This feature reduces the number of hand-offs as the satellite-fixed cells sweep over the terminal, augments frequency re-use and enhances the communications and sensing capacity of the satellite.

An appreciation of other aims and objectives of the present invention and a more complete and comprehensive understanding of this invention may be achieved by studying the following description of a preferred embodiment and by referring to the accompanying drawings.

BRIEF DESCRIPTION OF THE DRA WINGS

Figures 1 and 2 are schematic illustrations of the Earth and of satellites operating in orbits which are not Equatorial orbits at geosynchronous altitude.

Figure 3 reveals satellites in Earth orbit and the footprints they produce on the Earth's surface. Figure 4 is a schematic diagram that shows a single footprint and a single elongate spanning cell. The alignment of the elongate spanning cell has been corrected by an angle φ which compensates for the changing relative motion of the Earth's surface at various latitudes.

Figure 5 is a schematic diagram that reveals a single footprint and several elongate spanning cells.

Figure 6 presents another schematic diagram showing a single footprint, several elongate spanning cells and a number of multiple elongate segments within each elongate spanning cell.

Figures 7, 8 and 9 supply graphs which illustrate the functional relationships of the correction angle versus argument of perigee for a near-polar orbit, surface velocity versus latitude and latitude of nadir versus argument of perigee for an inclined orbit.

Figure 10 is a schematic illustration of a satellite in Earth orbit. This figure also portrays the satellite's track on the Earth, which resembles a spiral.

Figure 11 is a perspective view of a satellite which is designed for use with the preferred embodiment of the invention.

Figures 12, 13, 14, 15 and 16 are schematic block diagrams of components aboard the satellite shown in Figure 1 1 which are employed to practice the present invention.

BEST MODE FOR CARRYING OUT THE INVENTION

Satellites & Geometry

Figures 1 and 2 show a constellation 10 of satellites 12 circling the Earth E as they travel along polar, equatorial, or inclined orbits 11. Figures 1 and 2 also depict the Equator and several Great Circles. The present invention may be practiced using only one or many satellites 12. The present invention pertains only to satellites that use satellite-fixed cells, i.e., satellites that are constantly in motion with respect to the Earth. As a consequence, the orbits 1 1 shown in Figures 1 and 2 may be any orbit which is not an Equatorial orbit or an orbit at geosynchronous altitude. When used in this Specification and the Claims that follow, the term "orbit" refers to any trajectory of movement above the atmosphere of a celestial body, including circular and elliptical trajectories.

Figure 3 exhibits three satellites 12 moving along an orbital pathway 11 above the Earth E in a direction of travel T. When the line defining the direction of travel T in orbit is projected down onto the Earth's surface, which is rotating west to east, a continuous line is formed on the surface at the satellite's nadir N. This line is called the satellite's track T. Each satellite 12 includes a radiation interface 14 which is capable of transmitting and/or receiving radiated energy 15a, 15b and 15c. The term "radiated energy" pertains to any form of energy, including a signal, which is transported from one position to another. Radiated energy includes all forms of waves and particles irrespective of wavelength, including electromagnetic and optical radiation.

If the satellite 12 is a communications satellite, the radiation interface 14 comprises transmit and receive antennas. If the satellite employed to practice the present invention performs remote sensing tasks, the radiation interface 14 comprises sensors or detectors. The spacecraft may also be used to direct some beneficial transmission toward the Earth E. Examples of such transmissions may include messages from a store-and-forward telecommunication system, compressed musical works for transfer to compact disks or illuminating radiation for a resource-mapping radar. In one of the embodiments of the invention, the satellites 12 communicate with personal, mobile and fixed terminals P, M, F and gateways G on or near the surface. The word "terminal" is intended to encompass any device capable of sending or receiving radiated energy. Terminals include gateways. While most terminals are located on the Earth, they may be near the Earth, e.g., they may reside on an aircraft.

The area on the surface of the Earth that is illuminated by the radiated energy 15 emitted by a communications satellite 12 is known as the satellite's footprint 16. In the case of a remote sensing satellite, the interface 14 collects some form of radiation from an elliptical region on the surface of the Earth known as a "reception zone." A reception zone 17 defines the area at a given time which is "viewed" by a sensor aboard a satellite 12. A reception zone is substantially the same as a footprint 16, but a footprint 16 is defined by radiated energy 15 directed toward and impinging upon the Earth, while a reception zone 17 is defined by a surface on the Earth from which energy is radiated that may also be detected by a satellite at a particular instant. The footprint and reception zone boundaries are set by the radiation pattern of the radiation interface 14. The boundaries are selected by a selected roll-off of signal from a maximum value at or near the center of the footprint 16 or zone 17. The three elliptical regions depicted in Figure 3 may be either footprints 16a, 16b and 16c or reception zones 17a, 17b and 17c.

Figure 4 exhibits a single footprint 16 containing a single elongate spanning cell 18. In accordance with the present invention, the spatial points in each footprint 16 or reception zone 17 are mapped into long strips referred to as linear spanning cells 18 by electronic transformation of the spatial points by the radiation interface. The elongate spanning cells extend across along an entire footprint 16 or reception zone 17. Each cell 18 has two parallel sides along its longest dimension and a longitudinal axis 20 halfway between its two longest sides and parallel to them. The two relatively short sides of the cell 18 are collinear with an ellipse, because the boundary of the footprint 16 or the reception zone 17 which each cell 18 spans is always an ellipse itself. The edges of the elongate spanning cells are shaped by beam shaping and steering electronics aboard the satellite 12.

A footprint 16 and any elongate spanning cells 18 within it are constantly moving across the Earth's surface as the satellite 12 which creates the footprint 16 flies across the sky overhead. The instantaneous position of an elongate spanning cell 18 relative to its footprint 16 is determined by the track T' of the satellite 12 over the ground and by a correction angle φ. The correction angle is chosen to align the satellite radiation interface 14 longitudinal axis 20 with the satellite's motion over the Earth's surface. If the Earth were not turning, no correction would be needed.

Figure 4 shows the instantaneous positions of both an uncorrected and a corrected elongate spanning cell extending along a footprint 16. The uncorrected elongate spanning cell 18 is lightly shaded, and is centered along the projected line of the satellite's direction in orbit marked by reference character T. The elongate spanning cell which has been aligned by altering its position by correction angle φ is heavily shaded, and is centered along the line T, marked by reference character 24. The correction angle φ is measured at the intersection I of the satellite's track T and the line T, the projection of the satellite's direction in orbit. For the sake of clarity, the magnitude of the correction angle φ has been exaggerated in Figure 4.

Figure 5 portrays a footprint 16 containing a number of linear spanning cells 18, while Figure 6 portrays both linear spanning cells 18 and multiple linear segments 22 within a footprint 16 (or reception zone 17). The multiple linear segments 22 are provided by dividing the cell 16 with additional shaped and steered beams 15. Multiple linear segments 22 add additional "space division multiple access" to the communication system. This provides more reuse of allocated frequencies, hence a larger number of users can be accommodated simultaneously.

Calculating the Correction Angle φ

The following discussion concerns the derivation of the appropriate correction angle φ based on the satellite's latitude. The derivation applies to the special case of a near-polar orbit.

If the longitudinal axis of the cell is parallel to the orbital plane of the satellite (which is fixed in space, and does not rotate with the Earth), and the Earth is turning under the satellite 12 as it proceeds in its orbit 1 1, from the point of view of an observer on Earth the satellite's track T' is not parallel to the lines of longitude, but lies at a small angle to them. This small angle, called the "correction angle" and represented by the variable φ, varies from about four degrees at the Equator to zero at the poles. A graph 200 of the correction angle φ 202 versus latitude 206 is presented in Figure 7. The angle φ reverses when the satellite 12 passes over a pole. As a result of Earth's rotation, the terminal appears to move across the cell. If the linear spanning cells 18 are one degree of latitude (about 115 km) wide, and the satellite is in a near-polar orbit with an altitude of about 700 km, the terminal will seem to move about 3/4 of the way across the cell 18 in the cell's three minute dwell time at the Equator, decreasing toward the poles. The dwell time is the period of time the satellite footprint spends on a given point on the Earth's surface.

This phenomenon will increase handoff activity, since about half of the terminals will have to change cells from this cause in their three minute dwell time. As described above, this problem can be substantially eliminated by swinging the cell pattern a few degrees in each direction during the satellite's orbit. This compensation may be accomplished using a slow mechanical movement of the antenna or an electrical means, depending on the antenna design, and is described in complete detail below.

In the equations that follow, the frame of reference is the "fixed stars", i.e., a reference frame based on inertial space. The Earth is turning on its axis and satellites are moving in their orbits with respect to this frame of reference. Since the orbits of satellites of Earth move through space with the Earth's center of mass, we have assumed for convenience that the frame of reference moves with the Earth in its orbit.

The surface of the Earth E as seen from this frame of reference moves from West to East. The velocity is highest at the Equator, and decreases in proportion to the cosine of the latitude to zero at the poles. Because one nautical mile (about 6076.1 feet) is defined as one minute of longitude at the Equator, it is very convenient in the following analysis to work in nautical miles and in nautical miles per hour (knots). The surface velocity of the Earth at a latitude L, V_S(L), in knots, as seen by a distant observer in the reference frame is always from West to East, and is given by Equation One:

V (L)= cos( ) (Equation One) ^s 24

Where L is the latitude in degrees of the point on the Earth's surface at which the velocity is being measured.

A graph 214 of the Earth's surface velocity 216 as a function of North or South latitude 218 as shown in Figure 8.

In the simple case of a circular polar orbit, an orbit of constant altitude which passes over both the North Pole and the South Pole, the motion of the satellite is always at right angles to the motion of the Earth's surface, and, assuming a spherical Earth, the orbital period is the time it takes for the satellite to travel completely around the Earth, may be found using Equation Two:

(Equation Two) Where τ is the period of orbit; r_e is the radius of the Earth, i.e., 6378.165 km; h is the orbital altitude in km; and

GM is the Earth's mass times the gravitational constant, G=KmY²; GM=398,603.2.

For a satellite 12 in low Earth orbit at 700 km, τ=98.773 minutes. (This is the orbital period, i.e., the times it takes for the satellite to travel completely around the Earth.)

The orbital velocity of a satellite in a circular orbit is given by Equation Three:

F_o=360— (Equation Three)

For a satellite in a 700 km orbit, the orbital velocity is then V₀=l .312x10" knots.

The motion of the satellite's nadir across the surface of the Earth is the vector sum of the satellite's velocity and the Earth's surface velocity. At any given latitude, the angle at which the Earth's surface is moving with respect to the direction of the satellite's motion in the inertial reference frame is the angle portion of that vector sum. If the satellite 12 is moving Northward in its orbit, and the orbit is at a small angle to the Earth's axis, the Earth's surface will move from left to right below the satellite, and to fit that motion, the satellite's radiation pattern must be aligned in a Northwest-Southeast direction. As the satellite approaches the North Pole, the necessary angle of the correction decreases, and is zero when.directly over the North Pole NP. As the satellite proceeds Southward, the correction angle must swing to the satellite's right, to a Southwest-Northeast alignment, since West is now on the satellite's right. The correction angle reaches a maximum as the satellite crosses the Equator. The angle then decreases to zero as the satellite crosses the South Pole SP, reversing to a Northwest- Southeast alignment after the South Pole SP is crossed. The discussion presented below utilizes the navigator's convention that True North is zero degrees, with the direction angle increasing as direction changes toward the East. In other words, the angle of direction changes from zero degrees through 360 degrees as the compass needle moves clockwise from the starting point of True North. The correction angle, φ(L), which is the angle of the motion of the Earth's surface with respect to the satellite at latitude L, may be determined using Equation Four:

V_sL(ω) φ( ) =90 X-arctan (Equation Four) o

Where V₀ is the orbital velocity of the satellite; i is the orbital inclination, /. e. , the angle measured between the plane containing the satellite's orbit and the plane containing the Earth's Equator; ω is the argument of perigee, i.e., the satellite's orbital position as given by the number of degrees of a 360-degree orbit it has traveled since crossing the Equator in a Northbound direction; and L(ω) is the latitude of the satellite at argument of perigee ω. L(ω) is equal to ω for satellites in polar orbits, for a value of ω of ninety degrees or less.

The graph 200 presented in Figure 7 was generated using a value of i of ninety degrees, which corresponds to a "pure" polar orbit. In the case of a polar orbit, the latitude of the satellite's nadir is the same as its orbital position. Consequently, for the special case of a satellite in polar orbit, L(ω)=ω from the Equator to the North Pole.

For this polar orbit, the correction angle varies as the satellite 12 makes one complete orbit as shown in Figure 7. This figure presents data for a Northbound satellite in a circular polar orbit starting at the Equator (latitude= 0 degrees, ω= 0 degrees), proceeding Northward to the North Pole (latitude= 90 degrees, ω= 90 degrees), then Southward to the Equator (latitude again 0 degrees, ω= 180 degrees), continuing Southward to the South Pole (latitude= -90 degrees, ω= 270 degrees), and finally Northward again to the Equator (latitude= 0 degrees, ω= 0 degrees). Negative angles indicate motion to the left of the satellite as it proceeds in its orbit, while positive angles indicate motion to the right.

For a satellite in an inclined orbit, the satellite's nadir never reaches the pole, so that the maximum latitude is less than ninety degrees. The latitude of the nadir can be found at any moment from Equation Five:

sin( ) =sin(z) *sin( +ω) (Equation Five)

where L is the latitude of the nadir; i is the inclination of the orbit; f is a very small anomaly which may be ignored for the purposes of this calculation; and ω is the argument of perigee of the satellite. The graph 208 depicted in Figure 9 is based on Equation Five and plots the latitude of nadir 210 versus the argument of perigee 212 for a satellite in an inclined orbit.

Using the Correction Angle φ to Point the Satellite Antenna: A Preferred Method Employing a GPS Receiver 5 Once the correction angle φ has been determined, the correct angle of the radiation interface 14, which may be an antenna 34, with respect to the body of the satellite 12 must be ascertained. Other information that is required includes the relative speed and direction of the satellite, and the direction of the Earth's surface at the nadir.

The preferred method of acquiring the needed information employs a navigation receiver on the satellite

10 12 which receives and inteφrets the signals of the Global Positioning System (GPS). Such receivers are readily available from a number of manufacturers. Thousands of personal versions of GPS receivers are produced each month by companies such as Magellan™ and Garmin™. Space-qualified versions are produced by companies such as Interstate Electronics™, Motorola™ and Rockwell International™. These receivers provide continuous information on their position with respect to the Earth's surface, already corrected for the Earth's motion, and all

15 such receivers either provide, or can readily be adapted to provide, an output signal which gives a continuous value of the receiver's speed and direction with respect to the Earth's surface. The satellite 12 which is used to implement the present invention will include attitude sensing and stabilization mechanisms that supply an output signal which give a continuous value of the receiver's speed and direction with respect to the Earth's surface. Other systems on board the satellite 12 will contain attitude sensing and stabilization mechanisms which furnish 0 an output signal proportional to the orientation of the satellite's structure with respect to its direction of motion. These two signals can be added or subtracted to give a signal directly proportional to the relative angle of the satellite's body to the satellite's direction of motion over the Earth's surface T'. The antenna can then easily be rotated by a conventional servo-mechanism to place its long axis in line with that direction of motion.

Figure 10 is a schematic depiction of a satellite 12 in a ten day polar orbit. The track of the satellite on 5 the ground appears as a spiral T'.

Figure 1 1 reveals one of many satellite designs that may be utilized to realize the benefits of the present invention. The satellite 12 comprises a body 30, a solar array 32 and an antenna 34 which includes a narrow reflector with a parabolic cross-section. The internal components of the satellite 12 include an inertia wheel 36, an antenna positioning servo 38 and a radio-frequency feed to antenna 39. 0 Figures 12 and 13 are schematic block diagrams 39 and 45 depicting generalized groups of electronic circuits aboard the satellite 12. Figure 12 shows a GPS receiver 100 which produces an output that includes an altitude and latitude signal 40. This signal is fed to a converter 42 which, in turn, produces a control signal 44 that governs the action of an antenna positioning servo 38. The GPS receiver 100 is described in great detail in the text which follows and in Figure 16. 5 Figure 13 shows an alternative circuit that includes a radio receiver 46 which produces a synchronizing pulse 48 that is fed to timers 50. The timers 50 generate an elapsed time code 52 which is supplied to a converter 54. The converter emits a control signal 56 which directs the action of an antenna positioning servo 38. Figure 14 is a schematic block diagram 57 that portrays the GPS receiver 100 and accompanying electronic components in greater detail. The GPS receiver 100 generates two digital signals. The first signal comprises a number which is the speed of the satellite over the Earth's surface, and an angle which is the angle between the satellite's velocity and the underlying meridian of longitude. These quantities are components of the satellite's velocity vector, represented as box 62 in Figure 14. The second digital signal is the latitude of the satellite's nadir, shown as box 64 in Figure 14.

A microprocessor 66 receives these two signals 62 and 64 from receiver 100, in addition to the angle of the satellite's longitudinal axis with respect to the Earth's axis 60 which is supplied by an attitude sensor 58. The microprocessor 66 multiplies the vector constant equal to the surface velocity of the Earth's Equator by the cosine of the latitude. This product yields the vector of the surface velocity at the satellite's nadir, V„. The microprocessor 66 also calculates the resultant of the satellite velocity vector and the surface vector V_n. The angle portion of this resultant, corrected by the angle of the satellite's longitudinal axis with the Earth's axis, is the beam offset angle 68. This beam offset angle 68 is fed to a servo-mechanism 38, which rotates the antenna array 34 so that its long dimension makes the beam offset angle 68 with the direction of the satellite's motion. Figure 15 offers another schematic block diagram 71 that is specifically designed for use with satellites

12 that travel in circular orbits at a constant altitude. A radio receiver 46, a star tracker 72 or a GPS receiver 100 may be utilized to generate an Equator-crossing pulse 74 which resets a timer 76. The timer 76 sends a "time since crossing" (τ_c) signal 78 to a function generator 80. The τ_c signal 78 is proportional to the time that has elapsed since the satellite 12 crossed the Equator. The function generator 80 includes a memory which stores a quantity τ,_Λ, which is the time it takes the satellite to complete one half of its orbit around the globe, and also stores the satellite's velocity vector. The function generator output is equal to the angle of the satellite's apparent motion over the Earth's surface, taking into account the surface velocity of the Earth at the satellite's nadir. The surface velocity, V, is the velocity at the Equator, V, in Equation Six:

V'= VcosL (Equation Six)

where L is the latitude at the satellite's nadir.

The latitude in this case can be set equal to:

(Equation Seven)

since the cosine function inteφrets angles over ninety degrees. An adder 84 algebraically sums the correction angle from the function generator's output with the actual offset of the satellite's longitudinal axis with respect to the Earth's axis 82 and its direction of motion with respect to the Earth's axis. The result is the beam offset angle 86, which is used to a control servo-mechanism 38 that points the satellite antenna 34.

Figure 16 is a schematic block diagram 100 of the internal components and functions of a GPS receiver. In general, the GPS receiver is a superheterodyne with a voltage-controlled local oscillator. The existing 5 constellation of GPS satellites use a code-division modulation scheme (like CDMA), which allows all of the GPS satellites to use the same bandwidth segment. Signals from different GPS satellites are detected by matching them with a code that is peculiar to each satellite. This function is performed by an active detector 1 12, which receives an input from an antenna 102 through a low-noise amplifier 104, a mixer 106, a band-pass filter 108 and an intermediate frequency amplifier 110.

10 A satellite selector 118 uses an algorithm to pick a set of GPS satellites, usually four, which are in view and whose positions are widely-enough spaced to provide a good navigation solution. The satellites chosen by the selector 118 are indicated by arrow 120 in Figure 16. The satellite selector 118 bases its selection on satellite position information from the almanac memory 132 and a present-time signal 126 from the local clock 124. The local clock 124 is connected to a temperature-compensated crystal oscillator ("TXCO") 125. All the internal

15 timing signals used in the GPS satellite are derived from the TXCO 125. The selector 1 18 then passes the satellite's identity to the code generator 122, which generates the appropriate code for the desired satellite. This code is shifted in phase with respect to the local clock 124 until it matches the received signal code as determined by the active detector 112. This condition is maintained by a code tracking control loop 123.

Since the GPS satellites are moving rapidly, the satellite selector 118 also calculates the frequency offset

20 1 16 of each satellite due to Doppler shift, and adjusts a local oscillator 1 14 to place the satellite's signal in the receiver's passband.

The detected signal from the satellite is a bitstream which contains a fragment of the almanac 132, the satellite's ephemerides 134, a health bit 128 and another administrative information, and very precise time marks. This bitstream is separated into its components by a signal decoder 130.

25 A GPS satellite orbit calculator 136 uses the almanac 132 and ephemeride 134 information and Kepler's equations to calculate the position 138 of each chosen GPS satellite in its orbit with a precision of a few meters. This information is transmitted to the satellite selector 118 and to the navigation microprocessor 140.

The navigation microprocessor 140 receives a timing signal from the local clock 124 and a timing signal from each satellite for each chosen satellite. Since the speed of radio signals is a known constant, the navigation

30 computer 140 can calculate the distance to each chosen satellite. Because the local clock 124 is not highly accurate, these ranges are in error, and are called "pseudo-ranges." These pseudo-ranges are measured utilizing the code phase shift 127 provided by the code generator 122 over loop 123.

If four satellites were chosen by the selector 118, the navigation computer 140 has four pseudo-ranges. If, however, the ranges were accurate, three would be sufficient to find an exact position. To find an exact

35 solution, the computer solves a system of four equations in four unknowns: latitude 142, longitude 146, altitude 144 and the local clock error. The result is a position 142, 144 & 146 and precise time 148 output. The algorithms used in these calculations make allowance for the Earth's rotation, so that the position found is a point on the Earth's surface. In a commercial GPS receiver, positions are generated frequently, usually at intervals of a second or less. The receiver utilizes the most recent two or three positions and their associated times to calculate speed and direction of motion with respect to the Earth's surface. This task is accomplished in the speed and course computer 150. The speed and course output 152 can be combined with knowledge of the orientation of the body of the satellite to its direction of motion, which is derived from an attitude computer 154, to determine the appropriate pointing angle 156 of the antenna 34.

Alternatively, two GPS receivers 100 and 158 can be employed, separated by several feet on the satellite's structure. The satellite's attitude and any rotation of its structure can be then be determined by comparing the positions of the two receivers 100 and 158.

Alternative Methods of Pointing the Satellite Antenna

In an alternative embodiment of the invention, the satellite 12 is furnished with information describing its velocity with respect to the Earth's surface from an Earth-based radar system. This information can be combined with information from the satellite's attitude-sensing equipment to generate an antenna pointing signal. Another alternative utilizes circuits aboard the satellite 12 that generate a continuous relative- velocity signal. An on-board computer is used to continuously solve Kepler's equations for the satellite's orbit and adds the Earth's motion, derived from the latitude of the nadir as calculated above. The computer is updated periodically by an Earth-based signal which provides the ephemerides of the satellite's orbit from radar or visual observations. Yet another alternative incoφorates a method that is useful for a satellite in a stabilized circular orbit.

A device onboard the satellite is programmed with the equation of its fixed orbit. The output signal of the device is the relative velocity and angle of the satellite's course over the Earth's surface. The device receives synchronizing pulses periodically (as a practical matter, as infrequently as monthly for some orbits) from radar or visual sightings. The resulting information, which for such an orbit repeats itself each orbit and which in principle can be generated by a mechanism as simple as a cam, can be used to point the satellite antenna.

Comparison with Prior Inventions

The dwell time of a terminal in an elongate spanning cell is on the order of two to eight times longer for the present invention than the time a terminal would reside in a cell of Jan et al. as shown in U.S. Patent No. 5,268,694. This comparison was calculated with reference to Jan's description in Column 3, Lines 55-60 and to Figure 2 for a footprint diameter = 4,075 km, a satellite velocity = 25,000 km/hr, approximately eight Jan cells along the footprint diameter, and a cell diameter of approximately 4,075/8 = 500 km. The time of dwell of a Jan et al. cell on a terminal is 500 x 60/25,000 = 1.2 minutes. Using the same orbital parameters, the time of dwell of cell of the present invention on a terminal is 4,075 x 60/25,000 = 10 minutes. The cells of the present invention when divided into four elongate segments, have a dwell time of 10/4 = 2.5 minutes. A comparison of time of dwell with the cell of the present invention and Rouffet is not meaningful, since Rouffet's beams are made to hop around within the footprint to prevent interference between user terminals operating on the same frequency. CONCLUSION

Although the present invention has been described in detail with reference to particular preferred and alternative embodiments, persons possessing ordinary skill in the art to which this invention pertains will appreciate that various modifications and enhancements may be made without departing from the spirit and scope of the Claims that follow. The various orbital parameters and satellite designs that have been disclosed above are intended to educate the reader about particular embodiments, and are not intended to constrain the limits of the invention or the scope of the claims. The List of Reference Characters which follows is intended to provide the reader with a convenient means of identifying elements of the invention in the Specification and Drawings. This list is not intended to delineate or narrow the scope of the Claims.

INDUSTRIAL APPLICABILITY

The Cellular Satellite System described above will enhance the ability of satellite communication systems, particularly those availing satellite fixed cell techniques, to carry higher message traffic volumes from a greater number of users. It can be seen that by pointing the satellite radiation interface 14 to maintain the elongate spanning cells 18 in alignment with the satellite's track T' over the Earth's surface, a terminal P,M,G in the satellite's track T' remains (dwells) within an elongate spanning cell 18 for approximately the entire time the satellite footprint 16 passes over the terminal P,M,G. The dwell time of a terminal P,M,G within the track-aligned, elongate spanning cell 18 is many times longer than the dwell time of a terminal P,M,G in a hexagonal, circular or elliptical shaped satellite-fixed cell created by known satellites in comparable orbits. Even when the elongate spanning cell 18 is subdivided into multiple linear segments 22, a terminal P,M,G dwells for more than two times longer than in a satellite-fixed cell created by known satellites in comparable orbits. This feature saves time to access a satellite by reducing the number of hand-offs as the satellite-fixed cells 16 sweep over a terminal P,M,G. Saving time augments frequency re-use, and thereby enhances the communications and sensing capacity of the satellite 12 resulting in economic benefit to the users and proprietors of the system.

LIST OF REFERENCE CHARACTERS

Constellation Orbit Satellite Radiation interface a, 15b, 15c Radiated energy a, 16b, 16c Footprint formed on Earth a, 17b, 17c Reception zone Elongate spanning cells Longitudinal axis of elongate spanning cell Multiple elongate segments Corrected position of longitudinal axis Satellite body Solar array Parabolic reflector Inertia wheel Antenna positioning servo Radio-frequency feed to antenna Altitude and latitude signal Converter Control signal Radio receiver Synchronizing pulse Timers Elapsed time code Converter Control signal Attitude sensor Angle of satellite longitudinal axis with Earth's axis Satellite velocity vector Latitude Microprocessor Beam offset angle Star tracker Equator-crossing pulse Timer 78 Time since crossing

80 Function generator

82 Angle of satellite longitudinal axis with Earth's axis

84 Adder

86 Beam offset angle

100 Global Positioning Satellite receiver

102 Receiver antenna

104 Low noise amplifier

106 Mixer

108 Bandpass filter

110 Intermediate frequency amplifier

1 12 Active detector

114 Local oscillator

1 16 Doppler offset

1 18 Satellite selector

120 Chosen satellite number

122 Code generator

123 Code tracking control loop

124 Local clock

125 Temperature-compensated crystal oscillator

126 Time signal

127 Code phase shift

128 Health information

130 Signal decoder

132 Almanac memory

134 Ephemeride memory

136 GPS satellite orbit calculator

138 Positions of all GPS satellites

140 Navigation microprocessor

142 Latitude

144 Altitude

146 Longitude

148 Precise time

150 Speed and course computer

152 Speed and course

154 Attitude computer

156 Antenna angle 158 Second receiver

200 Graph of correction angle versus latitude

202 Correction angle

204 Argument of perigee

206 Latitude

108 Graph of latitude of nadir versus argument of perigee

210 Latitude of nadir

212 Argument of perigee

214 Graph of surface velocity of Earth versus North or South latitude

216 Surface velocity of the Earth

218 North or South latitude

E Earth

F Fixed terminal

G Gateway

GC Great circle

M Mobile terminal

N Nadir

NP North Pole

P Portable terminal

SP South Pole

T Direction of travel of satellite

T Track of satellite on ground φ Compensation angle

Claims

CLAIMSWhat is claimed is:

1. A method for controlling at least the transmission of a beam of radiated energy (15a, 15b, 15c) to a terminal (P,M,F,G) from a position above the Earth (E) comprising the steps of:

operating a satellite (12) in an Earth orbit (11);

said satellite (12) flying in a direction (T) having a track (T') with respect to a point on the Earth's surface while flying in a direction (T) at an altitude above the Earth (E) which is not an

Equatorial geosynchronous altitude;

said satellite (12) having a radiation interface (14);

said radiation interface (14) being capable of transmitting said beam of radiated energy (15a, 15b, 15c) to said terminal (P,M,F,G);

forming a footprint (16a, 16b, 16c) using said radiation interface (14) which is capable of transmitting said beam of radiated energy (15a, 15b, 15c) to said terminal (P,M,F,G);

partitioning said footprint (16a, 16b, 16c) into at least one elongate spanning cell (18);

said elongate spanning cell (18) extending along entire said footprint (16a, 16b, 16c); and having a longitudinal axis (20); and

directing said beam of radiated energy (15a, 15b, 15c) to said elongate spanning cell (18) in said footprint (16a, 16b, 16c);

characterised in that the method comprises aligning said longitudinal axis (20) of said elongate spanning cell (18) continually generally parallel to said track (T') of said satellite (12) on the Earth's surface.

2. A method for controlling at least the reception of radiated energy (15a, 15b, 15c) at a position above the Earth (E) comprising the steps of:

operating a satellite (12) in an Earth orbit (11);

said satellite (12) having a track (T) with respect to a point on the Earth's surface while flying in a direction (T) at an altitude above the Earth (E) which is not an Equatorial geosynchronous altitude;

said satellite (12) having a radiation interface (14);

said radiation interface (14) being capable of receiving radiated energy (15a, 15b, 15c);

mapping a reception zone (17a, 17b, 17c) on the Earth's surface using said radiation interface (14);

said reception zone (17a, 17b, 17c) containing all positions from which radiated energy (15a,

15b, 15c) may be received by said radiation interface (14) aboard said satellite (12) at a particular instant of time;

partitioning said reception zone (17a, 17b, 17c) into at least one elongate spanning cell (18);

said elongate spanning cell (18) extending along entire said reception zone (17) and having a longitudinal axis (20); and

directing said beam of radiated energy (15a, 15b, 15c) to said elongate spanning cell (18) in said reception zone (17a, 17b, 17c);

characterised in that the method comprises aligning said longitudinal axis (20) of said elongate spanning cell (18) continually generally parallel to said track (T') of said satellite (12).

3. A method in accordance with claim 1 or 2, wherein each elongate spanning cell (18) has two long sides substantially parallel to said longitudinal axis (20) whereby to constitute a linear spanning cell.

4. A method in accordance with any preceding claim, further comprising the step of:

dividing said spanning cell (18) into multiple segments (22) mutually located along the direction of said longitudinal axis (20); and

directing said beam of radiated energy (15a, 15b, 15c) to each of said multiple segments (22) in said spanning cell (18).

5. A method in accordance with any preceding claim, in which the step of continually aligning said longitudinal axis (20) of said spanning cell (18) includes:

correcting the alignment of said spanning cell (18) by shifting the position of said longitudinal axis (20) to a corrected position (24) by an angle having a value ╧å:

said value ╧å being measured between the projection of the track (T') of said satellite (12) on the Earth's surface and a Great Circle (GC) on the Earth's surface;

said Great Circle (GC) containing a projection of the direction (T) of the satellite orbit (1 1) and the nadir (N) of the position of said satellite (12);

such that the alignment of said spanning cell (18) is compensated for the changing relative motion of the Earth's surface at various latitudes.

6. A method in accordance with claim 1 or 2, further comprising the step of:

configuring the position of said radiation interface (14) to be continually generally aligned parallel to the relative motion of the Earth's surface, the track (T') of said radiation interface (14) over the surface of the Earth being generally a spiral.

7. A system for controlling at least the transmission of a beam of radiated energy (15a, 15b, 15c) from a satellite to a terminal (P,M,F,G) from a position above the Earth (E) comprising:

a satellite (12) in an Earth orbit (11);

said satellite (12) flying in a direction (T) having a track (T) with respect to a point on the Earth's surface while flying in a direction (T) at an altitude above the Earth (E) which is not an

Equatorial geosynchronous altitude;

said satellite (12) having a radiation interface (14);

said radiation interface (14) being capable of transmitting said beam of radiated energy (15a, 15b, 15c) to said terminal (P, M, F, G); said radiation interface (14) having a footprint (16a, 16b, 16c) on the Earth's surface;

said terminal (P, M, F, G) lying within said footprint ( 16a, 16b, 16c);

said satellite (12) also having beam control means for partitioning said footprint ( 16a, 16b, 16c) into at least one elongate spanning cell (18);

said elongate spanning cell (18) extending along entire said footprint (16a, 16b, 16c) and having a longitudinal axis (20);

means for directing said beam of radiated energy (15a, 15b, 15c) to said elongate spanning cell (18) in said footprint ( 16a, 16b, 16c); and

characterised in that the system comprises means for aligning said longitudinal axis (20) of said elongate spanning cell (18) continually generally parallel to said track (T') of said satellite (12) on the Earth's surface.

8. A system according to claim 7, wherein said beam control means operates such that each elongate spanning cell (18) has two long sides substantially parallel to said longitudinal axis (20) whereby to constitute a linear spanning cell.

9. The satellite system in accordance with claim 7 or 8, comprising means for further dividing said elongate spanning cell (18) into multiple segments (22) mutually located along the direction of said longitudinal axis (20);

and means for directing said beam of radiated energy (15a, 15b, 15c) to said terminals (P, M, F, G) within each of said multiple segments (22) in said spanning cell (18).

10. The satellite system in accordance with any of claims 7 to 9, in which:

said means for aligning operates by shifting the position of said longitudinal axis (20) to a corrected position (24) by an angle having a value ╧å;

11. The satellite system in accordance with claim 10, in which the angle ╧å is the angle of which the value is 90 degrees minus orbital inclination angle (i) minus arctangent surface velocity (V_s) at a given latitude (L(╧ë)) divided by orbital velocity (VJ.

12. The satellite system in accordance with claim 7, in which means are provided for continually generally aligning the position of said radiation interface (14) parallel to the relative motion of the Earth's surface, the track (T) of said radiation interface (14) on the surface of the Earth generally forming a spiral thereon.

13. A system for controlling at least the reception of a beam of radiated energy (15a, 15b, 15c) to a satellite from a terminal (P,M,F,G) from a position above the Earth (E) comprising:

a satellite (12) in an Earth orbit (11);

Equatorial geosynchronous altitude;

said satellite (12) having a radiation interface (14);

said radiation interface (14) being capable of receiving said beam of radiated energy (15a, 15b, 15c) from said terminal (P,M F,G); said radiation interface (14) having a footprint (16a, 16b, 16c) on the Earth's surface;

said terminal (P,M,F,G) lying within said footprint (16a, 16b, 16c);

said satellite (12) also having beam control means for partitioning said footprint (16a, 16b, 16c) into at least one elongate spanning cell (18);

means for pointing toward said beam of radiated energy (15a, 15b, 15c) received from said terminal (P,M,F,G) in said elongate spanning cell (18) in said footprint (16a, 16b, 16c); and

14. A system according to claim 13, wherein said beam control means operates such that each elongate spanning cell (18) has two long sides substantially parallel to said longitudinal axis (20) whereby to constitute a linear spanning cell.

15. The satellite system in accordance with claim 13 or 14, comprising means for further dividing said elongate spanning cell (18) into multiple segments (22) mutually located along the direction of said longitudinal axis (20);

and means for pointing toward said beam of radiated energy (15a, 15b, 15c) received from said terminals (P,M,F,G) within each of said multiple segments (22) in said spanning cell (18).

16. The satellite system in accordance with any of claims 13 to 15, in which:

said Great Circle (GC) containing a projection of the direction (T) of the satellite orbit (11) and the nadir (N) of the position of said satellite (12);

17. The satellite system in accordance with claim 16, in which the angle ╧å is the angle of which the value is 90 degrees minus orbital inclination angle (i) minus arctangent surface velocity (V_s) at a given latitude (L(╧ë)) divided by orbital velocity (VΓÇ₧).

18. The satellite system in accordance with claim 13, in which means are provided for continually generally aligning the position of said radiation interface (14) parallel to the relative motion of the Earth's surface, the track (T') of said radiation interface (14) on the surface of the Earth generally forming a spiral thereon.