U.S. patent application number 11/421794 was filed with the patent office on 2006-12-07 for system and method for implementing a constellation of non-geostationary satellites that does not interfere with the geostationary satellite ring.
This patent application is currently assigned to Virtual Geosatellite LLC. Invention is credited to Jack Anderson, David Castiel, John E. Draim.
Application Number | 20060276128 11/421794 |
Document ID | / |
Family ID | 26850388 |
Filed Date | 2006-12-07 |
United States Patent
Application |
20060276128 |
Kind Code |
A1 |
Castiel; David ; et
al. |
December 7, 2006 |
System and method for implementing a constellation of
non-geostationary satellites that does not interfere with the
geostationary satellite ring
Abstract
Provided is an improved system and method for implementing a
constellation of satellites in inclined elliptical orbits. The
satellites are operated during the portion of their orbits near
apogee to emulate the characteristics of geostationary satellites.
The orbits are configured to form a number of closely spaced
repeating ground tracks around the earth. In each ground track the
satellites operate only in arcs well above or below the equator to
provide a large number of non-geostationary orbital slots that
substantially increase global satellite capacity without
interfering with the existing geostationary satellite ring. Minimum
spacing is maintained between satellites in each active arc and
between satellites in the active arcs of adjacent ground tracks to
ensure that the satellites in the non-geostationary constellation
do not interfere with each other.
Inventors: |
Castiel; David; (Washington,
DC) ; Anderson; Jack; (Potomac, MD) ; Draim;
John E.; (Vienna, VA) |
Correspondence
Address: |
DORT PATENT CORPORATION
1700 Diagonal Road, Suite 300
Alexandria
VA
22314
US
|
Assignee: |
Virtual Geosatellite LLC
Washington
DC
|
Family ID: |
26850388 |
Appl. No.: |
11/421794 |
Filed: |
June 12, 2006 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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09658215 |
Sep 8, 2000 |
6954613 |
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11421794 |
Jun 12, 2006 |
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11095745 |
Sep 16, 2005 |
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11421794 |
Jun 12, 2006 |
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60153289 |
Sep 10, 1999 |
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Current U.S.
Class: |
455/12.1 |
Current CPC
Class: |
Y02D 30/70 20200801;
H04B 7/195 20130101 |
Class at
Publication: |
455/012.1 |
International
Class: |
H04B 7/185 20060101
H04B007/185 |
Claims
1. A satellite communications system, comprising: a ground station,
including communications equipment and an antenna, located at a
position on the earth; a plurality of satellites in orbits around
the earth having apogees and perigees, each of the satellites
having communications equipment thereon configured to communicate
with the ground station only during a predetermined portion of the
satellite's orbit proximate to apogee, the orbits of the plurality
of satellites being configured to form at least two ground tracks
on the earth displaced from each other longitudinally, each of the
ground tracks repeating daily and having a number of active arcs,
each active arc corresponding to the portion of the orbit of each
satellite during which the communications equipment on the
satellite is enabled to communicate with the ground station, the
orbits of the plurality of satellites being further configured such
that at all times there are at least two of the satellites in each
of the active arcs and such that at all times each of the
satellites in any one of the active arcs is separated by at least a
predetermined angle, as observed from the ground station, from each
other satellite in the same active arc and from any satellite in
any other active arc.
2. A system according to claim 1, wherein the orbit of each of the
plurality of satellites has a mean motion that is one of 2, 3 and
4.
3. A system according to claim 1, wherein the orbits of each of the
plurality of satellites is inclined at critical inclination.
4. A system according to claim 1, wherein the argument of perigee
of the orbits of each of the plurality of satellites is in the
range of 195 degrees to 345 degrees for apogees in the northern
hemisphere and in the range of 15 degrees to 165 degrees for
apogees in the southern hemisphere.
5. A system according to claim 1, wherein each of the plurality of
satellites has throughout its orbit an orbital height lower than a
height necessary for geostationary orbits.
6. A system according to claim 1, wherein the plurality of
satellites are equally spaced in mean anomaly within their
respective ground tracks.
7. A system according to claim 1, wherein the orbits of the
plurality of satellites are further configured such that the
portion of the orbits during which the communications equipment on
the satellites is enabled to communicate, is separated from the
equatorial plane of the earth by at least a predetermined
amount.
8. A system according to claim 1, wherein the communications
equipment on the plurality of satellites is further configured to
communicate at frequencies allocated to geostationary
satellites.
9. A system according to claim 1, wherein each of the plurality of
satellites has a power system configured to generate a first amount
of power when the communications equipment on the satellite is
enabled and a second amount of power more than the first amount of
power when the communications equipment is not enabled, to store
excess power generated when the communications equipment is not
enabled, and to enable the communications equipment with both the
stored excess power and the generated first amount of power.
10. A constellation of satellites, comprising: a plurality of
satellites in orbits around the earth having apogees and perigees,
each of the satellites having communications equipment thereon
configured to communicate only during a predetermined portion of
the satellite's orbit proximate to apogee, the orbits of the
plurality of satellites being configured to form at least two
ground tracks on the earth displaced from each other
longitudinally, each of the ground tracks repeating daily and
having a number of active arcs, each active arc corresponding to
the portion of the orbit of each satellite during which the
communications equipment on the satellite is enabled to
communicate, the orbits of the plurality of satellites being
further configured such that at all times there are at least two of
the satellites in each of the active arcs and such that at all
times each of the satellites in any one of the active arcs is
separated by at least a predetermined angle, as observed from the
earth, from each other satellite in the same active arc and from
any satellite in any other active arc.
11. A constellation according to claim 10, wherein the orbit of
each of the plurality of satellites has a mean motion that is one
of 2, 3 and 4.
12. A constellation according to claim 10, wherein the orbit of
each of the plurality of satellites is inclined at critical
inclination.
13. A constellation according to claim 10, wherein the argument of
perigee of the orbits of each of the plurality of satellites is in
the range of 195 degrees to 345 degrees for apogees in the northern
hemisphere and in the range of 15 degrees to 165 degrees for
apogees in the southern hemisphere.
14. A constellation according to claim 10, wherein each of the
plurality of satellites has throughout its orbit a orbital height
lower than a height necessary for geostationary orbits.
15. A constellation according to claim 10, wherein the satellites
in each of the two or more ground tracks are equally spaced in mean
anomaly.
16. A constellation according to claim 10, wherein the orbit of
each of the plurality of satellites is further configured such that
the portion of the orbits during which the communications equipment
on the satellites is enabled to communicate, is separated from the
equatorial plane of the earth by a least a predetermined
amount.
17. A constellation according to claim 10, wherein the
communications equipment on each of the plurality of satellites is
further configured to communicate at frequencies allocated to
geostationary satellites.
18. A constellation according to claim 10, wherein each of the
plurality of satellites has a power system configured to generate a
first amount of power when the communications equipment on the
satellite is enabled and a second amount of power more than the
first amount of power when the communications equipment is not
enabled, to store excess power generated when the communications
equipment is not enabled, and to enable the communications
equipment with both the stored excess power and the generated first
amount of power.
19. A method for satellite communications, comprising: orbiting a
plurality of communications satellites about the earth, the orbits
having apogees and perigees; and enabling each of the plurality of
communications satellites to communicate only during a
predetermined portion of the orbits proximate to apogee; wherein
the orbits of the plurality satellites form at least two ground
tracks on the earth displaced from each other longitudinally, each
of the ground tracks repeating daily and having a number of active
arcs, each active arc corresponding to the portion of the orbit of
each satellite during which the communications equipment on the
satellite is enabled to communicate; and wherein the satellites are
orbited such that at all times at least two of the satellites are
in each of the active arcs and such that at all times each of the
satellites in any one of the active arcs is separated by at least a
predetermined angle, as observed from the earth, from each other
satellite in the same active arc and from any satellite in any
other active arc.
20. A method according to claim 19, further comprising: configuring
the orbits of each of the plurality of satellites to have a mean
motion that is one of 2, 3 and 4.
Description
REFERENCE TO PRIORITY DOCUMENTS
[0001] This Patent Application is a continuation of, and claims
priority under 35 USC .sctn.120 to, co-pending U.S. patent
application Ser. No. 11/095,745, entitled "Fixed satellite
constellation system employing non-geostationary satellites in
sub-geosynchronous elliptical orbits with common ground tracks,"
filed Mar. 30, 2005, which is a continuation claiming priority
under 35 USC .sctn.120 to U.S. application Ser. No. 09/658,215,
entitled "Fixed satellite constellation system employing
non-geostationary satellites in sub-geosynchronous elliptical
orbits with common ground tracks," filed Sep. 8, 2000, now U.S.
Pat. No. 6,954,613, issued Oct. 11, 2005, which claims priority
under 35 USC .sctn.119(e) to Provisional Patent Application Ser.
No. 60/153,289, filed on Sep. 10, 1999, all of the above-reference
patent applications are incorporated by reference for all
purposes.
BACKGROUND
[0002] Satellite communications systems often require that a
station on the ground communicate with a satellite. The satellite
tracking is simplified when the satellite appears to be maintained
stationary relative to the Earth. Geosynchronous ("geo") satellites
have this characteristic. However, geo-satellites require high
altitude orbits. These high altitude orbits require large payloads
and launches, and also can have relatively long propagation delays
during communication.
SUMMARY
[0003] The present disclosure describes an array of
non-geostationary satellites in sub-geosynchronous, inclined
elliptical orbits. Each of the satellites communicates with a point
on the earth. At least a plurality of the satellites is in an
elliptical orbit with the earth at one focus of the ellipse.
[0004] At and near their apogee points, the satellites move slowly
relative to the Earth. These satellites appear virtually
geostationary to users within at least part of the desired coverage
area.
[0005] The disclosed embodiments use three sub-constellations, each
with 5 satellites. Three total sub-constellations are used. Two of
these sub-constellations are used for Northern Hemisphere
operation. A third constellation is for Southern Hemisphere
operation. The satellites are active over only part of their total
time of their orbits. The active time of the orbit is when the
satellites are closest to their apogees.
[0006] These active times can occur when the satellites are at
latitudes above 45.degree. These satellites are hence seen at high
elevations over much of their primary service areas.
[0007] This system is also effectively transparent to the
geostationary fixed satellite services and can be separated from
the geostationary arc preferably by at least 40.degree. at all
times within the service area of the system.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] FIG. 1 shows a basic layout of the multiple elliptical
orbits of the present invention;
[0009] FIG. 1A shows a graphical depiction of the satellite's
angular motion along its orbit as a function of the semi-major axis
of the elliptical orbit.
[0010] FIGS. 2A & 2B shows a block diagram of the satellite
communication equipment used according to the present
invention;
[0011] FIG. 2C shows a flowchart of operation of the satellites of
the present invention;
[0012] FIG. 3 shows the characteristics of a basic ellipse;
[0013] FIGS. 4A-4F show characteristics of the three-satellite
orbit of the present invention;
[0014] FIG. 4G shows characteristics of this orbit which prevent
interference with geosynchronous satellites in an inclined
orbit;
[0015] FIG. 4H shows characteristics of this orbit which prevent
interference with geosynchronous satellites in an equatorial
orbit;
[0016] FIGS. 5A-5E show characteristics of the five satellite orbit
of the present invention;
[0017] FIG. 6 shows an overall view of the ten satellite orbit of
the present invention;
[0018] FIGS. 7A-7G show the positions of the satellites of the ten
satellite embodiment within their repeating ground tracks;
[0019] FIG. 8 shows the operating elevation angles for the ten
satellite orbit, and their angular isolation from geo satellites;
and
[0020] FIG. 9 shows ground tracks of the preferred orbits.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
[0021] The disclosed system defines a communication system
including ground communication equipment and a special
constellation of satellites in elliptical orbits at lower altitudes
than those necessary for geosynchronous, which simulate to the
characteristics of a geosynchronous orbit from the viewpoint of the
ground communication equipment on the earth. The inventors
recognized that satellites which orbit in certain elliptical orbits
spend most of their time near the apogees of their orbits: the time
when they are most distant from the earth. These satellites spend
only a minority of their time near their perigee. For example, an
elliptical satellite in a 12-hour orbit spends eight of those hours
near its apogee. By appropriately choosing characteristics of the
satellite orbit, the satellite can be made to orbit, during that
time, at a velocity that approximates the rotational velocity of
the earth. The disclosed system defines a communication system
using a constellation of satellites chosen and operating such that
a desired point on the earth always tracks and communicates with a
satellite at or near apogee.
[0022] Another important feature of the disclosed system is the
recognition of how this mode of operation of the satellite changes
its power characteristics.
[0023] Geosynchronous satellites are used virtually 100% of the
time (except when in eclipse) and hence their power supplies must
be capable of full-time powering. This means, for example, if the
satellite requires 5 Kw to operate, then the power supply and solar
cells must be capable of supplying a continuous 5 Kw of power. The
satellites of the disclosed system, however, are not used 100% of
the time. During the perigee portions of the satellite orbit, the
satellites are typically not using most of their transmit and
receive capability and hence, the inventors recognized, do not use
a large part of their power capability.
[0024] The inventors of the disclosed system recognized this
feature of the satellites, and realized that the satellites could
be storing the power that is being produced during this time of
non-use. Therefore, the inventors realized that the size of the
power supply could be reduced by a factor of the percentage of time
that the satellite is not used.
[0025] The power sources can be any known means, including solar
cells, nuclear reactors, or the like. If the satellite is used half
the time, then the power source need only be sized to provide half
the power. At times when the satellite is not being used, the power
source provides power to a battery storage cell, which holds the
power in reserve for times when the satellite is being used.
[0026] Like geo systems, the satellite of the disclosed system is
virtually continuously in the same location. Unlike geo-based
systems, however, the ground communication equipment of the
disclosed system does not always communicate with the same
satellite. The satellites move slightly relative to the earth, i.e.
they are not always precisely at the same point in their apogees.
One important advantage of the disclosed system is that the one
satellite at apogee later moves to perigee, and still later to
apogees at other locations overlying other continents and areas.
Hence, that same satellite can later communicate with those other
areas. Therefore, this system allows a store-and-dump type system.
The information can be stored on board the satellite and later
re-transmitted when the satellite overlies those other areas. This
system also allows all the satellites in the array to communicate
with the other satellites in the constellation, through
intersatellite links. This feature is desirable for real time
communications.
[0027] This system has a number of other distinct advantages.
Importantly, the system operation allows selecting specific
geographic locations to be preferentially covered; for example,
continents can be followed by the constellation to the exclusion of
other areas, e.g. ocean areas between the continents. The
communication equipment on the continent always communicates with
one satellite at apogee, although not always the same satellite.
From the point of view of the ground station, the satellite appears
to hover over the ground. This satellite system operates virtually
like a geosynchronous satellite system. Importantly, these
satellites according to the disclosed system orbit at about half
the altitude of the geo systems. A geo orbit orbits at 36,000 miles
altitude: the virtual geo satellite orbits at average altitudes of
16-18,000 miles. Also, geo satellites require "apogee motors", to
boost them from their original orbits into the final geo orbit.
These apogee motors can double the weight of the satellite. This
yields a communications system which costs less dollars per launch
capability because of the reduced weight to boost and less size.
Also, since the geo satellites orbit at a higher altitude, they
operate at a higher power, and use a larger illuminating antenna,
all other conditions on the around being equal. These satellites
also have a much larger overall size. This size of the satellites
increases as the square of the distance. Therefore, the geo
satellite needs to be at least twice as large and twice as powerful
as a low altitude satellite. The power supply conservation
techniques of the disclosed system allow the satellite to be made
even smaller.
[0028] The system also provides satellites with very high elevation
angles. Maximizing the elevation angle prevents interference with
existing satellites such as true geosynchronous satellites.
[0029] This is another feature of the disclosed system which allows
these satellites to operate in ways which avoid any possibility of
interference with the geo band. Another objective and important
feature of the disclosed system is its ability to re-use satellite
communication channels. Regulatory agencies such as the FCC
allocate frequency bands by allocating a specific frequency band
for a specific purpose. The geo satellites, for example, receive an
assignment of a frequency band. Thereafter, the regulatory agency
will consider that other satellites located in the same orbital
position can not use this frequency because of possibility of
interference. Hence, frequencies in adjacent bands which might
interfere with that assigned band will not be allocated for new
satellite use. With the disclosed system, there is a large angular
separation between the geo-sats and those covered by the invention.
Thus, the same frequencies ca be allocated anew. Another feature of
the disclosed system is the location of the earth stations and
satellites in a way which prevents interference with the geo bands.
Specifically, the disclosed system defines embodiments using both
inclined orbits and non-inclined (equatorial) orbits. The inclined
orbit embodiment of the disclosed system only communicates with the
ground stations when a line drawn between the ground station and
current position of the satellite will not intersect any point
within x.degree. of the ring of geosynchronous satellites, where x
is the required separation between the communication for geo
satellites and the communication for the satellites of the
disclosed system. During other times, the equatorial component of
the communication is shut off. The satellite only communicates when
it is near apogee. During those times, the rotational velocity of
the satellite approximates the rotational velocity of the earth,
and hence the satellite tends to hang overhead relative to the
earth.
[0030] For non-inclined (equatorial) orbits, the ground stations
are placed in a position such that the communication does not
intersect the ring of equatorial orbits, by ensuring that satellite
apogees are at lower altitudes than apogees of geostationary
satellites.
[0031] The system is controlled by on-board processor 280, which
determines the position in the orbit and the steering of the
antenna from various parameters. Processor 280 carries out the
flowchart shown in FIG. 2a which will be described herein. The
overall system is powered by power supply 290 which supplies power
to all of the various components and circuitry which require such
power. Power supply 290 includes a source of power, here shown as a
solar array 292, and an energy storage element here shown as a
battery array 294. Importantly, according to the disclosed system,
the solar array 292 is sized to provide only some amount of power
less than that required to power the satellite communication. The
amount by which the solar array can be less is called herein the
power ratio of the device. The power ratio depends on the kind of
orbit that the satellite will have, and how long the satellite will
be transmitting during each elliptical orbit. The preferred power
ratio is 0.5: this will power a satellite which is communicating
half the time, and the other half the transmitter and receiver on
board the satellite is off and the solar array is providing power
to charge battery 294.
[0032] The flowchart of operation is shown in FIG. 2a. Step 350
represents controlling the antenna. This requires that the
processor keep track of the satellite's position in the orbit. Step
352 determines if the satellite is in a position in its orbit where
it is active (transmitting and/or receiving). If so, flow passes to
step 354 where power is drawn from power supply and the battery. If
the satellite is not powered, then power is used to charge the
battery at step 356.
[0033] The system also allows selective expansion of the
communications coverage by adding additional satellites into
additional elliptical orbits.
[0034] The virtual geo satellite system of the disclosed system
also enables complete communications coverage of the earth without
requiring a ground network. The same satellite services all
different portions of the earth at different times of day. The
coverage of the earth repeats over a 24 hour period. A preferred
embodiment receives information relayed from the ground, relays it
to the earth area below it, then stores the information, and later
reads back the stored information to retransmit that same
information to other areas of the earth. The system of the
disclosed system increases the satellite coverage at high density
geographic locations using fewer satellites than was possible with
previous constellations by fixing the satellite apogee passages
over given geographic regions defined by both longitude and
latitude.
[0035] Integral values for mean motion of the satellites in the
array ensures that the ground track repeats on a daily basis. The
ground tracks preferably repeat each day so that the orbit apogee
passes in the same location relative to the geographic target area.
This system maximizes the time of coverage and elevation angles for
that pass.
[0036] Before describing the minimum satellite arrangement
according to the disclosed system, the nomenclature used herein to
describe the characteristics of satellite orbits will be first
described. The "mean motion" is a value indicating the number of
complete revolutions per day that a satellite makes. If this number
is an integer, then the number of revolutions each day is uniform.
This means that the ground tracks of the satellites repeat each
day: each ground track for each day overrides previous tracks from
the preceding day.
[0037] Mean motion (n) is conventionally defined as the hours in a
day (24) divided by the hours that it takes a satellite to complete
a single orbit. For example, a satellite that completes an orbit
every three hours ("a 3-hour satellite") has a mean motion of
8.
[0038] The "elevation angle" .delta. is the angle from the
observers horizon up to the satellite. A satellite on the horizon
would have 0.degree. elevation while a satellite directly overhead
would have 90.degree. elevation. Geo satellites orbit near the
equator, and usually have a 20-30.degree. elevation angle from
points in the United States.
[0039] The "inclination" I is the angle between the orbital plane
of the satellite and the equatorial plane. Prograde orbit
satellites orbit in the same orbital sense (clockwise or
counter-clockwise) as the earth. For prograde orbits, inclination
lies between 0.degree. and 90.degree. Satellites in retrograde
orbits rotate in the opposite orbital sense relative to the earth,
so for retrograde orbits the inclination lies between 90.degree.
and 180.degree.
[0040] The "critical inclination" for an elliptical orbit is the
planar inclination that results in zero apsidal rotation rate. This
results in a stable elliptical orbit whose apogee always stays at
the same latitude in the same hemisphere. Two inclination values
satisfy this condition: 63.435.degree. for prograde orbits or its
supplement 116.565.degree. for retrograde orbits.
[0041] The "ascending node" is the point on the equator where the
satellite passes from the southern hemisphere into the northern
hemisphere. The right ascension of the ascending node ("RAAN") is
the angle measured eastward in the plane of the equator from a
fixed inertial axis in space (the vernal equinox) to the ascending
node.
[0042] The "argument of perigee" is a value that indicates the
position where orbital perigee occurs. When using equatorial
orbits, 00 argument of perigee is used for all the orbits. Inclined
orbit arrays use non-zero arguments of perigee. Arguments of
perigee between 0.degree. and 180.degree. locate the position of
perigee in the northern hemisphere and hence concentrate the
coverage in the southern hemisphere. Conversely, arguments of
perigee between 180.degree. and 360.degree. locate the perigees to
the southern hemisphere and hence concentrate the coverage on the
northern hemisphere.
[0043] An embodiment of the disclosed system evenly spaces the axes
of the ellipses. The spacing between RAANs is called "S" and
calculated by S=360/n=120.degree.
[0044] The disclosed system positions the satellite coverage based
on both longitude and latitude of the desired continental area to
be covered by the orbit. This is done, first, by synchronizing the
orbit apogee to pass over the targeted geographical region for each
successive satellite. We select a suitable value for the mean
anomaly, which is a fictitious angle relating to the elapsed time
in orbit. 360.degree. represents the completion of the orbit. In
this example, the mean anomalies are also S=120.degree. apart.
[0045] This "mean anomaly" M relates the amount of time it takes
the satellite to rotate S.degree. around the earth (here
120.degree.). The mean anomaly required for the 12-hour satellites
to rotate to S.degree. is 8 hours; two-thirds of a period. This
corresponds roughly to the amount of time the satellite remains in
apogee. Taking the initial satellite near apogee, therefore,
(180.degree. mean anomaly) the next satellite should be backed up
by 240.degree. This means that after 8 hours that satellite will be
at 180.degree. Since 180.degree. minus 240.degree. is negative
60.degree. which equals 3000, this is the value of mean anomaly M
for satellite number 2. This system is used to select values for
the constellation in a similar manner for each succeeding
satellite.
[0046] Arrays with more satellites ("higher order arrays") can also
be made using the same rules as those discussed above. Successively
larger numbers of satellites can be used to provide more coverage,
more overlapping coverage, or smaller integral mean motion values.
As the values of M get larger, the eccentricity of the ellipses
become smaller. This is because the perigee altitude is fixed at
about 500 km to avoid re-entry and decay into the earth's
atmosphere; longer periods have higher apogee altitudes greater
supportable eccentricities.
[0047] FIG. 1A shows how the satellite ellipse is selected to have
an angular rate in the plane of the equator, at apogee, which
approximates the angular rate of the earth. The dotted line in FIG.
1A represents the angular rate of a geo satellite, and hence at
this angular rate a satellite would approximate the angular speed
of the earth. The ellipse is selected to have a semi-major axis
length to set the minimum angular rate of the satellite at apogee.
At apogee, the satellite angular rate should approximate the
rotational velocity of the earth. In reality, this rotational
velocity will be either a little faster or a little slower than the
earth. At this time, therefore, the satellite appears to hang
relative to the earth. All elliptical orbits, including whose
described herein, are also subject to effects of long-term
perturbations. If effects of these long term perturbations are not
compensated, this could cause continental coverage to drift with
the passage of time.
[0048] These perturbation effects are mainly effects from the
Earth's J2 rotation harmonic. The earth is not a perfect sphere; it
actually bulges at the equator. This causes gravitational effects
on objects which orbit the earth. For posigrade orbits
(i>90.degree.) the line of nodes will regress. For inclinations
greater than critical (63.4.degree.>i>116.6), the line
between the perigee and apogee (line of apsides) will regress; for
other inclinations, I<63.4.degree. or I>116.6, the line of
apsides will progress. Exactly at the critical angles I=63.4 or
I=116.6, the line of apsides will remain stable a very desirable
feature in maintaining apogee at a certain latitude. In the
equatorial plane, the combined effect of these two major
perturbations cause the apogee to advance or move counter-clockwise
from the sense of looking down from the celestial north pole. All
of the satellites in a given array design would be affected
similarly. Fortunately, this effect could be compensated by
slightly increasing the period of each satellite in the array by an
amount which offsets the J2 perturbation. This affects the system
by causing a point on the earth to take a slightly longer time to
reach the satellite's next apogee arrival point. This effect is
compensated by slightly increasing the satellite's period. The
advance of perigee is suppressed by setting the inclination at one
of the critical values.
[0049] A first embodiment of the invention uses N=3 satellites,
where N is the total number of satellites, preferably in the
equatorial plane, to cover N-1=2 continents. The rules for spacing
and phasing the satellites will be given in the general form that
can be used later for more complicated constellations or arrays.
The mean motion integer sets the minimum number of satellites in
the array and n.sub.c the number of continents that are followed.
Here n.sub.c=2 provides a satellite period equal to 12 sidereal
hours. N (the minimum number of elliptic satellites in the array)
is determined by using the relationship N=n.sub.c+1. Thus, N=3.
This is the minimum number of satellites that need to be in the
array; we can also set the number of satellites in the array N to
be any integer greater than n+1.
[0050] The apogee passage is synchronized over the targeted
geographical region, for each successive satellite, moving
counterclockwise as viewed from the celestial North Pole. This is
accomplished by selecting a suitable value for the mean
anomaly.
[0051] Refinements: Additional features augmenting the usefulness
of the above simpler version include:
[0052] 1) Inclining the elliptical orbital planes at the critical
inclination angles (63.435 or 116.535.degree.), with phasing to
maintain a single repeating ground track. The single repeating
ground track for the simplified non-inclined example above is
simply the line of the equator.
[0053] 2) Taking advantage of the higher apogees in allowing more
direct cross-linking between satellites than with present
low-altitude circular arrays. Usually, a single cross-link
suffices, even when the longitude difference between end points is
180.degree. (on the opposite side of the earth).
[0054] 3) Placement of apogees over a selected latitude and
longitude for optimal coverage of a potential market area. This is
done through proper selection of all the orbital parameters, with
particular attention given to selection of argument of perigee,
.omega.
First Embodiment
[0055] The orbits of the disclosed system are shown in FIG. 1. The
satellite 100 is shown in an elliptical orbit 102 around the earth.
The communication equipment on the satellite 100 communicates with
earth ground station 104, and also beams the information to earth
ground station 106. Satellite 110 is shown in a separate
independent elliptical orbit communicating with ground stations 112
and 114 on the earth. Note also that the satellite 100 can
communicate directly to the satellite 110 via communication link
120.
[0056] The preferred characteristics of these orbits are described
in Table I.
[0057] TABLE I Satellite No. P1 P2 P3 Semi-Major 26553.98 km
26553.98 km 26553.98 km Axis, a=Inclination, I=0 deg 0 deg 0 deg
Arg. Perigee, 270 deg 270 deg 270 deg w=Eccentricity, 0.51 0.51
0.51 e=Rt. Ascension, 0 deg 120 deg 240 deg RAAN=Mean 180 deg 300
deg 60 deg Anomaly, MA=Satellite 100 also includes store and dump
hardware thereon as described herein. This allows the satellite to
obtain program information so that later in its orbit, when at the
position 130, it can send its same information to ground station
132.
[0058] A detailed block diagram of the electronics in the satellite
is shown in FIG. 2. This block diagram shows elements which carry
out communication between the ground station 104, the satellite
100, and the remote user station 106. The inter-satellite links 120
are shown from the satellite 100 to the satellite 110. The video
input to be distributed is received as video input 200, and input
to a video coder 202 which produces digital coded video
information. This digital coded video is multiplexed with a number
of other channels of video information by video multiplexer 204.
The resultant multiplexed video 206 is modulated and appropriately
coded by element 208 and then up-converted by transmitter element
210. The up-converted signal is transmitted in the Ku band, at
around 14 GHz, by antenna 212. Antenna 212 is pointed at the
satellite 100 and received by the satellite's receive phased array
antenna 214. Antenna 212 is controlled by pointing servos 213.
[0059] The received signal is detected by receiver 216, from which
it is input to multiplexer 218. Multiplexer 218 also receives
information from the inter-satellite transponders 240.
[0060] The output of multiplexer 218 feeds the direct transponders
250, which through a power amplifier 252 and multiplexer 254 feeds
beam former 256. Beam former 256 drives a transmit, steerable
phased-array antenna 260 which transmits a signal in a current geo
frequency band to antenna 262 in the remote user terminal 106. This
signal preferably uses the same frequency that is used by current
geo satellites. The phased array antenna is steered by an on-board
computer which follows a pre-set and repeating path, or from the
ground. This information is received by receiver 264, demodulated
at 266, and decoded at 268 to produce the video output 270.
[0061] The satellite includes another input to the multiplexer from
the steerable antenna, via the intersatellite link 120 and receiver
240. Transmit information for the the intersatellite link is
multiplexed at 242 and amplified at 246 prior to being
multiplexed.
[0062] Output 222 of input multiplexer represents a storage output.
The satellite electronics include the capability to store one hour
of TV program information. The TV channels typically produce
information at the rate of 6 megabytes per second. The channels are
typically digitally multiplexed to produce information on 4-6
channels at a time. Therefore, the disclosed system preferably uses
22 gigabytes of storage to store over 1 hour of information at
about 4.7 megabytes per second. The information stored will be
broadcast over the next continent. The storage unit 224,
accordingly, is a wide SCSI-2 device capable of receiving 4.7
megabytes per second and storing 22 GB.
[0063] Upon appropriate satellite command, the output of the
storage unit is modulated and up-converted at 226.
[0064] This basic system shown in FIG. 2 can be used in one of the
preferred satellite arrays of the disclosed system. These arrays
will be discussed herein with reference to the accompanying
drawings which show the characteristics of these satellite
arrays.
[0065] This first embodiment uses a simplified 12-hour equatorial
plane satellite array n=2, N=3. The mean motion n of 2 means that
each satellite completes an orbit around the earth twice per
day.
[0066] An important enhancement of an N=3 case is obtained by
modifying the characteristics of the orbits so that the satellites
coalesce over the covered areas at the moments when satellite
coverage changes. The term coalesce as used herein means that as
one satellite moves out of range of the ground tracking, the next
satellite moves into range at that same position. In fact, the two
satellites come very close to one another at that point--within
1.degree. from the view of the satellite. This simplifies the
ground tracking, since the switchover between satellites does not
require much antenna movement.
[0067] FIGS. 4A-4F show the basic three-satellite "rosette" formed
by the three elliptical orbits. The earth 300 is located at one of
the foci of each of the three ellipses of the respective
satellites. Satellite 302 communicates with point 304 on the earth.
Satellite 302 orbits the earth in ellipse 306. The satellites 1, 2
and 3 respectively have ascending nodes of 0, 120 and 240, and
respectively have mean anomalies of 180, 300, and 60.
[0068] Similarly, satellite 310 orbits the earth in ellipse 312,
and satellite 320 orbits the earth in ellipse 322. Satellites 310
and 320 are both in a position to provide coverage to the second
covered continent area 314. Note that satellites 310 and 320 are in
their coalesced position--they are very close positionally, to one
another. Satellite 320 is moving away from apogee while satellite
310 is moving toward apogee. The tracking antenna is hence
commanded to switch between tracked satellites a the time when
satellites 310 and 320 are positionally very close, but having
adequate angular separation to avoid self-interference. According
to the disclosed system, this switchover occurs when the satellites
are within 5.degree. of each other.
[0069] The satellites all orbit in a counter-clockwise direction
relative to the sense shown in FIG. 4. The earth also orbits in the
counter-clockwise direction. The semi-major axes of the ellipses in
FIG. 4 are shown as axes 308, 314, and 316, respectively.
[0070] In order to describe these orbits, first the characteristics
of an ellipse will be described. FIG. 3 shows ellipse 400, having a
focus 402. The satellite orbits along the path of the ellipse 400,
with the center of the earth being at the focus position 402 ("the
occupied focus").
[0071] The apogee 404 and the perigee 406 of the orbits are defined
by the points on the ellipse which are farthest from and closest to
the focus of the ellipse, respectively. The amount of difference
between these distances define the eccentricity of the ellipse. The
semi-major axis 408 is defined as half of the long axis of the
ellipse. This semi-major axis runs through the two foci of the
ellipse, to split the ellipse into two halves. The two lengths
along the semi-major axis, from one edge of the ellipse to the
occupied focus of the ellipse are called the "radius of perigee"
and the "radius of apogee"; the latter being the longer.
[0072] As the eccentricity of an ellipse approaches zero, the
ellipse becomes less elliptical, eventually approaching a circle
(e=0) when the eccentricity is zero. The semi-major axis of a
circle is the radius of the circle.
[0073] The characteristics of the ellipse/object in elliptical
orbit are calculated as follows.
[0074] The apogee, r.sub.a=a.multidot.(1+ECC).
[0075] Perigee r.sub.p=a.multidot.(1-ECC).
[0076] A more eccentric ellipse (higher value of eccentricity ECC)
has a greater difference between the values P and R. Hence, such an
ellipse is less like a circle. The characteristics of the ellipse
are therefore determined as a function of its eccentricity.
[0077] The position of a satellite in orbit follows Kepler's laws
of motion which states that the orbiting element will sweep out
equal areas of the orbit in equal times. This results in the
satellite moving very rapidly when it is at an approaching perigee,
but very slowly when it reaches apogee. For a twelve hour
elliptical orbit, therefore, it can be seen that the satellite will
spend most of its time near apogee. The numbers on the ellipse of
FIG. 3 represent time indications of hours passed in a 12 hour
orbit, e.g., they indicate the number of hours since zero that have
elapsed in a 12 hour orbit.
[0078] The preferred ellipse for the 3-satellite elliptical orbit
has an eccentricity of about 0.51. This value best allows the
satellites to coalesce.
[0079] The earth rotates once in every 24 hour period, and hence
takes eight hours to rotate between the major axes of the three
equally spaced ellipses (120.degree. spacing); FIG. 4A shows the
point to be covered 304 is initially pointing directly towards
satellite 302 which is at apogee at time 0:00. As time passes, both
the satellite 302 and the earth will rotate.
[0080] As time passes, the satellites move from the position shown
in FIG. 4A. FIG. 4B shows the position one hour later at time 1:00.
Satellite P1 has moved away from apogee, although it has moved
relatively little. Satellite P2, on the other hand, is now moving
much more rapidly at this time, since it is approaching perigee,
while P3 is still near the apogee position.
[0081] An observer on or near the equator sees the nearest
satellite appear to climb in altitude from almost directly
overhead, towards apogee, all the while staying almost directly
overhead at an elevation angle of 80-90. The satellite is actually
rotating more slowly than the earth during this time: it is
appearing to move from east to west, rather than west to east as
most low or medium altitude satellites move in the sky.
[0082] FIG. 4C shows a view of the satellites one hour later at
time 2:00. The tracked locations 304 and 314 each still view a
satellite near its apogee position. Satellite P3 continues to move
towards apogee and hence appears to hang overhead. P1 is still
around apogee and thus also appears to hover.
[0083] FIG. 4D shows yet another hour later at time 3:00. P3 is
still at apogee, but P1 is approaching perigee. Notice that P2 is
coming out of perigee and approaching the coalescence point at
which P1 and P3 will cross paths. That crossing of paths is shown
in FIG. 4E, time 4:00, when P1 and P2 have coalesced in their
positions at the time when point 304 switches over between coverage
by satellite P1 and P2. At that time, the satellites are within
1.degree. of one another as viewed from the ground.
[0084] The above has described the satellite P1 moving from
directly overhead the point to be covered, to the point where
satellite P1 no longer covers the point to be covered. Therefore,
the satellite is transmitting for eight of the twelve hours of its
orbit; 2/3 of the time.
[0085] This cycle repeats. As the satellites continue to orbit,
different satellites take similar positions to those shown in FIGS.
4A-4E. FIG. 4F shows the cycle starting to repeat with satellite P2
moving toward apogee, satellite P1 moving toward perigee, and P3
hovering relative to the earth near its apogee.
[0086] FIGS. 4A-4F demonstrate the important features recognized by
the inventors of the disclosed system, whereby the satellites spend
most of their time at apogee. At the highest points of apogee, the
velocity of the satellite very nearly matches that of the earth,
and so the satellite appears to hang overhead. The satellite is
preferably tracked while its angular velocity differs from the
earth's angular velocity by 20% or less.
[0087] Importantly, the covered areas on the earth always see
either a satellite directly overhead or two satellites which are
very nearly directly overhead. FIGS. 4A-4F show how this system
actually appears to the communications point 304 to be virtually
geosynchronous. The communications point communicates with
different satellites at different times in the satellite orbit. The
communications point is always communicating with one
satellite.
[0088] The satellites follow repeating ground tracks, since the
cycle of satellite movement shown in FIGS. 4A-4F continually
repeats. Importantly, this allows the ground tracking antenna 212
to continually follow the same path, starting at a beginning point,
tracking the satellite, and ending at the coalesce point. After the
satellites coalesce as shown in FIG. 4A, the antenna begins its
tracking cycle. The inventors of the disclosed system have
optimized this system for preventing interference with geo
satellites.
[0089] Specifically, consider FIG. 4G which shows a multiplicity of
satellites in inclined elliptical orbits. The disclosed system
preferably operates to monitor satellites at and near their apogee
positions. The satellites near perigee are moving too rapidly, and
hence are not tracked. More generally, the system of the disclosed
system operates such that the satellites are only being used at
certain times during their orbits. In this embodiment, those
certain times are when the satellites are at apogee. Non
geosynchronous circular arrays are commonly used at present; they
are actually much less efficient, since with zero eccentricity they
spend a significantly greater time on the side of the earth away
from the populated continents. The arrays of the disclosed system,
on the other hand, spend most of the time at or near apogee over
the populated continents of interest, and a relatively small time
(at high angular velocities) passing through perigee in regions of
no commercial interest.
[0090] The satellites are only used when their geometry is such
that there is no possibility of the line of sight between the
ground station and the satellite interfering with the
geosynchronous band of satellites. This allows the satellite
communication to take place on the same communication frequency
band normally assigned to geosynchronous satellites.
[0091] Moreover, the disclosed system teaches that when the
satellites are not communicating, either because the satellites are
no longer at their tracked apogee portion and/or when the
satellites are in a region where they might interfere with
geosynchronous satellites, the main transmission is turned off.
During this time, the power supply is used to charge the battery.
This means that the power supply can be made smaller by some factor
related to the duty cycle of the satellite.
[0092] Another consideration is since the satellites only
communicate while near apogee, they are never eclipsed by the
earth. The satellites can always receive sunlight for solar
operation while transmitting and receiving.
[0093] For example, FIG. 4G shows satellites in orbit. In the
example given in FIG. 4G, the satellites are only tracked when they
are in the position of the orbit above the line 450. The only
possibility of interference with geo satellites comes when the
tracking beam is within 10.degree. to 30.degree. of the geo band.
So long as an angular separation greater than this amount is
maintained, there can be no interference. Therefore, the disclosed
system allows re-using the frequency bands which are usually
assigned to geosynchronous satellites in a position where
interference with the existing satellites can not occur.
[0094] The same rules are used to construct higher order arrays
with successively larger integer mean motions and hence shorter
periods. These arrays require a larger number of satellites, but
provide somewhat better coverage of the earth. Since more
satellites are used in these higher order arrays, each satellite
need spend a lesser amount of its time at apogee. This allows
orbits to be formed wherein the values of eccentricity are allowed
to become smaller as the mean motion increases. The ultimate limit
is atmospheric drag, which limits perigee altitudes to about 500
kilometers. This would correspond to a 1500 kilometer apogee
elliptical orbit with a resulting eccentricity of
(r.sub.a-r.sub.p)/(r.sub.a+r.sub.p) which is approximately 0.067.
This described orbit is not practical since its period is about 1
hour and 45 minutes which is not an integral value for the mean
motion. The next nearest value for mean motion would be n=14. The
n=14 orbit, however, would be so slightly elliptic that it would
not offer much advantage over the circular orbits.
[0095] Practically, those arrays having mean motions of 3, 4, 5, 6,
7 and 8 are most preferred according to the disclosed system. The
most preferred orbits according to this invention include the
three-satellite orbits, the four-satellite orbits, and the
five-satellite orbits. A particularly advantageous embodiment uses
two arrays of five satellite orbits.
[0096] As discussed above, all of these orbits include long-term
perturbations which would, if not compensated, cause the desired
continental coverage to drift off with the passage of time. The two
major perturbation effects are due to the earth's J.sub.2 harmonic;
and include:
[0097] Regression of the line of nodes (for posigrade orbits), and
Advance of perigee.
[0098] For inclined orbits, the advance of perigee can be
suppressed by setting the inclination, i, at either 63.435 or
116.565.degree.
[0099] The combined effect of these two major perturbations in the
equatorial plane, due to the J.sub.2 harmonic term has the net
effect of causing the apogee to advance in a counter-clockwise
direction looking down from the celestial North Pole.). All the
satellites in a given array design would be affected alike.
Fortunately, this effect can be compensated by increasing slightly
the period of each satellite in the array in a way such that the
earth takes a slightly longer time to reach the next satellite's
apogee arrival point. This is compensated by adding this extra time
to the satellites' periods. The exact amount will vary, and is a
function of a number of variables, including the orbital periods,
inclinations, and eccentricities. For inclined elliptic orbits (at
critical inclination angles), there will be no rotation of perigee
in either direction. However, there will be a regression of the
line of nodes which must be compensated by a small adjustment in
orbital period. This will cause the plane of the orbit to rotate
clockwise in the sense looking down from the North Pole. If that
happens, the satellite would pass over a selected meridian at a
slightly earlier time each day (or each repeat cycle), unless we
adjust the period of the satellite. In this case, we would shorten
the period of the satellite, which effectively `stretches` out the
trajectory ground trace and causes the ground track to repeat
exactly over the life of the satellite.
[0100] As described above, third order effects due to tesseral
terms may need to be compensated by small orbit maintenance
maneuvers using minuscule amount of fuel.
[0101] The preferred four-satellite array is shown in FIGS. 5A-5E.
This array shows four satellites used to track three continents.
These satellites orbit in elliptical orbits having an eccentricity
of 0.6. FIGS. 5B and 5D show the satellite coalescing which occurs
according to this embodiment.
[0102] FIG. 6 shows an overall view of the 10 satellite array; and
FIGS. 7A-7E show the ground tracks for a satellite array with 5
satellites having a period, T, equal to 6 hours. This array is
preferably used with two sets of five satellites, yielding a
ten-satellite, six hour constellation.
[0103] The preferred communications system uses a ten satellite
system, each having six hour orbits, and each optimized for users
in the Washington, D.C. area. This still, however, provides
coverage throughout the rest of the continental United States, and
the entire northern hemisphere as well as that part of the southern
hemisphere down to about 10 deg South latitude.
[0104] The system uses ten equally-spaced prograde satellite orbit
planes. All satellite orbits are at the `critical` inclination
angle of 63.435.degree. to prevent rotation of the line of
apsides.
[0105] The ground track is adjusted so as to pass directly over
Washington, D.C. by adjusting the right ascensions of all the
orbits while maintaining their equal spacing. The argument of
perigee is adjusted to obtain apogees over or nearly over the
targeted latitude and longitude.
[0106] FIG. 6 shows an overview of the orbital constellation. It
can readily be seen that the satellites favor the Northern
Hemisphere by spending more time, and reaching a higher altitude in
the Northern Hemisphere. FIG. 6 shows a snapshot of time at 0:00
hours, and it should be seen that all satellites except for
satellites P5 and P1 are over the Northern Hemisphere at that
time.
[0107] FIGS. 7A-7G show a Cartesian, or Mercator, plot of the world
showing the repeating ground tracks. The satellite array has a
repeating ground track that repeats every 24 hours. The satellites
appear to `hover` or dwell along four equally-spaced meridians, one
of which is at the longitude of Washington, D.C.; the others being
spaced at 90.degree. intervals from Washington.
[0108] FIG. 8 shows the minimum elevation angle to the highest
satellite over Washington, D.C., as a function of time. Every 24
hour period has ten elevation angle peaks of satellites on a
descending (from north proceeding towards the equator) at or near
the observer's zenith (90 deg). The lower, sharper peaks in the
figure represent other satellites on ascending passes; they are at
lower altitudes and thus going faster. These ascending satellites
are not actively transmitting to users on the ground at the times
when they are on ascending passes.
[0109] The preferred system uses a total of ten (10) satellites in
critically-inclined (i=63.4 deg) 6-hour orbits, phased and oriented
to provide optimal earth coverage. As will be seen, this geometry
also provides a very high elevation angle, and hence avoids
interference with the existing geo communications satellite band.
The preferred orbits have apogee and per-gee altitudes of 20074 and
654 kilometers, respectively.
[0110] From a user's viewpoint, the satellites are accessed
sequentially at nominal 2 hour and 24 minute intervals at exactly
the same point in the northwestern sky (the `start point` of the
tracking segment), and are tracked in a roughly northwest to
southeast trajectory to a point in the sky well short of
intersecting the geo band of satellites. The satellites remain at
apogee during the time while they are being tracked from the
ground. Hence, these satellites are only tracked, and communicated
with, while their velocity closely matches the velocity of the
earth. When the satellites begin to approach the perigee stage, and
hence their velocity increases relative to the earth's rotation to
differ therefrom by more than 25%, for example, they are no longer
being tracked by the communication equipment on the earth. At this
end point of the tracking segment, the ground communications
antenna is directed back to tracking its start point to repeat the
sequence as the next-appearing satellite is acquired. Tracking
along the active arc segment is accomplished at less than 2
deg/min. For the present array, this results in every ground
communications antenna effecting ten switchovers per day. As
explained above with reference to FIG. 1, the steering operation of
the disclosed system preferably uses phased array steering of the
antenna. However, more-conventional antenna steering is also
contemplated.
[0111] Importantly, the trajectory segments appear exactly the same
to the user for every satellite, since the azimuth-elevation trace
is repeated for each satellite. This system defines significant
advantages. Its operating altitudes are half that of existing geo
systems. This greatly reduces link margins and emitted power
requirements for the satellites.
[0112] Apogees are placed on the meridians of longitude of the
heavily-populated areas for which the constellation is optimized.
Apogee points may also be adjusted to approximate the targeted area
latitudes as well. The satellite tracking arcs over the targeted
areas remain roughly overhead (within 30-40.degree. of zenith),
with slow angular movement during periods when the satellite is
active. The trajectories for mid-latitude (20-50.degree. North
latitude) observers located directly under the apogee points in the
high-population targeted areas are approximately north-south
oriented.
[0113] All ten ground tracks are identical, and only the satellite
that is currently covering the repeating ground tracks charge. The
repeat cycle is 24 hours. Since the satellites move from one
geographic area to another, information once transmitted can be
re-broadcast at another location.
[0114] The Mercator plot of FIGS. 7A-7E show that the entire system
actually follows one ground track, repeating after 24 hours. It
actually `folds over` from the left edge of the world map to the
right edge, giving it the appearance of multiple traces.
[0115] Table II gives the orbital parameters, or ephemerides, of
the entire array of ten satellites:
[0116] TABLE II SYSTEM ORBITAL PARAMETERS Sat RMN MA # a(km) i(deg)
e,(ecc.) w,(deg) (deg) (deg) 1 16742 63.435 0.58 315 0 0 2 16742
63.435 0.58 315 072 072 3 16742 63.435 0.58 315 144 144 4 16742
63.435 0.58 315 216 216 5 16742 63.435 0.58 315 288 288 6 16742
63.435 0.58 315 180 0 7 16742 63.435 0.58 315 252 072 8 16742
63.435 0.58 315 324 144 9 16742 63.435 0.58 315 036 216 10 16742
63.435 0.58 315 108 288
[0117] Some adjustments will be required to account for long term
orbital perturbations as described above. This adjustment is common
in satellites requiring precise repeat cycles such as
Topex-Poseidon, or the Canadian Radarsat.
[0118] Similar views to those from the above can be drawn for the
preferred ten-satellite array. An important point of the
ten-satellite array, moreover, is that there is good
inter-satellite connectivity for cross-linking.
[0119] FIG. 7A shows the position of the satellites at time 00:00.
Compare this with FIG. 7B, which shows the same satellites
twenty-four minutes later. The satellite P4, which is substantially
over Washington, D.C., has moved very little, albeit P5 will be
picking up speed as it approaches perigee. P4 appears to hang over
Washington, D.C., since it is near the apogee portion of its orbit
and its velocity very closely matches the velocity of the
earth.
[0120] In contrast, during the same short period of time, the
satellite P1, at perigee, has moved very quickly and very far along
its orbit. Similarly, satellite P8 (over Europe), P5 (over Southern
Africa) and P9 have moved very little. Twenty-four minutes later,
FIG. 7C shows that satellite P4 has started to move away from the
United States, but satellite P7 is now in place, very close to its
apogee. This is evident from its position twenty-four minutes after
that, shown in FIG. 7D, where satellite P7 has moved only very
little, and is still well-covering the United States. At time 1:36
shown in FIG. 7E, the satellite P7 is over Washington, D.C.
[0121] The satellite P7 is still over Washington D.C. at time 2:00
hours, shown in FIG. 7F. The satellite starts to move at time 2:24,
shown in FIG. 7G.
[0122] The disclosed system intends that the satellites be used for
communication during only some part of the time while they are in
orbit. During other times in orbits, the satellites are not being
used for communication, but instead are charging their energy
storage. This feature of the invention has been described above,
but will be described in more detail herein with reference to FIGS.
2A, 4G and 4H.
[0123] FIG. 4G shows a view of the earth from, for example, the
view of the satellite from the sun. This figure shows all of the
satellite orbits, and their elliptical orbital paths. The
geosynchronous satellites are in equatorial planes shown as the geo
ring 800. Communications equipment on the earth communicates with
this geo ring 800. Moreover, sometimes the geo satellites are
perturbed by the earth's oblateness, hence effectively forming
orbits which are slightly inclined. The geo rings should therefore
be considered at occupying a 5' position bordering their nominal
position.
[0124] Ground communications equipment on the earth communicates
with this geo ring. The cone of communications to the geo ring is
shown as 802.
[0125] When the ground communication equipment on the earth
communicates with the satellites P1-P5, it should be seen that they
are aimed at a position of the sky, 804, which is completely
separated from the geo ring 802. According to the disclosed system,
a distance is maintained between the satellites and the geo ring
800. The angular separation e the minimum acceptable angular
separation which can ensure no interference between the geo ring
and the satellites of the disclosed system. An embodiment uses an
angular separation of 30.degree., which is an amount which will
obviate any possibility of interference problem. More generally,
however, any angular separation greater than 15.degree. would be
acceptable.
[0126] Taking the satellite P3 as an example, therefore, the
satellite can only be used according to the disclosed system when
it is in its orbit between the points labelled 808 and 810.
However, the virtual geo system which is preferably used according
to the disclosed system uses these satellites during even less of
their orbit, only between the points 812 and 814. When the
satellite is in the other positions of its orbit, the satellite is
not consuming power or transmitting. Therefore, this prevents any
possibility of interference with the geo satellite systems.
[0127] The operation of the equatorial satellites is similar. The
equatorial satellite array is shown in FIG. 4h. The equatorial
satellite is shown as satellite ring 850. If the ground station is
on the equator, shown as ground station 852, then it would, at
least at some times, interfere with satellites in the geo ring
shown as 854. However, if the ground station is separated from the
equator by at least 30.degree., such as shown as position 856, then
at least part of the satellite ring has no chance of interference
with the ring 854. Therefore, the satellite calculates geometries
such as to obviate interference with the satellite ring. Therefore,
more generally, the disclosed system operates as shown in FIG. 2a.
The antenna is controlled at step 350, and from the antenna control
the position of the satellite relative to geo are determined at
step 870. This can be determined, for example, from the pointing
angle of the antenna. Step 872 determines if there is any
possibility of interference between the two. This is determined
from a numerical difference between the pointing angle and the
position of the geo ring. If there is any possibility of
interference, control passes to step 874 where the satellite
communications is disabled. If interference is not possible at step
872, then the satellite is enabled at step 874. An enabled
satellite can be, but is not necessarily, turned on. For example,
in the virtual geo embodiments, the enabled satellite will be
maintained in the "off" position during some of the time when it is
enabled. Therefore, step 352 determines if the satellite is
powered. This may be determined from the repeating ground track, or
other information. If the satellite is not powered at step 352, the
battery is charged at step 356. If the satellite is powered, then
power is drawn from both the supply and the battery at step
354.
Second Embodiment
[0128] Another embodiment, also referred to herein as the "VIRGO"
embodiment, uses satellite sub-constellations with prograde
elliptical orbits of approximately 8 hour periods. Each of the
satellites within a sub-constellation has the same ground tracks as
the other satellites within the subconstellation, or repeating
ground tracks.
[0129] Each sub-constellation includes several satellites in each
of the individual ground tracks. The satellites are spaced such
that as one satellite leaves a service area, another satellite
replaces it in the same ground track.
[0130] As will be established herein, each satellite is in
communication with the ground station during a portion of the
trajectory where the satellite is at or near its apogee. During
this time, the relative motion of the satellite, i.e. the perceived
motion of the satellite relative to the Earth, is slow. The
satellite travels through a relatively small angular arc, e.g.,
40%, during its active phase.
[0131] As the one satellite departs from its active phase in the
descending direction, the ground user can switch to the
next-appearing satellite in the ascending portion of the active
phase of this next satellite. Continuity of coverage is thus
provided by this switch-over.
[0132] During its active phase, each satellite is virtually
geostationary. That means that it appears relatively stationary to
a user on the earth.
[0133] The concept behind the virtual geostationary orbit can be
illustrated with analogy to the walking juggler. A juggler's clubs
cluster together and move very slowly at the highest point in their
trajectories. At the low point of the trajectories, the juggler is
catching and transferring the clubs hand-to-hand rapidly. At the
high point of the trajectory, however, the clubs move much
slower.
[0134] The satellites in a virtually geostationary constellation
are intentionally placed in stable elliptical orbits with their
apogees over the intended users. Like the juggler, these portions
rise over the service area and appear to hang there. Additionally,
each satellite is active for only a predetermined portion of its
orbiting time, closest to its apogee portion. The satellites are
spaced such that when one satellite in the subconstellation reaches
its inactive portion, another satellite in the subconstellation
becomes active. Hence, the satellites are spaced such that one
ascending satellite replaces another descending satellite leaving
the service area.
[0135] Since the satellites are in 8 hour orbits, each satellite
peaks three times in each 24-hour day. Each of the peaks is located
to follow a populated region. Using a Northern Hemisphere apogee
orbit as an example, each satellite ascends, reaches its turn on
point and begins operating, goes through its peak ("apogee") and
then descends. The satellite eventually reaches its turn off point.
The satellite is then replaced, after its time of "hanging", by the
next satellite in the array. The first satellite then falls rapidly
into the Southern Hemisphere and quickly rises into the next
Northern Hemisphere peak. Each satellite's peak is placed over one
of the three Northern Hemisphere Continental masses each day. In
order to provide coverage to countries in the Southern Hemisphere,
the embodiment employs another grouping of 5 satellites having
their apogees in the Southern Hemisphere.
[0136] Each of the subconstellations is a mean motion 3 array. Each
of the satellite peaks is separated from other satellite peaks by
120.degree. of longitude (360.degree./3).
[0137] The longitudes selected for apogee placement of this array
are 79.degree. W, 41.degree. E, and 161.degree. E longitude. These
five satellites serve the populated areas of South America, South
Africa, Australia and New Zealand. Each satellite, in a single day,
appears at apogee three times. This requires three satellites out
of a total of five to be active at any time. Overall, each
satellite must then be active 3/5 of the time over a full day, or
14.4 hours. Since this represents one day's total active time, and
the satellite has been active over three geographic region, each
region will be covered by a single satellite for 4.8 hours. In
other words, each 8-hour satellite period, the satellite will be
active for a 4.8 hour period--or 2.4 hours on either side of the
apogee.
[0138] The satellites in this array have a duty cycle of 60%; that
is, they are actively communicating 60% of the time. Their on/off
switching times occur 2.4 hours on either side of the apogee. This
corresponds to a latitude of 46.degree., and an altitude of 18044
km. The active phase for each satellite occurs at latitudes greater
than 46.degree. and altitudes greater than 18044 (up to and
including apogee at 27288 km). The satellites remain well clear of
the GEO band, while active, so there is no possibility for
electronic interference with GEO communication satellites.
[0139] Because of the operating features discussed above, VIRGO
satellites operate only when the satellites are at least 40.degree.
separated from the line of sight of geo satellites. Hence, existing
Ku and C frequency equipment can be used without interfering with
other communication.
[0140] The elliptical planes in the two Northern Hemisphere
sub-constellations are inclined at 63.40 with respect to the plane
of the equator. This means that the apogees will always appear to
be roughly at 63.40 North latitude.
[0141] The two 5-satellite sub-constellations are called Aurora 1
and Aurora 2. These are used to provide continuous coverage of this
type. The third subconstellation is called Australis. Two or three
spare satellites are placed into "parking" orbits where they can be
boosted into different orbits if necessary.
[0142] The VIRGO.TM. orbital characteristics are as follows.
[0143] TABLE 1 VIRGO.TM. ORBITAL CHARACTERISTICS Aurora I.TM.
Aurora II.TM. Australis.TM. Spare Sats n=1-5 Sats n=1-5 Sats n=1-5
Satellites Semimajor 20281 20281 20281 7285 Axis Eccentricity 0.66
0.66 0.66 0.05346 Inclination 63.435 63.435 63.435 63.435 Right
341.5 255.3 52.2 0 Ascension of 53.5 327.3 124.5 the Ascending
125.5 39.3 196.5 180 Node 197.5 111.3 268.5 269.5 183.3 340.5 30
Argument of 270 270 90 270 Perigee 270 270 90 270 270 270 90 90 270
270 90 270 270 90 Mean Anomaly 0 108.2 0 0 144 252.2 144 0 288 36.2
288 0 72 182.2 72 216 324.2 216
[0144] This apogee of these VIRGO.TM. satellites is at 27,300
kilometers. This is approximately three-quarters the altitude of
geostationary satellites. This lower altitude provides less
propagation delay to orbit.
[0145] The ground tracks of this embodiment are shown in FIG. 9.
These produce the following locations of VIRGO.TM. active arcs.
[0146] TABLE 2 LOCATIONS OF THE VIRGO.TM. ACTIVE ARCS
(Sub-Satellite Longitudes in Degrees East) AURORA I.TM. AURORA
II.TM. NORTHERN NORTHERN AUSTRALIS.TM. HEMISPHERE HEMISPHERE
SOUTHERN HEMISPHERE 8-53 78-123 19-64 Europe India-China Africa
128-173 198-243 139-184 Japan Alaska-Hawaii Australia-NZ 248-293
318-3 259-304 Con.US N. Atlantic South America
[0147] Further information on the ground track is shown in the
following.
[0148] TABLE 3 VIRTUAL-GEO ORBITAL ELEMENTS, PREFERRED EMBODIMENT
All Satellites: Semi-Major Axis (a)=20381 km; Eccentricity,
e,=0.66; Inclination, I,=63.435.degree. Sat. Ground Track No. RAAN
omega. MA #1 (West. US) VG1a 350 270 0 #1 (West. US) VG2a 62 270
144 #1 (West. US) VG3a 134 270 288 #1 (West. US) VG4a 206 270 72 #1
(West. US) VG5a 278 270 216 #2 (East. US) VG1b 263.8 270 108.2 #2
(East. US) VG2b 335.8 270 252.2 #2 (East. US) VG3b 47.8 270 36.2 #2
(East. US) VG4b 119.8 270 180.2 #2 (East. US) VG5b 191.8 270 324.2
#3 (S.A., Australia) VG1 c 61 90 0 #3 (S.A., Australia) VG2c 133 90
144 #3 (S.A., Australia) VG3c 205 90 288 #3 (S.A., Australia) VG4c
277 90 72 #3 (S.A., Australia) VG5c 349 90 216
[0149] Although only a few embodiments have been described in
detail above, other embodiments are contemplated by the inventor
and are intended to be encompassed within the following claims. In
addition, other modifications are contemplated and are also
intended to be covered.
* * * * *