WO2015160416A2

WO2015160416A2 - Communication satellite system

Info

Publication number: WO2015160416A2
Application number: PCT/US2015/013941
Authority: WO
Inventors: Jan A. King
Original assignee: King Jan A
Priority date: 2014-01-31
Filing date: 2015-01-30
Publication date: 2015-10-22
Also published as: WO2015160416A3; WO2015160416A9

Abstract

A method of providing satellite coverage to the Northern and Southern Polar Regions includes launching multiple satellites into orbit with each satellite's orbit moving in the direction of one of the poles. The satellites are sequentially launched into orbit at differing values of the argument of perigee (ω). Each launch is spaced apart in time by a predetermined amount. Each satellite has an elliptical orbit and an orbital inclination that causes an orbital drift rate such that, over time, the apogee of the elliptical orbit moves in the direction of one of the poles at a predetermined rate.

Description

COMMUNICATION SATELLITE SYSTEM

CROSS-REFERENCE TO RELATED APPLICATION

This application claims the benefit of U.S. Provisional Application No. 61/934,078 filed on January 31, 2014 and entitled "ENHANCED COMMUNICATIONS SATELLITE SYSTEMS AND METHODS", the contents of which are incorporated herein by reference in its entirety.

TECHNICAL FIELD

This invention relates to an enhanced communication satellite system for the

Northern and Southern Polar Regions.

BACKGROUND

The Polar Regions of the Earth are becoming more important. Polar ice in both regions is receding earlier during each annual summer cycle. The Northern Polar Region is becoming available to shipping traffic during the summer months saving many thousands of miles per route, improving route and time efficiency, and overall productivity. Oil, gas and mineral exploration are increasing in the Northern Polar Region.

Current Geostationary Earth Orbit (GEO) satellites cannot be used for polar communications since satellites in Equatorial orbit cannot "see" the Polar Regions of the Earth. Traditional GEO satellite services are not available above 80° and are still poor or unavailable at 70° latitude. Even at latitudes as low as 60° appropriate GEO satellites may not be visible or available and frequently have no real high-speed broadband capability.

Known satellite systems have used highly elliptical orbits (HEOs) which have a particular relationship between the angle between the satellite's orbital plane and the Equator

(this parameter is called the inclination of the orbit). Further, the location of the apogee of the orbit (the highest point in the satellite's orbit) is placed so that the apogee occurs at the highest possible latitude (North or South depending on the Polar Region to be served). One such HEO orbit has the name "Molniya" and was developed by space scientists of the former Soviet Union. This orbit has been used to provide communications services at high latitudes for many years. By selecting an inclination parameter value of 63.43 degrees, this orbit does not rotate in its plane. At all other inclination angles that can be practically achieved, the orbit would rotate in its plane. (See Equation [1] below for a detailed explanation.)

One known system, developed by Dr. Karl Meinzer, was based on the realization that a spacecraft launched into a standard Geostationary Transfer Orbit (GTO) is very similar in shape to a Molniya orbit. Standard GTO differs in two ways - its inclination is not at the desired 63.43 degrees but, is rather about 27-29 degrees (for U.S. launched satellites), and its apogee lies on or very near the Equator, not over a high latitude location. Dr. Meinzer proposed modifying the orbital inclination after launch so that the inclination of the orbit is higher than the initial orbit inclination. The resulting orbit rotates more slowly than at its initial rate but, in the correct direction so that the apogee, over time, moves in the direction of the North Pole. By selecting the inclination correctly, the satellite, in a few years time, essentially becomes a Molniya orbit but with a slightly lower inclination. As the satellite orbit continues to drift, the satellite passes the ideal high latitude apogee position and the orbit coverage at high latitudes decreases. According to Dr. Meinzer, with a service life in orbit of five or so years, the satellite would spend much of its operational lifetime in a high Northern latitude position.

SUMMARY

Economic development in the Northern Polar Region and growth in science research in both Polar Regions requires improved communications services in order to be sustained. Safety of life and security in these regions depends upon good quality communications. However, no significant form of high-speed data service exists at very high latitude settlements on Earth.

According to one aspect, a method of providing satellite coverage to the Northern and Southern Polar Regions includes launching multiple satellites into orbit with each satellite's orbit moving in the direction of one of the poles.

Embodiments of this aspect may include one or more of the following features. The method includes sequentially launching multiple satellites into orbit. The multiple satellites are in orbit at differing values of argument of perigee (co). Each launch is spaced apart in time by a predetermined amount. Each satellite has an elliptical orbit and an orbital inclination that causes an orbital drift rate such that, over time, the apogee of the elliptical orbit moves in the direction of one of the poles at a predetermined rate.

The satellites are launched, for example, by ride sharing with a primary satellite going to GTO.

The method includes modifying a satellite's HEO orbit using the satellite's propulsion system. The satellite's orbit is modified to drift at a predetermined rate. For example, the predetermined rate is within the range of 4.0 to 145 degrees per year, the predetermined rate is 45 degrees +/- 3 degrees per 1.25 years, the predetermined rate is 90 degrees +/- 3degrees per 1.25 years.

The satellite's HEO orbit is modified by changing the orbital inclination. For example, the orbital inclination is changed to a final orbit within the range of 40 to 90 degrees. If the range of inclination of the final orbit is between 40 degrees and 63 degrees the orbit apogees of the satellites drift toward the North. If the inclination of the final orbit is between 64 and 90 degrees the orbit apogees of the satellites drift toward the South. The orbital inclination is changed, for example, to a final orbit of 55.60 degrees +0.46 to -0.52 degrees, to a final orbit of 48.53 degrees +0.87 to -0.97 degrees, to a final orbit of 73.48 degrees +0.92 to -0.76 degrees.

Each satellite drifts once around the Earth. An expended satellite is removed from orbit after the satellite has drifted once around the Earth and is replaced by a new satellite launched into approximately the same orbit as the expended satellite. A satellite is removed from orbit by applying a small velocity increment against the velocity vector of the orbit at its apogee.

Sequentially launching includes launching one satellite at a time. Each satellite drifts northward or southward. Alternatively, sequentially launching includes launching two satellites at a time. One of the two launched satellites drifts southward and one of the two launched satellites drifts northward. The satellites' positions in their orbits relative to their perigee position are equally spaced in angle (also called mean anomaly).

Eight satellites in orbit provide coverage at both Polar Regions, for example, of at least 100% of the time from each of two different satellites. Four satellites in orbit provide coverage at both Polar Regions, for example, of at least 100% of the time from one satellite and 30% of the time from each of two different satellites.

The time between satellite launches is in the range of .33 to 12 years. For example, the satellites are launched 1.25 years apart +/- 30 days or 0.75 years apart +/- 15 days.

The details of one or more embodiments of the invention are set forth in the accompanying drawings and the description below. Other features, objects, and advantages of the invention will be apparent from the description and drawings, and from the claims.

DESCRIPTION OF DRAWINGS

Fig. 1 is an illustration of an enhanced communication satellite system including a constellation of satellites.

Figs. 2-4 are graphs of the coverage provided by a single satellite in orbit.

Figs. 5A-5H illustrate orbit drift over time of a single satellite.

Figs. 6-12 are graphs of the coverage provided by a two, three, four, five, six, seven and eight satellites as satellites are added to a constellation over time.

Fig. 13 is a graph corresponding to Figs. 6-12 showing the coverage of the Northern Polar Region.

Fig. 14 is a graph corresponding to Figs. 6-12 showing the coverage of the Southern Polar Region.

Figs. 15A-15I illustrate orbit drift over time as each satellite is added to the constellation.

Figs. 16A and 16B are graphs of the coverage provided by an established three satellite constellation.

Figs. 17A and 17B are graphs of the coverage provided by an established four satellite constellation. Figs. 18A and 18B are graphs of the coverage provided by an established five satellite constellation.

Figs. 19A and 19B are graphs of the coverage provided by an established four satellite constellation having a modified orbit inclination.

Figs. 20A and 20B are graphs of the coverage provided by an established five satellite constellation having the modified orbit inclination.

Figs. 21 A and 21B are graphs of the coverage provided by an established eight satellite constellation having the modified orbit inclination.

Figs. 22 and 23 are graphs of the drift time for the rotation of the orbit by 45 degrees versus the orbit inclination.

Figs. 24 and 25 illustrate two representative satellite configurations.

DETAILED DESCRIPTION

The main drawback of the Meinzer orbital concept is that the apogee position (the argument of perigree (ω)) within the orbit plane is non-constant. The orbit rotates at a specified rate set by three orbital parameters (see U.S. Patent No. 5,199,672, herein incorporated by reference). The relevant equation governing the time rate-of-change of the argument of perigee of the orbit, (dco/dt) is:

Where:

a = orbit semi-major axis (km)

e = orbit eccentricity (dimensionless number from 0 to 1)

i = orbit inclination in degrees

Re = Radius of the Earth (6378.137 km)

dco/dt = first derivative of the argument of perigee (degrees/day) If equation (1) is set equal to zero and all of the other above parameters have values within their normal range, then in order for the inclination of the orbit to drive the first derivative of the argument of perigree (co) to zero the following must also be true:

and

i = 63.434949° or 116.575051°

We note in equation (2) that if the inclination of the orbit is less than 63.43 degrees, then the first derivative (dco/dt) is positive and the orbit drifts to the North. If the inclination of the orbit is more than 63.43 degrees then dco/dt is negative and the orbit drifts to the South.

A satellite communications ground station at the North Pole initially sees a satellite on a Meinzer orbit only 30% of each day, but, as the apogee of the orbit slowly moves to the North within its orbit plane, the coverage improves dramatically. When the satellite's apogee is at the highest latitude position, the North Pole ground station sees the satellite 85% of each day. But, after this position is reached (after a few years of time have passed) the coverage becomes poorer again. Eventually, the satellite orbit lies with its apogee back on the Equator (this time on the other side of the Earth) and the daily temporal coverage of the North Pole is back to only 30%. As the orbit drifts even further, the apogee drifts into the Southern Hemisphere and the coverage to the North becomes even worse. However, the coverage to the South Pole now begins to improve. In effect, the advantage that was held by satellite ground stations in the North now passes to satellite ground stations in the South. Hence, over long time periods (for example, 10 years) the satellite can provide service for both Polar Regions.

By establishing a network of man-made satellites in specific Earth orbits, collectively known as a constellation, communications services can be provided over a broad area and for extended periods of time to both Polar Regions of the Earth using the same system of satellites. The Meinzer orbit is made into a constellation system of orbits so that, once the constellation is built, both Polar Regions have 100% temporal coverage. One example of how this constellation can be built is:

a. Launching a first satellite into a GTO orbit (where the orbits apogee lies in the equatorial plane) and increasing the satellite's inclination parameter to a value of 55.6 degrees (for example, as other variants exist). This sets the drift rate of the orbit in its plane at 0.1000 degrees/day or approximately 45 degrees per 15 months.

b. After 15 months (for example), a second satellite is launched into a GTO orbit and the process in a. is repeated.

c. After another 15 months (for example) a third satellite is launched into a GTO orbit and the process is again repeated.

d. This process is repeated until 8 satellites (for example) are launched over a period of 8.75 years (for example), creating a full constellation.

e. More satellites are added on 15 month launch cycles (for example). As the 9th satellite is launched it replaces the first satellite which is 10 years old, has expended its useful life, and its major axis (or line of apsides) has drifted one time around the Earth. This old satellite can be removed from orbit by a very small delta-V maneuver thus removing all debris from space associated with this oldest satellite.

f. The process continues with one new satellite added every 15 months (for example) and one old one removed from orbit.

Referring to Fig. 1, a communication satellite system 10 that provides satellite coverage to the Northern Polar Region 12 and the Southern Polar Region 14 includes multiple satellites, for example, eight satellites 16a-16h, that are sequentially launched into orbit to form a satellite constellation. Each satellite launch is spaced apart in time by a predetermined amount, for example, about 1.25 years (15 months). Each satellite has an elliptical orbit with a high apogee and a low perigee, and an orbital inclination that causes an orbital drift rate such that over time the apogee of the elliptical orbit moves in the direction of one of the poles at a predetermined rate. At any given time, the satellites in the constellation each have a different value for the argument of perigee (ω), with all of the satellites equally spaced in mean anomaly. The communication satellite system 10 also includes ground antennae 18 at the Northern and Southern Polar Regions for communicating with the satellites. Users 20 in the vicinity of a ground antennae 18 can access global communication systems.

Initially, upon injection of a spacecraft, for example, the satellite 16a, into GTO, the orbits apogee lies almost exactly in the equatorial plane. As illustrated in Fig. 2, the starter orbit (at ω = 180° and with the orbit's apogee on the equator) sees both poles, each 30% of the time each day (the percentage temporal coverage), though not at the same time for both poles.

The location of the orbit's apogee in the equatorial plane is optimum for changing the orbit inclination with minimum fuel so the inclination is preferably changed right away. The launcher leaves a typical satellite headed for GEO in a GTO which has a rate of motion dco/dt = +0.6069 degrees/day. This is for a US launch from Cape Canaveral to a 28 degree inclined GTO. For this reason the launcher biases the argument of perigee (ω) to start at about 178 degrees. Hence, the orbit argument of perigee drifts at the rate of +0.6069 degrees/day and after about 3.5 days (3.5 x .607 = 2.1 degrees) the orbit argument of perigee equals 180 degrees - the optimum location for motor firing in order to change inclination. This is also approximately equal to eight orbits. The motor is preferably fired on the 7th, 8th or 9th apogee when the value of the argument of perigee is closest to 180 degrees.

After the inclination increasing burn, the orbit apogee begins to drift within its orbit plane in a prescribed way, either toward the North or the South depending on the initial launch angle. The satellite orbit slowly begins to change as the apogee moves in the

Northern or Southern direction. The orbital inclination is selected to provide a predetermined orbital drift rate. For example, an orbital inclination of 55.60° results in a rotation of the elliptical orbit that sets the drift rate of the orbit within its plane to a rate of 45 degrees per 1.25 years.

Inclination = 55.60°

dco/dt = 0.1000 day ~ 45 15 months

a = 24,771.1 Ion

e = 0.702220 coi = 178° to 180°

Ω = 0.0°

M = 180°

With the satellite's inclination parameter increased to set the desired drift rate of the orbit within its plane at approximately 45 degrees per 1.25 years, after 1.25 years in service, at ω = 225° the satellite provides the percentage temporal coverage at different ground station latitudes illustrated in Fig. 3. At this time, the argument of perigee (co) has advanced by 45 degrees (or l/8th of a rotation in its orbit plane), and the coverage of the Southern Polar Region has gone down to less than 5% of the time while the coverage of the Northern Polar Region has gone up to more than 70% of the time or about 17 hours of each day.

As illustrated in Fig. 4, by the time the satellite is 2.5 years old, the argument of perigee (co) has advanced by 90 degrees (or 1/4 of a rotation in its orbit plane). By this time the coverage of the Southern Polar Region has gone down to 0% of the time while the coverage of the Northern Polar Region has gone up to more than 80% of the time.

Figs. 5A-5H illustrate a "Daisy Wheel" showing the change in the orbit of the single satellite over a period of 8.75 years. The greatest satellite coverage is obtained in the region of the Earth facing the orbit apogee. Thus at time 2.5 years (Fig. 5C), the satellite provides the greatest coverage of the Northern Polar Region, and at time 7.5 years (Fig. 5G), the satellite provides the greatest coverage of the Southern Polar Region.

By launching multiple satellites on the same orbit and inclination over a period of time, over time 100% coverage can be obtained at both the Northern and Solar Polar

Regions.

Referring to Fig. 6, at 1.25 years after the launch of the first satellite, a second satellite is launched with its argument of perigee (co) at 180°, which is 45° degrees behind the current argument of perigee of the first satellite. The perigee altitude and the inclination of the second satellite are changed in the same way as its predecessor. Now the coverage of the Southern Polar Region is up to 32% of time (between the coverage of the two satellites) and the Northern Polar Region is covered 78% of the time by one satellite and 22% of the time by two satellites. At the same time, the Southern Polar Region is covered approximately 33% of the time by one satellite, which is about 8 hours of coverage for the Southern Polar Region (twice as much as currently available using current means). For the North Polar Region this amounts to more than 18.5 hours single coverage and 5.3 hours double coverage. Over the next year the coverage in the Northern Polar Region increases to 100% as the Southern Polar Region coverage slowly decreases again.

Referring to Fig. 7, after another 1.25 years, that is, 2.5 years after the original launch, the first satellite has advanced in its orbit one quarter of a turn in its plane and its apogee is at the highest latitude (closest to the pole) it will ever be, and the second satellite has advanced l/8th of a turn. At this time a third satellite is launched in the same way as its predecessors and the system now provides 100% coverage of the Northern Polar Region with at least 65% of the time the Northern Polar Region seeing two satellites and about 8% of the time the Northern Polar Region seeing all three satellites. From this time onward in the life of the system the Northern Polar Region always have 100% coverage.

Fig. 8 shows the coverage at year 3.75 when a fourth satellite has been launched. Coverage of the Southern Polar Region (about 25 to 35%) while being twice as good as any satellite service seen by SPS today, is still poor in comparison to the Northern Polar Region. Referring to Fig. 9, at year 5, with the launch of the 5th satellite, the apogee of the first satellite moves into the Southern Hemisphere. Now the coverage for the Southern Polar Region picks up dramatically. At year 6.25 (Fig. 10), with the launch of the 6^th satellite, the Southern Polar Region coverage is over 80% and double coverage of the Southern Polar Region is more than 30%. The coverage at year 7.5, after launch of the 7^th satellite, is shown in Fig. 1 1 , and the coverage at year 8.75, after launch of the 8^th satellite, is shown in Fig. 12. Both Polar Regions now have 100% coverage.

1.25 years later, when it is time for the 9th satellite launch, the oldest satellite is ten years old. The first satellite, which has been once around the orbit plane, can be retired. With a satellite rocket motor burn of only 70 m/sec against the orbit velocity, the first satellite can be dropped into the Pacific Ocean (as an example) on one of its perigees and replaced by a 9th satellite. Figs. 13 and 14 illustrate the percent temporal coverage of the Northern Polar Region and the Southern Polar Region, respectively, over time as the satellites are added to the constellation.

Figs. 15A-15I illustrate a "Daisy Wheel" showing the change in the orbit of each satellite over a period of 8.75 years with the addition of a satellite to the constellation each 1.25 years.

The number of satellites in the constellation can be greater than or less than eight. For example, three, four or five satellites can be in the constellation. An example of the coverage of a three satellite constellation having the satellites' arguments of perigee (ω) spaced 120° apart, i.e., having the satellite launch dates spaced 3 1/3 years apart, for an inclination value of 55.60° at various values of the arguments of perigee (shown at 0.833 years apart in the time of observation), is illustrated in Fig. 16A (co = 0°, 120°, 240°) and Fig. 16B (ω = 30°, 150°, 270°). An example of the coverage of a four satellite constellation having the satellites' arguments of perigee (co) spaced 90° apart, i.e., having the satellite launch dates spaced 2.5 years apart, for an inclination value of 55.60° at various values of the arguments of perigee (1.25 years apart in the time of observation), is illustrated in Fig. 17A (co = 0°, 90°, 180°, 270°) and Fig. 17B (co = 45°, 135°, 225°, 315°). An example of the coverage of a five satellite constellation having the satellites' arguments of perigee (co) spaced 72° apart, i.e., having the satellite launch dates spaced 2.0 years apart, for an inclination value of 55.60° at various values of the arguments of perigee (1.0 years apart in the time of observation), is illustrated in Fig. 18A (co = 0°, 72°, 144°, 216°, 288°) and Fig. 18B (co = 36°, 108°, 180°, 252°, 324°).

The orbital inclination, and thus the drift rate, can also be selected as desired. For example, an orbital inclination of 48.53° results in a rotation of the elliptical orbit that sets the drift rate of the orbit within its plane to a rate of approximately 90° per 1.25 years.

Inclination = 48.53°

dco/dt = 0.200°/day « 90 15monfhs

a = 24,771.1 km

e = 0.702220 coi = 178° - 180°

Ω = 0.0°

M = 180°

An example of the coverage of a four satellite constellation for an inclination value of 48.53°, with satellites launched every 1.25 years, is illustrated in Figs. 18 and 19. An example of the coverage of an eight satellite constellation for an inclination value of 48.53°, with satellites launched every 7½ months, is illustrated in Figs. 20 and 21.

The table below gives the drift time for the rotation of the orbit by 45 degrees (as would be associated with an eight satellite constellation) versus the final inclination of each satellite orbit. If one wishes to adjust the spacing between any two launches, the inclination is adjusted in accordance with this table or the corresponding graphs illustrated in Figs. 22 (North drifting) and 23 (South drifting). The inclination of the modified orbit is preferably in the range of approximately 40 to 90 degrees. If the constellation is to drift to the North, the range of inclinations for the final orbits would preferably be from 40 to 62 degrees. If the constellation is to drift to the South, the range of inclinations for the final orbits would preferably be from 65 to 90 degrees.

Relationship Between the Modified GTO Orbit inclination and Drift Time

for a 45°Change in the Orbit Argument of Perigee

Orbit Modified Inclinotlon: d /dt: Drift Time for&u of 45.0' Drift Direction:

35.00 deg, 0.3952 deg./day 113.9 days 0.31 years North

40.00 deg. 0.3246 deg./day 138.6 days 0.38 years ^' North

45.00 deg. 0.2517 deg./day 178.8 days 0.49 years North

48.53 deg. 0.2002 deg./day 224.8 days 0.62 years North

50.00 deg. 0.1789 deg./day 251.5 days 0.69 years North

55.00 deg. 0.1082 deg./day 415.9 days 1.14 years North

55.60 deg. 0.1000 deg./day 450.0 days 1.23 years North

57.50 deg. 0.0744 deg./day 604.8 days 1.66 years North

60.00 deg. 0.0420 deg./day 1071.4 days 2,93 years North

62.50 deg. 0.0111 deg./day 4054.1 days 11.10 years North

63.43 deg. 0.0000 deg./day ω value frozen at this inclination No Drift

65.00 deg. -0.0180 deg./day 2500.0 days 6.845 years South

67.50 deg. -0.0449 deg./day 1002.2 days 2.744 years South

70,00 deg. -0.0697 deg./day 645.6 days 1.768 years South

73.48 deg. -0.1000 deg./day 450,0 days 1.232 years South

75.00 deg. -0.1116 deg./day 403,2 days 1.104 years South

77.50 deg. -0.1285 deg./day 350.2 days 0.959 years South

80.00 deg. -0.1425 deg./day 315.8 days 0.865 years South

82.50 deg. -0.1535 deg./day 293.2 days 0.803 years South

85,00 deg. -0.1614 deg./day 278.8 days 0.763 years South

90.00 deg. -0.1678 deg./day 268.2 days 0.734 years South

In a four satellites constellation (instead of eight) the drift time allows for the rotation of the orbit argument of perigee by 90 degrees, which corresponds to a wait time between launches (corresponding to the drift time for a Δω of 90 degrees), which is twice as long as those given in the table above for a drift time for Δω of 45 degrees.

To maximize the coverage of the constellation and to assure that the maximum number of satellites are visible in the Polar Regions over all time, the mean anomalies of the satellites in their orbits (i.e., the position of a satellites in their orbits relative to the perigee position) should be equally spaced. For example, if satellite 1 of 8 has a mean anomaly position Ml at time T of 0° then satellite 2 should have a position M2 at time T of 45°; satellite 3 should have a position M3 at time T of 90° and so forth with satellite 8 having a position M8 at time T of 315°.

In an eight satellite constellation, if all eight satellites were to lie in a plane then that plane would be inclined to the Earth's orbit by either, for example, 55.6° or 48.5° depending on how fast it is desired that the Daisy Wheel rotate. In the situation with all orbits in the same plane, each satellite in its orbit is 45 degrees advanced around its own particular orbit from the next. For example, referring to Fig. 15H, satellite 4 in orbit 4 is shown at perigee (low point of orbit); satellite 3 in orbit 3 is at 45 degrees ahead of perigee (measured from the center of the earth); satellite 2 in orbit 2 is at 90 degrees ahead of perigee, and so forth. Each of the satellites is at a different altitude and at a different location over the Earth.

However, the satellite orbits need not lie in one plane. This is because as the satellites near the Polar Region the satellites can all "see" or view the Polar Regions equally well to first order regardless of the direction from which the satellite approaches the pole.

Traditional satellite launch vehicles and traditional propulsion systems onboard each satellite can be used to establish the constellation. A benefit of the constellation system is that ride sharing is permitted with a primary satellite going to GTO and then the secondary satellite's propulsion system is used to modify the HEO orbit so that the satellite drifts in its plane in the desired direction at the desired rate. This consideration is important because most commercial launches inject satellites into GTO orbit, making more opportunities available to launch secondary satellites 16a-16h.

The satellites are designed to launch on any geostationary launch opportunity with at least 220 Kg of payload margin. Most significantly, the performance of the modified GTO orbit used by the constellation system does not depend on the launch time of day of the primary satellite. That is, the performance of the final orbit is virtually independent of the right ascension (RAAN or Ω) of the primary mission. This occurs because the performance of a HEO system at high latitudes (where longitude lines converge) is nearly constant, independent of the ascending node of each spacecraft in the constellation (i.e., as discussed above, the view of the polar regions does not depend on the direction from which the satellite approaches the pole). While the Mean Anomaly spacing (also called phasing) of the orbit is more critical and requires nearly equal spacing in the time between the ascending nodes of satellites in adjacent orbits, the Ω of the orbits in the constellation are unimportant.

Therefore, follow-up launches of constellation satellites need not wait to team up with a primary satellite requesting a particular launch slot. Also, the launches are not seasonally dependent. Only the time between consecutive launches to GTO is of major importance. Example Orbits:

GEO satellites are only occasionally launched into circular geosynchronous orbits with zero inclination - and "parked" on the GEO arc - straight away. For a large majority of the launches the satellites themselves start with an initial GTO (geostationary transfer orbit) and with their own rocket motor create a GEO. GTOs vary somewhat but, most of these orbits have a low perigee (in the range from 200 to 700 km), an apogee at about GEO altitude, which is 35,786 km, and an inclination which depends a lot on the launcher and, varies from 51° for Russian launchers, to 24° to 29° for launches from the U.S., to 5° to 7° for the European Ariane 5 rocket launched near the equator in French Guiana. What GEO satellites do with their on-board rocket motors is, 1) Circularize the GTO and make the low perigee into an orbit with the same perigee altitude as the apogee altitude (35,786 km) and 2) they use the same rocket motor burn to, "take out" all of the initial inclination and adjust the final orbit inclination equal to zero degrees.

The GTO orbit is a good starting point for launching the satellites in the GTO-MOD constellation. To form the constellation, a satellite's orbit is modified by increasing the inclination of the orbit without raising the apogee. The perigee is raised slightly (to approximately 1000 km) to improve the stability of the orbit and to get the orbit out of the atmospheric drag region. This doesn't take very much change in orbit velocity.

The AV requirements and characteristics of the resulting orbits (as well as mass available) for various commercial launchers is shown in the table below:

Final Orbit nominal alternate

a 24,771 km

e 0.7021 km

hs 35,786 km

hp 1,000 km

i 55.60 deg. 48.53 deg.

ω 180 deg. 0 deg.

Ω Don't Care

If it is desired to have a new orbit drift southward rather than northward, the inclination of the orbit can be increased to a value greater than 63.43°, allowing the construction of a satellite constellation that favors the Southern Polar Region first. For example, by increasing the inclination of the satellites to a value of 73.50°, the new orbit will rotate southward at a rate of -0.1000 degrees/day or approximately 45° per 15 months.

The direction of the rotation of the orbit (i.e., whether the new orbit begins to drift to the North or the South) can also be adjusted by selecting the initial GTO orbit at launch to have an initial argument of perigee value (co) near 0.00 or near 180°. Launches having an initial ω near the value of 0.00° will begin to drift to the South, while launches having an initial co near 180° will drift to the North. A constellation of satellites can be established in a reduced period of time and provide improved initial coverage (satellite visibility) in both Polar Regions by simultaneously launching one satellite having an initial co value of 0.00° and another satellite having an initial co value of 180°. By periodically and simultaneously launching two satellites that drift in opposite directions, the constellation can then be completed in half the time required for individual launches. Once the requisite number of satellites have been launched in each rotation direction, a double constellation will have been created, one advancing clockwise, the other advancing counter-clockwise. This approach increases the cost but also increases the coverage (satellite visibility) during the formation time. The overall constellation communications capacity is also doubled.

Due to the spaced launching of the satellites in the constellation, capacity is added to the system as the demand increases in the Polar Regions. Furthermore, the time between launches allows any problems discovered in prior satellites to be fixed for the next satellite before the next launch. Improvements and enhancements can also be added during the launch interval. For example, one enhancement can be to increase the capacity of each follow-on satellite if the communications capacity demand increases faster than anticipated by building a bigger satellite. Alternatively, more satellites can be launched if more capacity is needed.

Each satellite is, for example, about the size of a bar refrigerator and weighs about 220 Kg when maximum fuel capacity is used, however, under typical launch conditions the mass is less than 150 Kg. The spacecraft EOL power is approximately 500 watts and the spacecraft total volume is about 0.5 m³.

System Parameters of an Example Spacecraft:

Redundancy: Full/4 Wheels/6 TWTAs User Beam G/T / Xpdr: -3.5 dB/K

An exemplary spacecraft is illustrated in Figs. 22 and 23. Figure 22 shows a spacecraft with a capacity of 2 transponders and approximately 110 Mbps (in each direction), while Figure 23 shows a spacecraft with a capacity of 4 transponders and approximately 220 Mbps (in each direction).

The satellites are oriented in their elliptical orbits with the main body of the satellite directed toward the center of the Earth (called NADIR) and directive Ka-band antennas (example of a useful frequency band for polar communications) on the satellites point directly at either a gateway location or toward the Polar Service Area. When the desired service area is the entire polar region the center of the Service Area beams are directed toward the North Pole or the South Pole at all times when that pole is visible to the satellite. In order to track the sun the spacecraft must also rotate about its NADIR axis (known as Yaw steering) and rotate the two solar array wings toward the sun to provide the 500+ watts of power needed for its mission.

The non-geostationary satellites of the constellation require tracking antennas onboard the spacecraft and at the ground stations. By having uplink and downlink frequencies at Ka-band, the required size of the ground station antenna as well as the satellite antenna is reduced when compared to using lower frequencies like Ku-band (11 to 15 GHz). This allows the provision of such services in the Polar Regions using only small satellite systems.

Each of the Service Area antenna beams on-board the satellites, covering the

Northern or Southern Polar Region service are 8.0 to 9.0° (half power beamwidth). This beamwidth covers the service area (63.4° latitude to the pole) when the satellite is at apogee. The antenna system on the satellite is range compensating: the beam roll-off experienced at the pole being served, is larger when the satellite is closer to the Earth (note that the beam size of the pattern is constant but, as the Earth gets bigger as seen by the satellite the projection on the Earth gets smaller). This is off-set by the reduced range (and, in particular, the reduced path loss) between the user and the satellite. This range compensation works well over about a factor of 2.5 in altitude change, from apogee down to about 15,000 km. While some losses in link performance occur at altitudes below this value, the link adapts to these conditions well. In any case, another satellite at a higher elevation can then be selected by each user as it becomes available. Hence, a single spacecraft antenna system may be used for all orbit altitude circumstances. No beam switching is thus required.

A number of embodiments of the invention have been described. Nevertheless, it will be understood that various modifications may be made without departing from the spirit and scope of the invention. Accordingly, other embodiments are within the scope of the following claims.

Claims

WHAT IS CLAIMED IS:

1. A method of providing satellite coverage to the Northern and Southern Polar Regions, comprising:

sequentially launching multiple satellites into orbit such that the multiple satellites are in orbit at differing values of argument of perigee (co), each launch being spaced apart in time by a predetermined amount;

each satellite having an elliptical orbit and an orbital inclination that causes an orbital drift rate such that, over time, the apogee of the elliptical orbit moves in the direction of one of the poles at a predetermined rate.

2. The method of claim 1 wherein launching comprising ride sharing with a primary

satellite going to GTO.

3. The method of claim 1 further comprising modifying a satellite's HEO orbit using the satellite's propulsion system.

4. The method of claim 3 wherein the satellite's orbit is modified to drift at a predetermined rate.

5. The method of claim 4 wherein the predetermined rate is within the range of 4.0 to 145 degrees per year.

6. The method of claim 5 wherein the predetermined rate comprises 45 degrees +/- 3

degrees per 1.25 years.

7. The method of claim 5 wherein the predetermined rate comprises 90 degrees +/- 3 degrees per 1.25 years.

8. The method of claim 3 wherein modifying the satellite's HEO orbit comprises changing the orbital inclination.

9. The method of claim 8 wherein the orbital inclination is changed to a final orbit within the range of 40 to 90 degrees.

10. The method of claim 8 wherein the orbital inclination is changed to a final orbit of 55.60 degrees +0.46 to -0.52 degrees.

1 1. The method of claim 8 wherein the orbital inclination is changed to a final orbit of 48.53 degrees +0.87 to -0.97 degrees.

12. The method of claim 8 wherein the orbital inclination is changed to a final orbit of 73.48 degrees +0.92 to -0.76 degrees.

13. The method of claim 1 wherein each satellite drifts once around the Earth.

14. The method of claim 1 further comprising removing an expended satellite from orbit after the satellite has drifted once around the Earth and is replaced by a new satellite launched into approximately the same orbit as the expended satellite.

15. The method of claim 1 further comprising removing a satellite from orbit by applying a small velocity increment against the velocity vector of the orbit at its apogee.

16. The method of claim 1 wherein sequentially launching comprises launching one satellite at a time.

17. The method of claim 16 wherein each satellite drifts northward.

18. The method of claim 16 wherein each satellite drifts southward.

19. The method of claim 1 wherein sequentially launching comprises launching two satellites at a time.

20. The method of claim 19 wherein one of the two launched satellites drifts southward and one of the two launched satellites drifts northward.

21. The method of claim 1 wherein the satellites' positions in their orbits relative to their perigee position are equally spaced.

22. The method of claim 1 wherein eight satellites in orbit provide coverage at both Polar Regions of at least 100% of the time from each of two different satellites.

23. The method of claim 1 wherein four satellites in orbit provide coverage at both Polar Regions of at least 100% of the time from one satellite and 30% of the time from each of two different satellites.

24. The method of claim 1 wherein the time between satellite launches is in the range of .33 to 12 years.

25. The method of claim 24 wherein the satellites are launched 1.25 years apart +/- 30 days.

26. The method of claim 24 wherein the satellites are launched 0.75 years apart +/- 15 days.