EP1512233A2 - Orbite fortement elliptique pour satellites de communication - Google Patents

Orbite fortement elliptique pour satellites de communication

Info

Publication number
EP1512233A2
EP1512233A2 EP02792458A EP02792458A EP1512233A2 EP 1512233 A2 EP1512233 A2 EP 1512233A2 EP 02792458 A EP02792458 A EP 02792458A EP 02792458 A EP02792458 A EP 02792458A EP 1512233 A2 EP1512233 A2 EP 1512233A2
Authority
EP
European Patent Office
Prior art keywords
satellite
orbit
constellation
inclination
satellites
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP02792458A
Other languages
German (de)
English (en)
Inventor
Alan Global Radio S.A. KIDD
Paul Heinerscheid
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Global Radio SA
Original Assignee
Global Radio SA
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Global Radio SA filed Critical Global Radio SA
Publication of EP1512233A2 publication Critical patent/EP1512233A2/fr
Withdrawn legal-status Critical Current

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
    • B64G1/1085—Swarms and constellations
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
    • B64G1/1007—Communications satellites
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/242—Orbits and trajectories
    • H—ELECTRICITY
    • H04—ELECTRIC COMMUNICATION TECHNIQUE
    • H04B—TRANSMISSION
    • H04B7/00—Radio transmission systems, i.e. using radiation field
    • H04B7/14—Relay systems
    • H04B7/15—Active relay systems
    • H04B7/185—Space-based or airborne stations; Stations for satellite systems
    • H04B7/195—Non-synchronous stations

Definitions

  • the invention is generally related to satellites. More particularly, the invention is related to communications satellite orbits.
  • Communications satellites are used as relay stations in space.
  • One use of a communications satellite includes the broadcast of media content, such as radio or television programming.
  • media content such as radio or television programming.
  • DAB digital audio broadcasting
  • One approach to transmission of radio programming is digital audio broadcasting ("DAB"), which attempts to provide radio programming free from interference or distortion caused by mountains, high-rise buildings, weather conditions, etc.
  • DAB may also transmit text, data, images and video.
  • Satellite communications may be affected by a satellite's orbit.
  • a satellite's orbit may determine quality of signal for a DAB provider.
  • GSO geosynchronous or geostationary orbit
  • HEO highly elliptical orbit
  • the satellite may be a communications satellite, such as a digital audio broadcasting ("DAB") satellite.
  • DAB digital audio broadcasting
  • the orbit may be a lower inclination variation of a Tundra orbit having a ground track with a teardrop shape. In one embodiment, the inclination of the orbit may be less than 55 degrees.
  • the satellite following the lower inclination HEO orbit may be a part of a satellite constellation.
  • the satellite constellation may include a three satellite or a six satellite constellation.
  • the lower inclination HEO orbit allows sufficiently high elevation for high quality of service in Europe.
  • Fig. 1 is a map illustrating one embodiment of an optimized satellite orbit
  • Fig. 2 is a map illustrating an enlarged section of the map of Fig. 1 ;
  • Figs. 3a and 3b illustrate the performance of a three satellite constellation following the orbit of Fig. 1 ;
  • Figs. 4a and 4b illustrate the performance of a six satellite constellation following the orbit of Fig. 1.
  • the optimized orbit of the present invention may include a lower inclination HEO orbit.
  • the lower inclination HEO orbit may be a lower-inclination variation of a Tundra orbit having a teardrop shape.
  • the communications satellite system following the lower inclination HEO orbit may include a three-satellite or a six-satellite constellation.
  • the three- satellite constellation may be launched first, and the more satellites may be added later to the three-satellite system to create the six-satellite system.
  • Fig. 1 is a map illustrating one embodiment of the highly elliptical orbit of the present invention.
  • the orbit of a communications satellite may have a ground track 100 of a teardrop shape.
  • the orbit may be a variation of the Tundra orbit having a lower inclination. In one embodiment, the inclination may be 55 degrees or less.
  • the orbit may be used to broadcast to a European coverage area 110.
  • the standard Tundra orbit is a 24-hour elliptical orbit, first investigated by the Russians and having an inclination of approximately 63.4 degrees.
  • the standard Tundra orbit ground track includes a small upper loop which a satellite may spend about eight hours traversing, making the satellite appear quasi stationary to an observer in the area covered by the upper loop. For example, the upper loop may be above Europe, thus making the satellite following the standard Tundra orbit appear quasi stationary to an observer in Europe.
  • tundra orbits experience less severe exposure of satellites to radiation, and fewer eclipses.
  • a satellite in a GSO may experience over 100 eclipses, the longest lasting about 72 minutes.
  • a satellite may experience 38 eclipses, the longest lasting about 75 minutes.
  • no eclipse will occur while the satellite is providing service to the European coverage area 110. This allows for a spacecraft design to conserve mass by reducing the number of batteries that are carried.
  • the higher orbit of the standard Tundra orbit requires more energy to launch than other orbits, and the 63.4 degree position is not ideal for European coverage.
  • the Tundra orbit has a much higher perigee than comparable HEOs.
  • the orbital constraints of the Tundra are not as great as the comparable HEOs, such as the Molniya, another Russian orbit.
  • the perigee of the Tundra orbit is considerably higher than that of the Molniya orbit, the precession of the orbit is less influenced by perturbations from the Earth.
  • the coverage in Europe may be optimized. The reduction in inclination greatly improves elevation angles in the majority of Europe.
  • Table 1 gives parameters for one embodiment of the lower inclination Tundra orbit of
  • a Telemetry, Tracking and Command (“TT& C") station may need to see the satellite during its perigee pass because some station keeping maneuvers may be performed only at perigee, the point of the orbit at which the satellite is closest to the Earth.
  • a station located near the equator may see the satellite during its entire orbit. For example, a station may be located at Libreville in Gabon, which is located near the Equator and at the apogee in longitude.
  • a TT&C station would have 24 hour coverage with a minimum elevation angle to the satellite of about 25 degrees.
  • a station located in southern Africa may have a good visibility of the satellite at perigee, but would not see the satellite at apogee, the furthest point from the Earth in the orbit.
  • a TT & C station may be located in Europe for a satellite providing service to the European coverage area 110.
  • a second TT & C station may be needed to provide redundancy as well as command capabilities when the satellite passes its perigee point.
  • a station located in Libreville, as described above may serve as a back up TT & C station for Luxembourg, for example.
  • the location of the second TT&C station may take many factors into account besides optimum viewing locations.
  • the energy needed to launch a satellite into the baseline Tundra orbit is about the same as that required to launch a satellite into a GSO. However, the higher inclination may impose some restriction on the launch sites that may be used.
  • the size and the weight of the satellite launched may imply some difficulties, a dual launch may be possible on larger launch vehicles.
  • the orbit and satellite constellation design described herein may be used in conjunction with off-the-shelf spacecrafts with suitable flight-proven designs such as the SB3000 by Alcatel, Eurostar 3000 by Astrium, 601 HP/702 by Boeing, A2100 by Lockheed Martin, FS1300 by SS/Loral, etc.
  • Fig. 2 is map illustrating an enlarged view of the European coverage area 110 of Fig. 1.
  • the lower inclination Tundra orbit is a 24 hour orbit, thus allowing for fewer satellites in a constellation for full coverage.
  • a minimum of four satellites are needed to provide full coverage over the European service area 110, while requiring eight satellites for full redundancy.
  • the lower inclination orbit may provide full coverage with only three satellites in a constellation, and full redundancy may be achieved with six satellites.
  • Table 2 illustrates a comparison of the minimum elevation angles provided by a "left- hook" Molniya orbit (not shown) and a 53 degree Tundra orbit over the European coverage area.
  • the values presented in Table 2 are for illustrative purposes, and do not represent final values. Many values may be varied once the final beam configuration is known. However, the final values achieved may be similar to those presented in Table 2.
  • the lower inclination Tundra orbit provides higher minimum elevation angles in almost every city, which is beneficial for reception in urban and suburban areas. This is especially evident when the non-redundant or early entry systems (i.e., the 4-satellite Molniya and the 3-satellite lower inclination Tundra) are compared.
  • Table 2 Minimum Elevation Angles with full coverage (Optimized for Individual City)
  • the early entry systems may be in service for several years before the full configuration (i.e., the 8-satellite Molniya or the 6-satellite 53 degree Tundra) become operational. Therefore, the elevation angles obtained for the early entry systems are critical. In one embodiment, minimum elevation angles greater than 75 degrees are desirable due to a probability of blockage in urban areas.
  • the elevations presented in Table 2 are for optimum handover for each of the cities shown. When a single beam is used to cover Europe, the optimum elevations may not be achieved by either constellation. In a multiple beam system, if two or more satellites are in continual view of the coverage area, such as the area shown in Fig. 2, resource sharing is possible between the two satellites. As a result, various beams may be handed over one at a time as the incoming satellite reaches an optimum position for each particular beam.
  • Table 3 illustrates the minimum elevation angles achieved by a lower inclination tundra constellation for a single European beam. A minimum elevation angle of 70 degrees may be achieved in most areas of Europe for a single beam system in a lower inclination tundra orbit in the three-satellite constellation. For a six-satellite constellation, a single beam system offers elevation angles above 75 degrees in all of the areas evaluated.
  • the satellite constellation would be a six-satellite system.
  • the satellite constellation may be a three satellite system.
  • the three-satellite constellation may be the initial system that is deployed. Once three satellites have been launched and evaluated, a service provider may begin operation and start providing service. However, to protect itself against failure, three additional satellites may be launched at a later date to form the final operating configuration ("FOC"). In addition to providing redundancy, the FOC may also provide better elevation angles, allowing the service provider to expand the coverage area illustrated in Fig. 2.
  • Fig. 3a is a map illustrating the elevation angles achieved by a three-satellite constellation. It may be observed that although the three-satellite constellation provides adequate elevation angles in the core coverage area, the minimum elevation angles decrease when moving away from the part of the European coverage area 110 immediately below the apogee point. This may also be seen in Tables 2 and 3.
  • Fig. 3b shows the change in elevation angles for a three-satellite constellation over a one day period as seen by a user in Luxembourg.
  • Graph 300 shows that the minimum elevation of 78.4 degrees for the three-satellite constellation occurs for only a relatively short amount of time. The minimum elevation occurs during handover between satellites in the constellation.
  • the three-satellite constellation does not provide in-orbit redundancy for 24 hour operation. Thus, there is no in-orbit redundancy in case of catastrophic failure.
  • the three-satellite constellation requires a total payload handover.
  • the southern beams may be transferred about an hour earlier from the outgoing satellite to the incoming satellite.
  • the altitude of the satellite at handover may vary with then number of satellites in the constellation. For a three-satellite constellation, the handover may occur four hours from either side of the apogee. The altitude of the satellite at those points would be 45,000 km. Therefore, the satellite altitude over the coverage area illustrated in Fig. 2 may vary between 45,000km and the apogee altitude of 49,300 km.
  • Fig. 4a is a map illustrating the elevation angles achieved by a six-satellite constellation. Once the final three satellites are launched to complete the FOC, the region with elevation angles greater than 70 degrees is greatly expanded.
  • Fig. 4b shows the change in elevation angles for a six-satellite constellation over a one day period as seen by a user in Luxembourg.
  • the minimum elevation angle for the six-satellite constellation is 83.7 degrees.
  • Graph 400 shows that the handover occurs more smoothly in the six-satellite constellation.
  • the minimum elevation is closer to the average elevation in the six-satellite constellation than in the three-satellite constellation.
  • the FOC may provide full in-orbit redundancy.
  • the six-satellite system there are generally three satellites in view of the coverage area 110 at any one time. In the European coverage area illustrated in Fig. 2, the third satellite may provide good elevation angles for either the Spanish or Italian beams.
  • the FOC may also allow power sharing, as well as gradual handovers.
  • Gradual handover is possible since two satellites may be in view of the European coverage area 110 at all times, which in turn allows sharing of resources between the two satellites.
  • various beams may be handed over one at a time as the incoming satellite reaches an optimum position for that particular beam.
  • the payload power requirements may be ramped up or down for the satellites during handover, rather than an abrupt switching on or switching off of the broadcasting payload.
  • no satellite would need to handle the entire traffic load required for the coverage area 110. This allows each satellite's power requirements to be reduced.
  • the six-satellite constellation there would always be at least one beam that may be off loaded onto another satellite even when one satellite in the constellation had failed.
  • this off loading would imply some reduction in elevation angles over the coverage area 110 for a period of time.
  • the reduction in time may be about 5 degrees, and the period of time for the reduction would not exceed 15 minutes.
  • Nominal handovers in a six-satellite constellation may occur two hours before and two hours after the time the satellite is at apogee, which is at 49,300 km for the lower inclination Tundra orbit. These handovers occur more frequently than would be the case in a three satellite system because in the six-satellite system, there are twice as many satellites providing the coverage of the service area.
  • the handover may be optimized on a beam by beam basis. Thus, there are multiple handovers during each satellite's time over the coverage area 110.
  • the exact pattern of handovers may be optimized by the system designer or contractor, and would depend on the final beam pattern chosen. However, the range of altitudes would be about the same as for the three-satellite constellation.
  • a single satellite supports the entire coverage area 110 for extended periods of time. Since the six-satellite constellation has three satellites visible from the coverage area 110 at all times, the six-satellite system may be sized to require less power than the three-satellite system.
  • each of the six satellites may require less payload power than the payload power of each satellite of a three-satellite system, depending on the beam configuration chosen.
  • a satellite in the six-satellite constellation would need to carry somewhat less power than the power of a spacecraft or satellite of a three-satellite system, depending on beam layout.
  • a satellite may provide backup service in the six-satellite system with less power than a satellite in a three-satellite system. This allows a spacecraft design that would be at the lower end of most manufacturer's range of spacecraft buses.
  • the Tundra orbit is not constrained to the 63.4 degree inclination of the Molniya orbit.
  • the 63.4 degree inclination may still be optimum for cancellation of precession (or change) of the orbit's argument of perigee, so as the inclination is reduced, extra station keeping fuel may be needed.
  • the amount of extra fuel needed would be small, however, in the range of 10 to 30 kg over a ten year life.
  • the optimum inclination for the lower inclination Tundra orbit would be approximately in the range of 53 degrees to 56 degrees.
  • An inclination of 54.74 degrees represents an optimum inclination for minimizing drift in mean anomaly (a satellite's position in the orbit) for HEO orbits.
  • Table 4 shows the elevation angles in a 55 degree constellation. There is a small degradation in the minimum elevation for the 55 degree constellations in comparison to the minimum elevation for the 53 degree constellations, presented in Table 2.
  • tundra orbit is used by Sirius to cover the United States
  • the tundra orbit described herein has a lower inclination, which optimizes the elevation angles over the European coverage area 110.
  • the lower inclination Tundra orbit of the present invention insures enough high elevation in Europe to clear most building obstructions and allow significant quality of service without having to rely on an extended and costly terrestrial repeaters network.

Abstract

La présente invention concerne une orbite optimisée pour satellite de communication, qui peut comprendre une orbite fortement elliptique (HEO) présentant une inclinaison optimisée. Dans un mode de réalisation, l'orbite HEO peut présenter une variation d'inclinaison inférieure à celle d'une orbite Tundra, une éphéméride en goutte d'eau et une inclinaison approximative comprise entre 53 et 56 degrés. Le satellite de communication qui suit l'orbite HEO à inclinaison inférieure de l'invention peut faire partie d'une constellation de satellites. La constellation de satellites peut comprendre une constellation de trois ou de six satellites. Dans un mode de réalisation, la constellation de satellites peut être initialement mise en oeuvre sous la forme d'une constellation de trois satellites et trois satellites supplémentaires peuvent être lancés ultérieurement afin de former une constellation de six satellites.
EP02792458A 2001-12-21 2002-12-20 Orbite fortement elliptique pour satellites de communication Withdrawn EP1512233A2 (fr)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
US34183801P 2001-12-21 2001-12-21
US341838P 2001-12-21
PCT/US2002/040662 WO2003061141A2 (fr) 2001-12-21 2002-12-20 Orbite fortement elliptique pour satellites de communication

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EP1512233A2 true EP1512233A2 (fr) 2005-03-09

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EP (1) EP1512233A2 (fr)
AU (1) AU2002357912A1 (fr)
WO (1) WO2003061141A2 (fr)

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7669803B2 (en) 2004-12-07 2010-03-02 Lockheed Martin Corporation Optimized land mobile satellite system for north american coverage
US7624948B2 (en) 2004-12-07 2009-12-01 Lockheed Martin Corporation Optimized land mobile satellite configuration and steering method
FR2918827B1 (fr) * 2007-07-10 2009-10-16 Astrium Sas Systeme pour le positionnement d'un utilisateur terrestre.
CA2716174C (fr) * 2010-10-01 2019-11-26 Telesat Canada Systeme satellitaire
US9550585B2 (en) 2014-06-09 2017-01-24 Space Systems/Loral, Llc Highly inclined elliptical orbit de-orbit techniques
US9365299B2 (en) 2014-06-09 2016-06-14 Space Systems/Loral, Llc Highly inclined elliptical orbit launch and orbit acquisition techniques

Family Cites Families (2)

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Publication number Priority date Publication date Assignee Title
US5439190A (en) * 1991-04-22 1995-08-08 Trw Inc. Medium-earth-altitude satellite-based cellular telecommunications
US5525999A (en) * 1994-11-28 1996-06-11 Motorola, Inc. Multi-receiver master control station for differential GPS and method

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See references of WO03061141A3 *

Also Published As

Publication number Publication date
WO2003061141A2 (fr) 2003-07-24
AU2002357912A8 (en) 2003-07-30
AU2002357912A1 (en) 2003-07-30
WO2003061141A3 (fr) 2003-12-18

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