Abstract: A constellation of Earth-orbiting spacecraft, the constellation having an orbital maneuver lifetime life (OML), includes a first spacecraft disposed in a first orbit and a second spacecraft disposed in a second orbit, each of orbit being substantially circular with a radius of approximately 42,164 km and having a respective inclination with respect to the equator specified within a range of 10° to 20°. The first orbit has, at beginning of life (BOL), a first right ascension of ascending node (BOL-RAAN1) and the second orbit has, at BOL, a second RAAN (BOL-RAAN2) the BOL-RAAN1 and the BOL-RAAN2 being separated by a first angular separation ?-RAAN1. A first stationkeeping delta-V (?V1) applied over the OML to the first spacecraft, in combination with a second delta-V (?V2) applied over the OML to the second spacecraft, maintains the ?-RAAN1 approximately constant and an actual inclination within specification, and ?V1 approximately equals ?V2.

Abstract: A spacecraft includes at least a first thruster support mechanism (TSM) and a second TSM, each TSM including a pointing arrangement, an elongated structural member and thruster for performing orbit raising north-south stationkeeping, east-west stationkeeping, and momentum management. A first pointing arrangement is articulable only by way of first and second revolute joints, the first revolute joint being rotatable about a first axis fixed with respect to the spacecraft. The second pointing arrangement is articulable only by way of third and fourth revolute joints, the third revolute joint being rotatable about a third axis fixed with respect to the spacecraft. The first axis and the third axis are asymmetrically arranged with respect to a spacecraft coordinate system origin such that the first and third axis are at acute angles to a spacecraft pitch axis and the acute angle of the first axis is less than that of the third axis.

Abstract: A constellation of Earth-orbiting spacecraft includes a first spacecraft disposed in a first orbit, a second spacecraft disposed in a second orbit, and a third spacecraft disposed in a third orbit. Each of the first orbit, the second orbit and the third orbit is substantially circular with a radius of approximately 42,164 km, and has a specified inclination with respect to the equator within a range of 5° to 20°. The first orbit has a first right ascension of ascending node RAAN1, the second orbit has a second RAAN (RAAN2) approximately equal to RAAN1+120°, and the third orbit has a third RAAN (RAAN3) approximately equal to RAAN1+240°. A fourth spacecraft is disposed in a fourth orbit that has a period of approximately one sidereal day, an inclination of less than 2°, a perigee altitude of at least 8000 km, and an eccentricity between approximately 0.4 and 0.66.

Abstract: A constellation of Earth-orbiting spacecraft, the constellation having an orbital maneuver lifetime life (OML), includes a first spacecraft disposed in a first orbit and a second spacecraft disposed in a second orbit, each of orbit being substantially circular with a radius of approximately 42,164 km and having a respective inclination with respect to the equator specified within a range of 10° to 20°. The first orbit has, at beginning of life (BOL), a first right ascension of ascending node (BOL-RAAN1) and the second orbit has, at BOL, a second RAAN (BOL-RAAN2) the BOL-RAAN1 and the BOL-RAAN2 being separated by a first angular separation ?-RAAN1. A first stationkeeping delta-V (?V1) applied over the OML to the first spacecraft, in combination with a second delta-V (?V2) applied over the OML to the second spacecraft, maintains the ?-RAAN1 approximately constant and an actual inclination within specification, and ?V1 approximately equals ?V2.

Abstract: A constellation of Earth-orbiting spacecraft includes a first spacecraft disposed in a first orbit, a second spacecraft disposed in a second orbit, and a third spacecraft disposed in a third orbit. Each of the first orbit, the second orbit and the third orbit is substantially circular with a radius of approximately 42,164 km, and has a specified inclination with respect to the equator within a range of 5° to 20°. The first orbit has a first right ascension of ascending node RAAN1, the second orbit has a second RAAN (RAAN2) approximately equal to RAAN1+120°, and the third orbit has a third RAAN (RAAN3) approximately equal to RAAN1+240°. A fourth spacecraft is disposed in a fourth orbit that has a period of approximately one sidereal day, an inclination of less than 2°, a perigee altitude of at least 8000 km, and an eccentricity between approximately 0.4 and 0.66.

Abstract: There is provided a method and system for simultaneous north-south stationkeeping and 3-axis momentum management for a geosynchronous orbiting spacecraft. The spacecraft has a first thruster, a second thruster, and at least three momentum wheels mounted on-board the spacecraft. The first and second thrusters are mounted adjacent to a north and a south face of the spacecraft, respectively, on an anti-earth side of the spacecraft and are aligned to produce thrust vectors slightly off from the spacecraft's center of mass. The first and second thrusters are independently fired at predetermined positions along the geosynchronous orbital path of the spacecraft. The thrust vectors provide attitude and orbital adjustment needed to correct north-south orbital position deviation and provide torque needed to desaturate the stored angular momentum of the momentum wheels.

Abstract: A satellite perigee velocity augmentation method wherein a satellite is steered throughout the perigee velocity augmentation maneuver such that its thrust vector is always oriented approximately opposite to the direction of motion of the satellite. The satellite is pitched around a predetermined axis (e.g. the Y-axis) such that the thrust vector is always pointing approximately opposite to the direction of motion of the satellite. The satellite is steered during the perigee velocity augmentation maneuver in order to achieve the desired attitude wherein the thrust vector is always oriented approximately opposite to the direction of motion of the satellite is accomplished by pitching the satellite around the predetermined axis such that the thrust vector is always pointing approximately opposite to the direction of motion of the satellite. By steering the satellite in accordance with the present perigee velocity augmentation method, overall efficiency achieved is increased.