Abstract: A method for implementing a satellite fleet includes launching a group of satellites within a launch vehicle. In an embodiment, the satellites are structurally connected together through satellite outer load paths. After separation from the launch vehicle, nodal separation between the satellites is established by allowing one or more of the satellites to drift at one or more orbits having apogee altitudes below an operational orbit apogee altitude. A satellite is maintained in an ecliptic normal attitude during its operational life, in an embodiment. The satellite's orbit is efficiently maintained by a combination of axial, radial, and canted thrusters, in an embodiment. Satellite embodiments include a payload subsystem, a bus subsystem, an outer load path support structure, antenna assembly orientation mechanisms, an attitude control subsystem adapted to maintain the satellite in the ecliptic normal attitude, and an orbit maintenance/propulsion subsystem adapted to maintain the satellite's orbit.

Abstract: The invention concerns an observation satellite (1) which is intended to be placed in orbit around a celestial body (2), and which comprises a reflecting device (5), a receiving device (6), a linking mechanical system (19), the whole forming a capture system (3) which is suitable to be able to orient the capture system (3) by gravity gradient in an aiming position in which the electromagnetic radiation corresponding to the information to be captured is received. The invention extends to an observation method and a fleet of such satellites (1).

Abstract: A spacecraft system that includes a primary space vehicle and a secondary space vehicle, both of which are designed to optimize payload capacity and launch weight of the primary space vehicle. The primary and secondary space vehicles combine to form an on-orbit space vehicle capable of performing functions and maneuvers that exceed the physical capabilities of the primary space vehicle at the time of its launch. The spacecraft system is designed to minimize propellant containment-related disturbances while maintaining a standard level of functionality. The primary space vehicle is designed to be incapable of independently performing a propellant-intensive orbit change maneuver. Instead the primary space vehicle is designed to couple to a secondary space vehicle having propellant and thrust capability sufficient to perform an orbit change maneuver when the primary and secondary space vehicles are coupled.

Abstract: An algorithm for deducing all possible scenarios of satellite members and possibilities thereof in a low earth orbiting (LEO) satellite constellation is described, which is achieved mainly by relying on the spherical geometry analysis and probabilities and statistics technologies, in an attempt to rapidly and precisely obtain the concerned scenarios and possibilities thereof observed on the earth ground. With any user-defined orbital parameters and a position of an observation station for the scenarios on the earth ground inputted, all the possible scenarios and possibilities thereof can be obtained with the algorithm.

Abstract: A method for implementing a satellite fleet includes launching a group of satellites within a launch vehicle. In an embodiment, the satellites are structurally connected together through satellite outer load paths. After separation from the launch vehicle, nodal separation between the satellites is established by allowing one or more of the satellites to drift at one or more orbits having apogee altitudes below an operational orbit apogee altitude. A satellite is maintained in an ecliptic normal attitude during its operational life, in an embodiment. The satellite's orbit is efficiently maintained by a combination of axial, radial, and canted thrusters, in an embodiment. Satellite embodiments include a payload subsystem, a bus subsystem, an outer load path support structure, antenna assembly orientation mechanisms, an attitude control subsystem adapted to maintain the satellite in the ecliptic normal attitude, and an orbit maintenance/propulsion subsystem adapted to maintain the satellite's orbit.

Abstract: A method for sensing and monitoring ICBM threats using a constellation of earth orbiting satellites. The method includes arranging a first and a second plurality of threat sensing satellites in earth orbit. The first plurality is selected to have a first prograde orbit plane with a first inclination angle between about 30° to 70° with respect to the earth's equator and an Argument of Perigee equal to about 270°. The second plurality is selected to have a first retrograde orbit plane with a second inclination angle between about 110° to 150° with respect to the earth's equator and an Argument of Perigee equal to about 270°. The second inclination angle is selected to be a symmetric conjugate of the first inclination angle. The first and second plurality of threat sensing satellites are distributed respectively within the first prograde and first retrograde orbits with a Walker distribution of T/P/1.

Abstract: A satellite constellation in one example comprises a first set of satellites configured in a first sub-constellation defined by a first set of orbital elements. The satellite constellation further comprises a second set of satellites configured in a second sub-constellation defined by the first set of orbital elements with a longitudinal offset.

Abstract: The present invention relates to an aircraft and to a method of getting the aircraft onto station. According to the invention, said aircraft (1) includes propulsion means (2) capable only of enabling said aircraft (1) to move and to orient itself at high altitude, and said aircraft (1) is taken to its station in the high atmosphere, in particular in the stratosphere, by means of an independent transporter (3).

Abstract: A system and method to enhance attitude estimation for transfer orbit maneuvers of a spinning satellite includes a transfer orbit earth sensor that can receive and convert an infrared radiance to an analog signal. The analog signal is provided to a hardware chord processor and an analog-to-digital converter. The hardware chord processor determines an earth chord length based on the analog signal. The analog signal is sampled in the analog-to-digital converter to provide a digital signal to a digital earth sensor pre-processor, which reshapes the digital signal to provide a pre-processed signal having nearly distinct peaks. A software chord processor is provided with the pre-processed signal. The software chord processor locates the peaks in the pre-processed signal and determines an earth chord length.

Abstract: A new approach for designing satellite constellations whereby each satellite follows a common ground track is being proposed for performing high-precision change detection imagery for long periods of time. By precisely prescribing the orbital parameters, i.e., the relationship between the right ascension of the ascending nodes (RAAN) and the phase angle difference between successive satellites, for example, sharpened change detection images may be taken from successive satellites in the constellation without the need to process out blurring by any special image re-working software. The relationship between the orbital parameters of the satellites is precisely tuned to the earth's rotation rate for the altitude of the satellites. A reduction in the total satellite count is achievable due to tiling the satellite coverage in near optimal arrangements.

Abstract: A new approach for designing satellite constellations whereby the satellites are distributed into contiguous groups forming longitudinal lengths at least equal to 360 degrees. Gaps between the contiguous groups are uniformly distributed along the tracks. These contiguous groups of satellites may be populated by having multiple satellites onboard a single launch vehicle, whereby the satellites are nested into multiple common ground track sub-constellations to form a constellation with multiple common ground tracks.

Abstract: A new method of constellation design based on combining repeating ground tracks with common ground tracks and/or sun-synchronous orbits is used for performing high-precision change detection imagery for longer periods of time while minimizing battery usage. By precisely prescribing the orbital parameters, sharpened change detection images may be taken without the need to process out blurring by any special image re-working software. The relationship between the orbital parameters of the satellites is precisely tuned to the earth's rotation rate for the altitude of the satellites. The unique set of earth orbits minimizes satellite battery requirement while optimizing the ability to perform coherent change detection (CCD).

Abstract: A vehicle (12) including a control system (18) is used for controlling vehicle attitude or angular velocity (38). The processor (24) is coupled to a star sensor or tracker (22) and a memory (30) that may include a star catalog (32), and an exclusion list (36). The exclusion list (36), a list of stars to be temporarily excluded from consideration when determining attitude or angular velocity or relative alignment of star sensors or trackers, is calculated on board. Such a calculation prevents the necessity for a costly, periodic, ground calculation and upload of such data. By manipulating the star catalog, or sub-catalogs derived from said catalog, based upon the exclusion list (36), measurements of such excluded stars are prevented from corrupting the attitude or angular velocity or alignment estimates formulated on board. The system uses multiple stayout zones for excluding stars from the exclusion list.

Abstract: A method for deploying multiple spacecraft is disclosed. The method can be used in a situation where a first celestial body is being orbited by a second celestial body. The spacecraft are loaded onto a single spaceship that contains the multiple spacecraft and the spacecraft is launched from the second celestial body towards a third celestial body. The spacecraft are separated from each other while in route to the third celestial body. Each of the spacecraft is then subjected to the gravitational field of the third celestial body and each of the spacecraft assumes a different, independent orbit about the first celestial body. In those situations where the spacecraft are launched from Earth, the Sun can act as the first celestial body, the Earth can act as the second celestial body and the Moon can act as the third celestial body.

Abstract: A method and system for maximizing satellite coverage at predetermined local times for a set of predetermined geographic location includes a processor operative to determine a period of rotation for each of the desired satellites in the satellite constellation. The processor also determines a time dependent coverage of the satellite constellation based on the period of rotation and the trajectory of each of the desired satellites. The trajectories of the desired satellites are tilted until the satellite constellation provides maximum coverage at the predetermined local times for the set of predetermined geographic locations. If a new satellite constellation is being designed, command signals are programmed into a computer of a launch vehicle containing the modified trajectory. If an existing satellite constellation is being modified, the ground station transmits command signals to the satellites for modifying the trajectory of the satellites in accordance with the tilted trajectory.

Abstract: A sparing system for space vehicle constellations. A plurality of space vehicles arranged in at least one mission orbit plane. The number of space vehicles is at least one greater than a minimum number necessary to provide at least a minimum level of service necessary to carry out a constellation mission. If one of the space vehicles fails the remaining space vehicles are repositioned in the mission orbit plane to provide the minimum level of service necessary to carry out a constellation mission.