Abstract: Apparatus for a space platform comprises a magnetic field generator and/or an electric field generator respectively configured to generate fields for influencing the trajectory of one or more items of space debris passing within a region of the apparatus. A system comprising a plurality of such space platforms may be placed in an orbit proximal to an orbit containing space debris. An individual space platform or system thereof may be used in a method for displacing earth orbital space debris out of the orbit and either towards the Earth's atmosphere where it is likely to be destroyed by the burning on re-entry into the atmosphere, or into a safer orbit from which it may be collected or in which it may be left.

Abstract: A method and a device is provided for the optimization of the mass of a satellite. The method includes: a step of calculation of an elliptical second orbit obtained by rotation of a first orbit about an axis connecting the periapsis and the apoapsis, the elliptical second orbit being associated with a second maximum eclipse duration less than a first maximum eclipse duration; a step of determination of a manoeuvre enabling the satellite to move to the second orbit; and a step of calculation of a second battery mass making it possible to maintain the satellite in operation during the second maximum eclipse duration and of calculation of a mass of fuel necessary to effect the manoeuvre.

Abstract: A method, a member satellite, and a tangible machine-readable medium are disclosed. An inter-satellite link subsystem 472 may maintain a fixed communication link between the member satellite and a partner satellite of the medium earth orbit centric satellite constellation in a regular, wraparound symmetric, spatially dimensional network. A terrestrial linking subsystem 476 may create a terrestrial link to a ground terminal.

Abstract: An improved approach to satellite-based navigation (e.g., GPS) is provided. In one embodiment, a method includes determining a nominal orbital path of a navigation satellite. The method also includes transmitting ephemeris data corresponding to the nominal orbital path from the navigation satellite to a plurality of navigation devices. The method further includes determining an actual orbital path of the navigation satellite locally at the navigation satellite. In addition, the method includes determining a deviation between the actual orbital path and the nominal orbital path locally at the navigation satellite. The method also includes autonomously adjusting the actual orbital path locally at the navigation satellite to reduce the deviation between the actual orbital path and the nominal orbital path.

Abstract: An apparatus for orbital change of a satellite incorporates a capture mechanism for space debris and a tether connecting the capture mechanism to a satellite. The tether is extendable to position the capture mechanism relative to the space debris. A controller is employed for timed release of the space debris by the capture mechanism for orbital change by the satellite.

Abstract: This project is related to the inventor, Cyrus Nejat, previous patent Application Number: U.S. Ser. No. 12/011,127. In this study, the main discussion emphasizes on building a spacecraft based on mathematics. The orientation of the low thrust engine spacecraft was structured in such a way that the Kinematics and kinetics equations be written for large elements, by means of the purposed Nejat Kinetics Expansions and Nejat Kinematics Expansions. These equations were used to design Nejat Space Station. Simplified CN (Cyrus Nejat) Equations of Motion used for the Special Spacecraft Rout in Space.

Abstract: A system for defending a planet having a satellite against impact from an incoming body includes an explosive system and a propulsion system. The explosive system is adapted for deployment on the satellite and detonation thereon with sufficient explosive force to produce at least one ejectum to which is imparted a velocity increment sufficient for the ejectum to exceed the satellite's escape velocity and enter orbit about the planet. The propulsion system is adapted to be secured to at least one ejectum and impart a velocity increment to the ejectum sufficient to leave orbit about the planet and enter an orbit intercepting the incoming body. The system is used in a planetary defense method in which the explosive system is deployed at a deployment site on the satellite and detonated to produce at least one ejectum having a velocity that exceeds the escape velocity of the satellite.

Abstract: A constellation, including a plurality of spacecraft, including a first, second and third spacecraft, each of the plurality of spacecraft including a broadcast capability, and each of the plurality of spacecraft in its own approximately 24-hour orbit. Each of the orbits has a substantially teardrop-shaped or oval-shaped ground track, is optimized based upon elevation angle or probability of signal availability, and has an apogee longitude of approximately 90° west to approximately 100° west. Each of the orbits has a semi-major axis of approximately 42,164 kilometers, an argument of perigee of approximately 270° , an inclination of approximately 40° to approximately 60° , and an eccentricity of approximately 0.16 to approximately 0.4. The orbits of each of the plurality of spacecraft are selected to bring each of the spacecraft to apogee at time increments of approximately eight hours.

Abstract: Spacecraft vehicles and methods for providing reorbiters capable of moving target objects are provided. More particularly, a reorbiter utilizing electrostatic or Coulomb force for acting on target objects and for moving the target objects into new orbits or altitudes are provided. The reorbiter may establish an electrostatic force by controlling the electrical potential of the reorbiter through active charge emission. The target object may acquire an electrical potential due to its interaction with the space plasma and photoelectrons or the reorbiter may impart additional charge to the target object. After establishing the electrostatic force, a propulsion system of the reorbiter is operated to provide thrust. Thrust is directed to change the orbit and/or position of the target object that is being moved by the electrostatic force between the reorbiter and the target object.

Abstract: Space tug vehicles and methods for providing space tugs and moving target vehicles are provided. More particularly, a space tug utilizing electrostatic or Coulomb force for acting on target vehicles and for moving the target vehicles into new orbits or altitudes are provided. The space tug may establish an attractive electrostatic force by controlling the electrical potential of the space tug so that it is opposite the electrical potential of a target vehicle. The target vehicle may acquire an absolute electrical potential due to its interaction with the space plasma and photoelectrons or the space tug may impart additional charge to the target vehicle. After establishing the attractive electrostatic force, a propulsion system of the space tug is operated to provide thrust. Thrust is directed to change the orbit and/or position of the target vehicle that is being pulled towards the space tug by the electrostatic force.

Abstract: A method and apparatus for maneuvering a satellite in orbit to alternately optimize the collection of solar energy and to take sensor data of terrestrial objects is disclosed The longitudinal axis of a large payload package is oriented perpendicular to the orbital plane to minimize the disturbance torque due to gravity gradient, and to allow simple rotation about the axis for attitude change between optimal Sun and optimal ground coverage.

Abstract: Spacecraft payload orientation steering is provided for an orbiting spacecraft in motion along an orbit track around a celestial body, the orbit track having a nominal inclination with respect to an equatorial orbit, a substantial eccentricity, and a drift angle with respect to the nominal inclination. Coordinates of an optimal payload target location as a function of a spacecraft position along the orbit track are determined, the target location being on the surface of the celestial body and having a substantial motion with respect to the surface and with respect to a spacecraft nadir. A payload of the spacecraft is substantially aligned with the determined coordinates by steering the satellite body to correct for at least one of the inclination drift angle, and the eccentricity, thereby adjusting the spacecraft orientation as a function of the spacecraft position along the orbit track.

Abstract: In a method for injecting a plurality of spacecraft into different circum-earth or interplanetary orbits individually in a single launch, a plurality of spacecraft coupled to an assist module are injected into an interplanetary orbit having a periodicity synchronous with the earth's revolution period. Then, in a maneuver allowing the assist module to re-counter with and pass near to the earth (Earth swing-by), the assist module appropriately performs an orbital change maneuver and a separation maneuver for each of the spacecraft in a sequential order. Through these maneuvers, a closest-approach point to the earth is changed for each of the spacecraft so as to drastically change a subsequent orbital element for each of the spacecraft. The assist module takes a sufficient time to determine a target orbit for each of the spacecraft with a high degree of accuracy until a half month to several days before a closest-approach time in the Earth swing-by.

Abstract: A calculating method 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 calculating method.

Abstract: Methods and apparatus are provided for active motion damping of a spacecraft whose attitude control system (SACS) has experienced a partial failure. In a preferred embodiment, the apparatus comprises, an attitude control processor (ACP), a motion damping controller (MDC) coupled to the ACP, a control moment gyro (CMG) comprising a gimbal loop controller (GLC) coupled to the MDC and a gimbal motor with input coupled to the GLC and gyroscopic torque output coupled to the satellite. Normally, a gimbal rate command (GRC) from the ACP passes to the GLC for execution to rotate the CMG gimbal and the coupled satellite. If the ACP output becomes invalid, the MDC assumes control and can modify CMG operation in several ways to safely decelerate a moving satellite without hazardous mechanical stress. Following such active motion damping, the CMG remains quiescent until a RESET is received whereby further valid GRCs are then executable.

Abstract: A control system of a spacecraft for controlling two or more sets of collinear control moment gyroscopes (CMGs) comprises an attitude control system. The attitude control system is configured to receive a command to adjust an orientation of the spacecraft, determine an offset for a momentum disk for each of the two or more sets of CMGs that maximizes torque, determine a momentum needed from the two or more sets of CMGs to adjust the orientation of the spacecraft, and calculate a total torque needed by taking the derivative of the momentum. The control system further comprises a momentum actuator control processor coupled to the attitude control system, the momentum actuator control processor configured to calculate a required gimbal movement for each of the CMGs in each of the two or more sets of collinear CMGs from total torque.

Abstract: A constellation, including a plurality of spacecraft, including a first, second and third spacecraft, each of the plurality of spacecraft including a broadcast capability, and each of the plurality of spacecraft in its own approximately 24-hour orbit. Each of the orbits has a substantially teardrop-shaped or oval-shaped ground track, is optimized based upon elevation angle or probability of signal availability, and has an apogee longitude of approximately 90° west to approximately 100° west. Each of the orbits has a semi-major axis of approximately 42,164 kilometers, an argument of perigee of approximately 270°, an inclination of approximately 40° to approximately 60°, and an eccentricity of approximately 0.16 to approximately 0.4. The orbits of each of the plurality of spacecraft are selected to bring each of the spacecraft to apogee at time increments of approximately eight hours.

Abstract: A space based orbital kinetic energy weapon system and method of using same is provided. The space based orbital kinetic energy weapon system includes a satellite having a control system configured to maintain an orbit in outer space around the earth and to deorbit the satellite on a desired trajectory corresponding to an earth based target upon a command, and a projectile object operably connected to the satellite. The projectile object includes a dense mass and a heat shield operably surrounding the mass such that at least a portion of the mass survives reentry into the earth's atmosphere and strikes the earth based target delivering its kinetic energy.

Abstract: A method of insuring against launch failure is provided to reduce the business risk to the satellite owner and insurance underwriters against satellite launch failure. The launch insurance includes provision for guaranteeing a rapid response recovery mission, preferably at no cost, in the event that a satellite suffers a boost failure during launch resulting in a fully-functional satellite being launched to an unintended inoperable orbit. Preferably, the launch insurance also provides for additional insurance against failure of the recovery mission to provide compensation to the satellite owner in the event that the recovery mission is also unsuccessful.

Abstract: Enhanced translational thrusting is provided by reaction engines configured to permit translational thrusting off or through the center of gravity of a spacecraft or other vehicle. Among other applications, this approach is useful for a Satellite Life Extension System (SLES) that provides maintenance services to orbiting satellites. By attaching to the satellite and conducting maneuvers to maintain its operational orbit and attitude, the SLES increases the working lifetime of the satellite. Since the engines of the SLES are redundant, the failure of a single engine will not jeopardize the overall success of the mission.

Abstract: This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR, and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time as well as the ability to intervene using manual override to teleoperate the robot.

Abstract: According to one embodiment of the invention, a stationkeeping method for a geostationary satellite includes determining a gravitational force of the sun on the satellite at a beginning of a stationkeeping operation and a gravitational force of the moon on the satellite at the beginning of the stationkeeping operation. An initial inclination vector of the satellite is determined at the beginning of the stationkeeping operation that accounts for a first set of one or more perturbations affecting the orbit of the satellite. A maneuver strategy is determined to correct for a second set of one or more perturbations affecting the orbit of the satellite without accounting for the first set of one or more perturbations affecting the orbit of the satellite. Finally, a maneuver is performed on the satellite according to the maneuver strategy.

Abstract: This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.

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: This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.

Abstract: This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.

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.