Class / Patent application number | Description | Number of patent applications / Date published |
244158400 | Spacecraft formation, orbit, or interplanetary path | 81 |
20080237399 | SATELLITES AND SATELLITE FLEET IMPLEMENTATION METHODS AND APPARATUS - 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. | 10-02-2008 |
20080251645 | ALGORITH FOR DEDUCING POSSIBILITIES OF ALL POSSIBLE SCENARIOS OF SATELLITE MEMBERS IN LOW EARTH SATELLITE CONSTELLATION - 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. | 10-16-2008 |
20100006704 | Method for Lightening the Weight of Fuel Stowed Onboard During an Interplanetary Mission - The method for lightening the weight of fuel stowed onboard during an interplanetary mission is characterized in that it consists:
| 01-14-2010 |
20100032528 | SYSTEM FOR CONTROLLING THE DEPLOYMENT OF SPACECRAFT REQUIRED TO FLY IN FORMATION, BY SIMULTANEOUS AND HIGH-PRECISION DETERMINATION OF THEIR POSITIONS - A system is dedicated to the control of the deployment of at least two spacecraft (ES | 02-11-2010 |
20100108818 | OPTIMIZED LAND MOBILE SATELLITE SYSTEM FOR NORTH AMERICAN COVERAGE - 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. | 05-06-2010 |
20110049302 | Retro-Geo Spinning Satellite Utilizing Time Delay Integration (TDI) for Geosynchronous Surveillance - Geosynchronous surveillance is conducted by injecting one or more observer satellites into a retro sub or super geosynchronous orbit at approximately zero inclination. The observer satellite spins about an approximately North-South axis in an Earth frame of reference to sweep a sensor's FOV around the geobelt. Sensor time delay integration (TDI) is synchronized to the observer satellite's spin-rate and possibly the sum of the spin-rate and the target inertial LOS rate to realize longer integration times. This approach facilitates faster scans of the entire geobelt, more timely updates of the catalog of tracked objects and resolution of small and closely spaced objects. An inexpensive small-aperture body-fixed visible sensor may be used. | 03-03-2011 |
20110226907 | OPTIMAL SPACE SITUATIONAL AWARENESS SYSTEM - A satellite system for observing space objects includes two or more satellites positioned in an Earth orbit and configured to observe objects in various orbits including those viewed (i) against the Earth's background; (ii) against a sunlit Earth background; and (iii) against a space background. An electromagnetic sensor may be provided on at least one of the satellites that is responsive to electromagnetic radiation having a wavelength that discriminates against substantial reflection of electromagnetic radiation from the Earth's atmosphere to observe the space object. A method of observing a space object using a satellite system is also disclosed. | 09-22-2011 |
20120168566 | System for global earth navigation using inclined geosynchronous orbit satellite - A global earth navigation satellite system may be provided. The global earth navigation satellite system may include a group of satellites including at least one inclined geosynchronous satellite disposed in at least one orbital plane distinguished based on an interval determined based on a longitudinal coordinate of the earth, and the at least one inclined geosynchronous satellite may be disposed in the at least one orbital plane at predetermined intervals, and may revolve around the earth at a predetermined inclination of satellite orbit so as to provide, over time, geometric shape change information associated with the earth, geometric shape change information associated with a low earth orbit satellite, and geometric shape change information associated with a geostationary satellite. | 07-05-2012 |
20120223189 | APPARATUS AND METHOD FOR GENERATING FLASH OF LIGHT TOWARD EARTH BY MEANS OF REFLECTION OF SUNLIGHT | 09-06-2012 |
20120261513 | Coarse and fine projective optical metrology system - Described herein is a projective optical metrology system including: a light target formed by a first number of light sources having a pre-set spatial arrangement; and an optical unit including an optoelectronic image sensor, which receives a light signal coming from the light target and defines two different optical paths for the light signal towards the optoelectronic image sensor. The two optical paths are such that the light signal forms on the optoelectronic image sensor at most an image of the light target that can be processed for determining at least one quantity indicating the mutual arrangement between the light target and the optical unit. | 10-18-2012 |
20120273621 | METHODS FOR OPTIMIZING THE PERFORMANCE, COST AND CONSTELLATION DESIGN OF SATELLITES FOR FULL AND PARTIAL EARTH COVERAGE - A system and method for highly efficient constellations of satellites which give single, double, . . . k-fold redundant full earth imaging coverage, or k-fold coverage for latitudes greater than any selected latitude is given for remote sensing instruments in short periods of time, i.e., continuous coverage, as a function of the parameters of the satellite and the remote sensing instrument for many different types of orbits. A high data rate satellite communication system and method for use with small, mobile cell phone receiving and transmitting stations is also provided. Satellite instrument performance models, full and partial satellite constellation models, and satellite cost models are disclosed and used to optimize the design of satellite systems with vastly improved performance and lower cost over current major satellite systems. | 11-01-2012 |
20130032673 | SOLAR POWER SATELLITE SYSTEM FOR TRANSMITTING MICROWAVE ENERGY TO THE EARTH AND METHOD OF ARRANGING A SOLAR POWER SATELLITE SYSTEM ABOUT THE SUN FOR SAME - Solar power satellite system for transmitting microwave energy to the earth and a method of arranging the solar power satellite system about the sun for same. The solar power satellite system comprises a space-based power generation unit disposed in a planetary orbit about the sun. A photovoltaic cell on the space-based power generation unit collects solar energy that is then converted to microwave energy to be beamed to the earth. A ground-based rectenna receives the microwave energy and converts the microwave energy to electricity that is transmitted to an end user. The solar power satellite system and method provides electrical power on earth day or night and regardless of atmospheric conditions. Also, surface area of the solar panel on the space-based power generation unit orbiting about the sun is much less than the surface area required of a ground-based solar panel or a solar panel in earth orbit. | 02-07-2013 |
20140367523 | SPACE DEBRIS REMOVING DEVICE AND SPACE DEBRIS REMOVING METHOD - Provided are a space debris removing device and a method which enable easy installation of a deceleration device to space debris undergoing a tumbling motion. The space debris removing device includes: a propulsion device ( | 12-18-2014 |
20150102173 | METHOD OF SOLAR OCCULTATION - A method of solar occultation, and in particular solar coronagraphy, employing a spacecraft | 04-16-2015 |
20150375876 | Three dimensional imaging arrangement - A spacecraft arrangement having sensors for providing the three dimensional imaging of space object to be imaged and having a plurality of spaced apart nodes defining an open imaging area therebetween through which imaging area the space object to be imaged passes. | 12-31-2015 |
20160101879 | Geostationary polar satellite - A Satellite Stationary above the North or South Geometric Pole. | 04-14-2016 |
20160380580 | Large-Scale Space-Based Solar Power Station: Multi-Scale Modular Space Power - A space-based solar power station, a power generating satellite module and/or a method for collecting solar radiation and transmitting power generated using electrical current produced therefrom is provided. Each solar power station includes a plurality of satellite modules. The plurality of satellite modules each include a plurality of modular power generation tiles including a photovoltaic solar radiation collector, a power transmitter and associated control electronics. The power transmitters can be coordinated as a phased array and the power generated by the phased array is transmitted to one or more power receivers to achieve remote wireless power generation and delivery. Each satellite module may be formed of a compactable structure capable of reducing the payload area required to deliver the satellite module to an orbital formation within the space-based solar power station. | 12-29-2016 |
20180022474 | Constellation Configuration for Constellations having a Large Number of LEO Satellites | 01-25-2018 |
20220135255 | SPACE SURVEILLANCE ORBIT - A satellite system includes a satellite in an orbit that is configured to reduce a number of exclusion regions and improve the observation coverage of resident space objects (RSOs) positioned in near Earth orbits. The satellite system includes at least one satellite positioned in a sun synchronous orbit (SSO) with a noon/midnight nodal crossing. The altitude of the SSO is between 1000 and 2000 kilometers and the satellite includes at least one sensor arranged on the satellite that is configured for detection, tracking, and/or identification. Using the noon/midnight nodal crossing is advantageous in that three main exclusion regions, the sun, eclipse, and Earth exclusion regions, are combined into only two exclusion regions for improved performance of the satellite system in observing RSOs. | 05-05-2022 |
244158500 | Orbit insertion | 21 |
20090206204 | SPIN-STABILIZED LANDER - The invention provides a novel, low-cost, spin-stabilized lander architecture capable (with appropriate system scaling tailored to the attributes of the target) of performing a soft-landing on a solar-system body such as Earth's Moon, Mars, Venus, the moons of Mars, Jupiter, Saturn, Uranus and Neptune, selected near-Earth and main-belt asteroids, comets and Kuiper belt objects and even large human-made objects, and also moving about on the surface of the target solar-system body after the initial landing in movement akin to hopping. | 08-20-2009 |
20090230249 | Method of Launching into Operational Orbit an Artificial Satellite and Associated Propulsion Device - A method is disclosed for placing a satellite in an operational orbit The satellite is equipped with its own satellite propulsion system as well as a detachable separate propulsion device The satellite and separate propulsion device are launched into a transfer orbit by means of a space launcher The separate propulsion device is controlled by a satellite. The satellite is transferred from the transfer orbit to an intermediate orbit by means of the separate propulsion device. The separate propulsion device is separated from the satellite in the intermediate orbit. The satellite then enters and operational orbit from the intermediate orbit by means of its own satellite propulsion system The intermediate orbit is disposed between the transfer and operational orbits, and is in relatively close proximity to the operational orbit but is far enough away from the operational orbit to prevent possible interferences | 09-17-2009 |
20090321578 | Planetary impact defense system - 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. | 12-31-2009 |
20100108819 | METHOD FOR DESIGNING AN ORBIT OF A SPACECRAFT - A method is disclosed for designing an orbit of a spacecraft which allows the spacecraft to take a small-radius halo orbit near a Lagrange point while avoiding the prohibited zone where the spacecraft may be shadowed or might be prevented from making communication. The method makes it possible to have a closed orbit although being similar to a Lissajous orbit, under a restricted condition where a propulsion force magnitude applied to a spacecraft is fixed, and where it rotates at a constant angular velocity, based on the equation of motion close to a Lagrange point. The method also provide a theory for guiding/controlling the orbit of a spacecraft, that is, the embodiment of the above orbit design method. | 05-06-2010 |
20100163683 | Space Elevator - A freestanding space elevator tower for launching payloads, tourism, observation, scientific research and communications. The space elevator tower has a segmented elevator core structure, each segment being formed of at least one pneumatically pressurized cell. The pressure cells may be filled with air or another gas. Elevator cars may ascend or descend on the outer surface of the elevator core structure or in a shaft on the interior of the elevator core structure. A payload may be launched from a pod or deck at the upper end of the space elevator tower. The space elevator tower is stabilized by gyroscopic and active control machinery. The space elevator tower maintains a desired pressure level through gas compressor machinery. Methods of constructing the space elevator are also disclosed. | 07-01-2010 |
20110272528 | METHOD AND SYSTEM FOR DELIVERING CARGOES INTO SPACE - A method for delivering space cargo including inserting one or more container spacecrafts (CSCs) into orbit. A CSC comprises a hull, a device for receiving cargoes (an artificial medium), a braking medium container, an arrangement for separating the cargoes and the braking medium, storage tanks, a propulsion system, a satellite solar power station, and also heat dissipaters for cooling the braking medium. To compensate for loss of speed during cargo capture, the CSCs use propulsion systems supplied with power from said power station. A reactive-type propulsion system in which a part of the incoming cargo is consumed can be used as such system. | 11-10-2011 |
20120091279 | Space Station Pyramid Structure - The space station pyramid design is to be erected around an Earthbound land mass (i.e. homes, towns and sanctuaries), thereby, encapsulating the land mass. Propulsion systems will raise the station into Earth's orbit. Artificial gravitation is provided through electrostatic forces within the structure. | 04-19-2012 |
20120119034 | Deorbiting a Spacecraft from a Highly Inclined Elliptical Orbit - Deorbiting of an earth-orbiting satellite is accomplished by executing an orbit transfer maneuver, the orbit transfer maneuver resulting in transference of the satellite from an operational orbit to a disposal orbit, where the disposal orbit is substantially circular and has a nominal radius of approximately, 31,000 kilometers. The operational orbit may be substantially geosynchronous and have at least one of (i) an inclination of greater than 10 degrees and (ii) a nominal eccentricity greater than 0.1. Alternatively, the operational orbit may be a medium earth orbit. | 05-17-2012 |
20120187249 | GAS GUN LAUNCHER - A gas gun launcher having a pump tube and a launch tube with a first end of the launch tube slidably inserted into a second end of the pump tube. A sliding seal is employed to retain the gas within the launch tube and the pump tube A launch tube alignment system is preferably automatic, again to enhance the accuracy of launches. And an embodiment of the gas gun launcher suitable for use in water such as an ocean or large lake preferably utilizes a neutrally buoyant launch tube and a neutrally buoyant pump tube. And a fast-closing muffler at the second end of the launch tube conserves the light gas utilized for launching a vehicle. | 07-26-2012 |
20120217348 | EARTH OBSERVATION SATELLITE, SATELLITE SYSTEM, AND LAUNCHING SYSTEM FOR LAUNCHING SATELLITES - Embodiments of an Earth observation satellite comprising a satellite main body having an elongated shape extending in the direction of a roll axis R of the satellite from a first base face of the satellite main body to a second base face of the satellite main body, the satellite main body having a plurality of lateral faces extending from the first base face to the second base face; a solar array; and propulsion means for compensating air-drag being arranged on the base face of the satellite main body. A first solar panel of the solar array is mounted on a first lateral face of the satellite main body, a second lateral face of the satellite main body is configured to radiate heat away from the satellite main body, and a third lateral face of the satellite main body has observation means for Earth observations. | 08-30-2012 |
20130015295 | Traveling wave augmented railgunAANM Campbell; Robert NeilAACI CorralesAAST NMAACO USAAGP Campbell; Robert Neil Corrales NM US - A railgun launcher with principle rail energization and fielding derived from a co-traveling energy pulse associated with a close-coupled parallel transmission line structure. Enhanced system efficiency, plus amelioration of simple railgun negative features, is enabled via the induction hybrid railgun methodology proposed. | 01-17-2013 |
20130187008 | System and Method for Controlling Motion of Spacecrafts - A motion of an object is controlled from a geostationary transit orbit (GTO) of an earth to an orbit of a moon. A first trajectory of the motion of the object is determined from an intermediate orbit of an earth to a neighborhood of a stable manifold of a first Lagrange point (L1). A second trajectory of the motion of the object is determined from the GTO to the intermediate orbit. A third trajectory of the motion of the object is determined from the neighborhood to the stable manifold to an L1 orbit, and a fourth trajectory of the motion of the object is determined from the L1 orbit to the orbit of the moon. A trajectory from the GTO to the orbit of the moon is determined based on a combination of the first, the second, the third, and the fourth trajectories. | 07-25-2013 |
20140110530 | Aerodynamic and Spatial Composite Flight Aircraft, and Related Piloting Method - An aircraft having propulsion units for both conventional aircraft flight in the atmosphere and for high-altitude operation as a rocket. The aircraft is divided into a payload compartment and a compartment containing rocket propulsion unit propellant or fuel, and includes a long transverse wing with a small back-sweep to favour lift in the dense layers of the atmosphere and to thus make it possible to climb to high altitudes at a subsonic speed before using the rocket propulsion units. The return flight portion is performed by gliding or controlled as for a conventional aircraft. | 04-24-2014 |
20140367524 | ORBIT INSERTION DEVICE FOR ARTIFICIAL SATELLITE AIMED TO EXPLORE A PLANET OF UNKNOWN CHARACTERISTICS - An orbit insertion device for artificial satellite aiming to explore a planet with unknown characteristics. A plurality of sensors send a plurality of measured states of the artificial satellite. A planetary attraction constant estimator calculates a planetary attraction constant based on the plurality of measured states. A drag force coefficient estimator calculates a drag force coefficient based on the plurality of measured states. A rate of convergence calculation unit calculates a rate of convergence based on the drag force coefficient. A fictitious control calculation unit calculates a fictitious control based on the planetary attraction constant, the drag force coefficient and the rate of convergence. A normalized thrust calculation unit calculates a normalized thrust based on the fictitious control. A plurality of thrusters generate a thrust based on the normalized thrust. | 12-18-2014 |
20150307213 | GAS GUN LAUNCHER - A gas gun launcher has a pump tube and a launch tube with a first end of the launch tube slidably inserted into a second end of the pump tube. The pump tube may hold a heat exchanger to heat a light gas used to launch a vehicle. A sliding seal can be employed to manage recoil and to retain the gas within the launch tube and the pump tube. A fast-closing muffler at the second end of the launch tube can conserve the light gas utilized for launching a vehicle, enabling the light gas to be recycled. A launch tube alignment system is preferably automatic, ensuring the survival of the launch vehicle. | 10-29-2015 |
20150329222 | CONFIGURABLE SPACE STATION MOMENTUM - A method of deploying a modular space station comprises placing an initial space station module in space in a first deployment, the initial space station module including a first control law and momentum component that provides an initial solution for guidance, navigation, and control (GNC) during the first deployment. A first space station modular segment is joined with the initial space station module in a second deployment to produce a first joint configuration of the space station. A second control law and momentum component provides a second solution for GNC of the first joint configuration during the second deployment. A second space station modular segment is joined to the first joint configuration in a third deployment to produce a second joint configuration of the space station. A third control law and momentum component provides a third solution for GNC of the second joint configuration during the third deployment. | 11-19-2015 |
20150353208 | HIGHLY INCLINED ELLIPTICAL ORBIT LAUNCH AND ORBIT ACQUISITION TECHNIQUES - Techniques for placing a satellite into a highly inclined elliptical operational orbit having an argument of perigee of 90° or 270° include executing an orbit transfer strategy that transfers the satellite from a launch vehicle deployment orbit to the operational orbit. The launch vehicle deployment orbit is selected to have an argument of perigee of approximately 90° greater than the argument of perigee of the operational orbit, an apogee altitude of approximately 14000 km and a perigee altitude of approximately 500 km. The orbit transfer strategy includes (i) an apsidal rotation of approximately 90°, at least a substantial part of the apsidal rotation being attained without expenditure of any satellite propellant; and (ii) an electric orbit raising maneuver to attain an apogee altitude and a perigee altitude required by the HIEO. | 12-10-2015 |
20160009425 | Systems and Methods for Optimizing Satellite Constellation Deployment | 01-14-2016 |
20160023780 | METHOD FOR STATIONING A SATELLITE AND IN-ORBIT TESTING OF ITS PAYLOAD - A method for stationing a satellite comprises a transfer phase, during which the satellite moves on an elliptical geosynchronous orbit, the orbit being deformed progressively by application of a thrust by electrical or hybrid electrical-chemical propulsion to bring it closer to a geostationary orbit. The transfer step comprises a substep during which, during a plurality of revolutions of the satellite, the thrust is stopped for a fraction of orbital period and tests of a telecommunications payload of the satellite are performed in the absence of thrust. | 01-28-2016 |
20160090197 | KINETIC ENERGY STORAGE AND TRANSFER (KEST) SPACE LAUNCH SYSTEM - A Kinetic Energy Storage and Transfer (KEST) vehicle and target vehicle kinetic energy transfer method are provided. The KEST vehicle is configured to transfer kinetic energy to the target vehicle, propelling the target vehicle into a higher orbit or beyond the Earth. This is accomplished by a catching mechanism that contacts the target vehicle. The catching mechanism may also include a braking mechanism configured to accelerate the target vehicle, and thus slow the KEST vehicle, as the catching mechanism and target vehicle travel along one or more tethers of the KEST vehicle. Alternatively, the catching mechanism may be attached to an end of the one or more tethers and be configured to slow the target vehicle as the one or more tethers bend. | 03-31-2016 |
20160114909 | SPACE CRAFT COMPRISING AT LEAST ONE PAIR OF SUPPORTING ARMS, SAID ARM PAIR BEING EQUIPPED WITH A HOLLOW MOUNTING MODULE, AND METHOD FOR EMPLOYING SUCH A CRAFT - Space craft comprising a body, at least one pair of supporting arms, a first device mounted on a first supporting arm and a second device mounted on a second supporting arm. The first arm is rotatably mounted on the body of the craft about an axis of rotation. The second arm is fixed to the body, and in which craft of the first device and the second device at least one is offset from the axis of rotation of the first arm. The pair of supporting arms further comprises a hollow module for the rotatable mounting of the first arm on the body. The mounting module comprising an opening through which the axis of rotation and the second supporting arm pass. | 04-28-2016 |
244158600 | Orbital control | 41 |
20090166476 | Thruster system - 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. | 07-02-2009 |
20090321579 | Space based orbital kinetic energy weapon system - 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. | 12-31-2009 |
20110036951 | HYBRID ELECTROSTATIC SPACE TUG - 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. | 02-17-2011 |
20110036952 | Electrostatic Spacecraft Reorbiter - 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. | 02-17-2011 |
20110260006 | Nejat kinematics and kinetics expansions spacecraft - 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. | 10-27-2011 |
20120018585 | MEDIUM EARTH ORBIT CONSTELLATION WITH SIMPLE SATELLITE NETWORK TOPOLOGY - A method, a member satellite, and a tangible machine-readable medium are disclosed. An inter-satellite link subsystem | 01-26-2012 |
20120018586 | Method and Device for Optimization of the Mass of a Satellite - 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. | 01-26-2012 |
20120085869 | APPARATUS, METHOD AND SYSTEM FOR REMOVING ORBITAL DEBRIS - 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. | 04-12-2012 |
20120097796 | Satellite orbit raising using electric propulsion - Apparatus and methods for raising the orbit of a satellite having electric propulsion thrusters, an Earth sensor and an inertial reference sensor such as a gyro. A satellite positioning system generates orbital data and a profile generator generates an ideal electric orbit raising profile of the satellite. The ideal profile is one that the satellite must follow so that the perigee, apogee and inclination of the satellite can be adjusted simultaneously in a mass-efficient manner. A state machine processes the ideal profile and a true anomaly to generate a desired electric orbit raising profile. Steering apparatus generates signals that are used to control the attitude of the satellite to follow the desired profile. The desired profile places the satellite in an Earth-pointed attitude when the satellite is at a predefined point in the orbit, slews the satellite from the Earth-pointed attitude to an ideal orbit raising attitude, steers the satellite according to the ideal profile during orbit raising, and steers the satellite from the desired attitude to the Earth-pointed attitude. | 04-26-2012 |
20120097797 | SPACECRAFT MOMENTUM MANAGEMENT - Three-axis spacecraft momentum management is performed for a spacecraft traveling along a trajectory, by an actuator including at least one thruster disposed on a single positioning mechanism. As the spacecraft travels along the trajectory, a desired line of thrust undergoes a substantial rotation in inertial space. When the spacecraft is located at a first location on the trajectory, the single positioning mechanism orients the thruster so as to produce a first torque to manage stored momentum in at least one of a first and a second of the three inertial spacecraft axes. When the spacecraft is located at a second location on the trajectory, the single positioning mechanism orients the thruster so as to produce a second torque to manage stored momentum in at least a third of the three inertial spacecraft axes. | 04-26-2012 |
20120138748 | TERMINATOR TAPE SATELLITE DEORBIT MODULE - A deorbit device comprising a passive electrodynamic conductive tape connected at one end to a spacecraft. | 06-07-2012 |
20130001365 | ORBITAL DEBRIS MITIGATION USING HIGH DENSITY PLASMA - The present invention is directed to mitigating orbital debris in outer space by concentrating the plasma existing in space. The debris mitigation method involves sending a satellite device | 01-03-2013 |
20130062471 | INCLINED ORBIT SATELLITE COMMUNICATION SYSTEM - A method of flying a constellation of inclined geosynchronous satellites at the same station longitude with specific spacing but without the possibility of collision and provides the basic equations defining the initial positions of satellites such that the satellites will continue to remain in synchronized positions relative to each other for a number of years with little or no north-south positioning. In preferred embodiments the number of satellites in the constellation is five or ten. Communication with the satellites in the constellation is provided with existing prior art tracking radio systems. | 03-14-2013 |
20130105632 | Method and System for Controlling a Set of at Least Two Satellites Adapted to Provide a Service | 05-02-2013 |
20130292516 | Propulsion System for Satellite Orbit Control and Attitude Control - A propulsion system for the orbit control of a satellite with terrestrial orbit having an angular momentum accumulation capacity comprises a propulsion unit able to deliver a force along an axis F having a component perpendicular to the orbit, and a motorized mechanism connected on the one hand to the propulsion unit and on the other hand to a structure of the satellite, the motorized mechanism being able to displace the propulsion unit along an axis V parallel to the velocity of the satellite, and able to orient the propulsion unit so as to make it possible to control: a component of the force along the axis V, for orbit control, an amplitude and a direction of couple in a plane perpendicular to the axis F, for control of the angular momentum. | 11-07-2013 |
20130306799 | SPACE DEBRIS REMOVAL USING UPPER ATMOSPHERE - The systems and methods of the invention modulate atmospheric gases to temporarily increase the amount of atmospheric particles in the path of the debris, in order to decelerate the debris and accelerate natural orbital decay to the point of atmospheric re-entry. In one aspect of the invention, clearing the space debris includes propelling a plume of atmospheric gases substantially orthogonal to the path of the debris such that the debris collides with the gaseous plume as it passes through the plume. Increased atmospheric drag from the gaseous particles of the plume in the path of the debris obstructs a forward propagation of the debris and gradually decelerates the debris, leading eventually to atmospheric recapture. Embodiments of the invention can be employed in any number of applications, including without limitation, clearing debris in the low-earth orbit (LEO) which is particularly susceptible to debris build-up, de-orbiting non-refuse payloads front orbits, and clearing debris from geosynchronous orbits. | 11-21-2013 |
20130313369 | PROPULSION SYSTEM FOR CONTROLLING THE ORBIT AND CONTROLLING THE ATTITUDE OF A SATELLITE - A propulsion system for controlling the orbit of a satellite in earth orbit comprises a thruster suitable for delivering a force along an axis F, and a motor-driven mechanism linked on the one hand to the thruster and on the other hand to a structure of the satellite, said motor-driven mechanism being suitable for displacing the thruster on either side of the plane of the orbit and suitable for orienting the thruster so as to make it possible to control a component perpendicular to the orbit of the force in two opposite directions, to control the inclination of the satellite, and in that said motor-driven mechanism is suitable for displacing the thruster along an axis V parallel to the velocity of the satellite, and suitable for orienting the thruster so as to make it possible to control a component of the force on the axis V, to control orbit. | 11-28-2013 |
20130327893 | APPARATUS AND METHOD FOR CONTROLLING GEOSTATIONARY ORBIT SATELLITE - An apparatus and method for controlling a geostationary orbit satellite is provided. The method including generating remote measurement data by measuring a state of a geostationary orbit satellite, transmitting the remote measurement data, receiving a remote command signal, and controlling an orbit and a pose of the geostationary orbit satellite relative to inclined geosynchronous space debris. | 12-12-2013 |
20140027576 | Earth Observation Constellation Methodology & Applications - The focus of this invention pertains to the methodology behind launching line-scanning satellite constellations that can image an entire planet such as the Earth at high temporal cadence (less than a week), at high spatial resolution (less than 10 m). Utilizing simple control and operation, our invention captures images of an entire planet in an effective and distributed manner. Additional benefits are realized by taking advantage of the distributed onboard storage and computing abilities of such a constellation to optimize the data collected, system latency, and data downlinked. | 01-30-2014 |
20140034784 | DEVICE AND METHOD FOR DEORBITING OF A SATELLITE - The present disclosure relates to the deorbiting of satellites in low orbit that have entered safe hold mode. A device makes it possible to decide in an autonomous manner and on the basis of information existing in the satellite, when and where to trigger a series of short thruster manoeuvres to modify the satellite orbit with the aim of deorbiting. | 02-06-2014 |
20140158830 | DEVICE FOR MOVING OR REMOVING ARTIFICIAL SATELLITES - A device for coupling with a space satellite before the satellite is launched for the purpose of de-orbiting said satellite and/or returning it to Earth. The device includes:
| 06-12-2014 |
20140339368 | ECCENTRICITY CONTROL FOR GEOSYNCHRONOUS SATELLITES - Eccentricity control for a geosynchronous satellite includes: setting initial conditions, duration, and schedule for the eccentricity control; defining a plurality of parameters including control loci for centroid, semi-major axis, semi-minor axis, uncontrolled eccentricity radius, right ascension of ascending node, and inclination, wherein the plurality of parameters are defined such that when the eccentricity control is applied, a mean geodetic longitude of the geosynchronous satellite is maintained within a predefined distance from a station longitude. | 11-20-2014 |
20140361123 | PROPULSION SYSTEM IN TWO MODULES FOR SATELLITE ORBIT CONTROL AND ATTITUDE CONTROL - A propulsion system for the orbit control of a satellite in Earth orbit driven at a rate of displacement along an axis V tangential to the orbit comprises two propulsion modules, fixed to the satellite, and facing one another relative to the plane of the orbit, each of the propulsion modules comprising, in succession: a motorized rotation link about an axis parallel to the axis V; an offset arm; and a plate supporting two thrusters, suitable for delivering a thrust on an axis, arranged on the plate on either side of a plane P at right angles to the axis V passing through a centre of mass of the satellite; each of the two thrusters being oriented in such a way that the thrust axes of the two thrusters are parallel to one another and at right angles to the axis V. | 12-11-2014 |
20140361124 | PROPULSION SYSTEM WITH FOUR MODULES FOR SATELLITE ORBIT CONTROL AND ATTITUDE CONTROL - A propulsion system for the orbital control of a satellite with terrestrial orbit travelling with a speed of displacement along an axis V tangential to the orbit comprises two propulsion assemblies, fixed to the satellite facing one another with respect to the plane of the orbit, each of the propulsion assemblies comprising two propulsion modules; each of the propulsion modules successively comprising: a motorized link for rotation about an axis parallel to the axis V, an offset arm, and a platen supporting a propulsion unit able to deliver a thrust oriented along an axis perpendicular to the axis V; the two propulsion modules of each propulsion assembly being linked to the satellite on either side and substantially at equal distances from a plane perpendicular to the axis V passing through a centre of mass of the satellite. | 12-11-2014 |
20150083865 | MULTIPLE SPACECRAFT LAUNCH SYSTEM - A system and method for propelling spacecraft is disclosed. An electrical propulsion system is mounted on a base stage. A plurality of spacecraft couplers are also mounted on the base stage. Each spacecraft coupler securedly attaches a spacecraft to the base stage. Each spacecraft includes an internal power source that is coupled to the electrical propulsion system via an electrical connection. The internal power source consists of solar panels and/or batteries. A power regulation circuit is coupled between the electrical propulsion system and each internal power source. The power regulation circuit is draws an equal and proportional amount of power from each spacecraft. The spacecraft are preferably satellites and the electrical propulsion system preferably propels the base stage and attached satellites from a lower-Earth orbit to a higher-Earth orbit so that the electrical propulsion system in each satellite need only be capable of providing propulsion for orbit maintenance and maneuvering. | 03-26-2015 |
20150144738 | SPACECRAFT FITTED WITH A DE-ORBITING DEVICE COMPRISING A DETONATION ENGINE - A spacecraft including at least one main propellant tank, a main engine fed with propellant by the main tank, and a de-orbiting device. The de-orbiting device includes a detonation engine fed with propellant by the main tank. | 05-28-2015 |
20150353209 | HIGHLY INCLINED ELLIPTICAL ORBIT DE-ORBIT TECHNIQUES - Techniques for deorbiting a satellite include executing an orbit transfer maneuver that transfers the satellite from an operational orbit to an interim orbit. The operational orbit is substantially geosynchronous and has (i) an inclination of greater than 70 degrees; (ii) a nominal eccentricity in the range of 0.25 to 0.5; (iii) an argument of perigee of approximately 90 or approximately 270 degrees; (iv) a right ascension of ascending node of approximately 0; and (v) an operational orbit apogee altitude. The interim orbit has an initial second apogee altitude that is at least 4500 km higher than the first apogee altitude, and the interim orbit naturally decays, subsequent to the orbit transfer maneuver, such that the satellite will reenter Earth's atmosphere no longer than 25 years after completion of the orbit transfer maneuver. | 12-10-2015 |
20160137317 | Satellite System and Method for Global Coverage - The present invention relates to satellite systems and more particularly, to the provision of a novel, non-geostationary satellite system and method for weather and climate monitoring, communications applications, scientific research and similar tasks, with global coverage. Contrary to the teachings in the art it has been discovered that global coverage may be obtained using a constellation of six satellites in two orthogonal, 24 sidereal hour orbits (geosynchronous) with inclinations of 70° to 90°, and eccentricities of 0.275-0.45. By placing three of the satellites in a first orbit with an apogee over the north pole, and three of the satellites in a second, orthogonal orbit with an apogee over the south pole, global coverage may be obtained. As well, the satellites in these orbits avoid most of the Van Allen Belts. | 05-19-2016 |
20160176545 | THRUSTER SUPPORT MECHANISM FOR SATELLITE PROPULSION | 06-23-2016 |
20160376033 | EFFICIENT STATIONKEEPING DESIGN FOR MIXED FUEL SYSTEMS - Apparatus and methods for stationkeeping in a satellite. The satellite includes a north electric thruster and a south electric installed on a zenith side. An orbit controller selects a duration of a burn of the north electric thruster proximate to an ascending node that differs from a duration of a burn of the south electric thruster proximate to a descending node. The orbit controller is configured to select an offset of the burn of the north electric thruster in relation to the ascending node that differs from an offset of the burn of the south electric thruster in relation to the descending node. | 12-29-2016 |
244158700 | Aerobraking | 4 |
20090218448 | SATELLITE AIR BRAKE WING STRUCTURE - A deployable wing structure for air-braking a satellite and which, once deployed, at least one wing structure element that forms a three-dimensional structure and includes at least two panels lying in secant planes and forming a dihedron. | 09-03-2009 |
20120175466 | SPACE DEBRIS REMOVAL USING UPPER ATMOSPHERE - The systems and methods of the invention modulate atmospheric gases to temporarily increase the amount of atmospheric particles in the path of the debris, in order to decelerate the debris and accelerate natural orbital decay to the point of atmospheric re-entry. In one aspect of the invention, clearing the space debris includes propelling a plume of atmospheric gases substantially orthogonal to the path of the debris such that the debris collides with the gaseous plume as it passes through the plume. Increased atmospheric drag from the gaseous particles of the plume in the path of the debris obstructs a forward propagation of the debris and gradually decelerates the debris, leading eventually to atmospheric recapture. Embodiments of the invention can be employed in any number of applications, including without limitation, clearing debris in the low-earth orbit (LEO) which is particularly susceptible to debris build-up, de-orbiting non-refuse payloads from orbits, and clearing debris from geosynchronous orbits. | 07-12-2012 |
20130082146 | SYSTEM AND METHOD FOR CREATING AN ARTIFICIAL ATMOSPHERE FOR THE REMOVAL OF SPACE DEBRIS - Systems and methods are disclosed herein for producing an artificial atmosphere for creating drag on space debris. The artificial atmosphere is produced by an artificial atmosphere delivery system (AADS), which comprises a combustible propellant and a propellant ignition device. The combustible propellant creates an artificial atmosphere when ignited. The combustible propellant is directed to a region near the path of the space debris by a targeting device. The propellant ignition device ignites the combustible propellant such that ignition of the combustible propellant does not impart significant momentum on the AADS. Igniting the combustible propellant creates the artificial atmosphere in the path of the space debris. | 04-04-2013 |
20180022476 | SATELLITE DEORBITING SYSTEM | 01-25-2018 |
244158800 | Automatic | 7 |
20120181386 | Eccentricity vector control with continuous or quasi-continuous maneuvers - A method of managing an eccentricity vector of a geosynchronous satellite orbit is provided. The method includes determining a desired target locus of an acquisition control of a satellite in a geosynchronous orbit, where the acquisition control ensures that an osculating trajectory of the satellite converges in mean to the target locus. Further, the method includes determining a solar pressure perturbation to the geosynchronous satellite orbit, honoring a hard eccentricity limit constraint of the satellite orbit using an ideal continuously controlled osculating trajectory, and controlling the eccentricity vector of the geosynchronous satellite orbit using a quasi-continuous control or a continuous control to mitigate or eliminate an annual solar pressure perturbation, where the quasi-continuous control or the continuous control maintains the satellite orbit within the hard limit osculating constraint and converges the eccentricity of the satellite orbit toward the ideal continuously controlled osculating trajectory of the geosynchronous satellite orbit. | 07-19-2012 |
20120248253 | MULTI-BODY DYNAMICS METHOD OF GENERATING FUEL EFFICIENT TRANSFER ORBITS FOR SPACECRAFT - A method of generating orbital transfers for spacecraft. The method provides an innovative technique for transferring spacecraft from one Earth orbit to another Earth orbit using significant solar gravitational influences. In one particular implementation, the multi-bodies in the transfer determination are the Earth (about which the spacecraft is to orbit) and the Sun (e.g., the Earth and the Sun are the first and second celestial bodies providing multi-body dynamics). The transfer orbit or trajectory is determined to make use of efficient tangential maneuvers by leveraging solar gravitational influences to improve transfer performance. Based on the generated transfer orbit, the spacecraft is controlled to perform one or more maneuvers to achieve a transfer orbit that traverses into a regime where the spacecraft's trajectory is significantly affected by gravity from both the Sun and the Earth. The spacecraft performs a near-tangential orbit insertion maneuver to enter the final orbit. | 10-04-2012 |
20130292517 | Autonomous satellite orbital debris avoidance system and method - An autonomous system for a satellite which calculates collision paths of debris from anywhere within the spheroid around the satellite by using its radar/ladar data and from data on its own orbit derived by onboard sensors such as star, earth and sun sensors or from stored data sent from its ground control station through the satellite's command subsystem. If a collision would be likely, the system calculates the minimum change in the satellite's orbit to avoid such collision and generates and executes commands for firing on-board orbital control thrusters to put the satellite in a suitable avoidance orbit. | 11-07-2013 |
20140077036 | System and Method for Maneuver Plan for Satellites Flying in Proximity - A technique to assist guidance techniques for a free-flying inspection vehicle for inspecting a host satellite. The method solves analytically in closed form for relative motion about a circular primary for solutions that are non-drifting, i.e., the orbital periods of the two vehicles are equal, computes the impulsive maneuvers in the primary radial and cross-track directions, and parameterizes these maneuvers and obtain solutions that satisfy constraints, for example collision avoidance or direction of coverage, or optimize quantities, such as time or fuel usage. Apocentral coordinates and a set of four relative orbital parameters are used. The method separates the change in relative velocity (maneuvers) into radial and crosstrack components and uses a waypoint technique to plan the maneuvers. | 03-20-2014 |
20140166814 | Method and system for sationing a satellite - A method for stationing a satellite, comprises: determining a predefined trajectory for stationing of the satellite based on a model of movement of the satellite; determining parameters of the predefined control law of approximation of the trajectory and by minimizing impact on the control law of deviation from a trajectory followed by the satellite using parameters of the control law; determining a state vector of the satellite; determining a deviation between the state vector and the predefined trajectory; determining Lagrange multipliers based on a current state vector of the satellite, on a deviation between current state vector and predefined trajectory and on parameters of the predefined control law; determining parameters of the current control law of the engines based on the Lagrange multipliers and by derivation of a parameter representative of an effect of engines on the real trajectory; and controlling engines based on parameters of the current control law. | 06-19-2014 |
20150041595 | Attitude and orbit control system and method for operating same - A hybrid network of kinematic sensors of an AOCS, made up of a star sensor including an optical camera head, and a processing unit provided as the central master processing unit, and additional kinematic sensors, each made up of a sensor element and a processing unit connected to the central processing unit via a first bus. An additional processing unit is equivalent to the processing unit and is a redundant central processing unit. The central processing units and—are connected via an additional bus of a spacecraft provided with the hybrid network with the aid of a central computer. The particular active central processing units-provide all kinematic sensors with a uniform time pulse via a synchronization line, and supply the central computer with hybridized kinematic measuring data formed according to a method for hybridization based on the synchronous kinematic measuring data of the star sensor and the measuring data of the other sensors. | 02-12-2015 |
20160376035 | EFFICIENT STATIONKEEPING DESIGN FOR MIXED FUEL SYSTEMS IN RESPONSE TO A FAILURE OF AN ELECTRIC THRUSTER - Apparatus and methods for stationkeeping in a satellite. The satellite includes a north electric thruster and a south electric installed on a zenith side, an east chemical thruster installed on an east side, and a west chemical thruster installed on a west side. An orbit controller detects a failure of one of the electric thrusters. In response to the failure, the orbit controller controls a burn of the remaining electric thruster proximate to an orbital node. The orbit controller controls a burn of one of the chemical thrusters at 90°±5° from the burn of the remaining electric thruster, and controls a burn of the other one of the chemical thrusters at 270°±5° from the burn of the remaining electric thruster. | 12-29-2016 |