Abstract: 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.

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 method of launching space vehicles by towing them aloft, then twirling them around a large transport aircraft (40) at the center of a formation (AA) of other tow aircraft (28, 34) and other devices of the invention. A lengthy, semi-rigid tow pipeline (14) serves as a conduit for the transfer of fuels and oxidizers, as the tow cable, and as an energy storage device that reflexes efficiently when it is flexed. The flexing of tow pipeline (14) is caused by a parachute (22) acting in conjunction with all the aircraft making the tighest turn they are capable of doing. Tow aircraft in certain arrays (28) are joined to tow pipeline (14) by sliding trollies (26) that also host canard rotor wings for the aerodynamic support of the main tube (12). The tow trollies (26) aid the sliding tow aircraft arrays (28) in gaining mechanical advantage to accelerate the space vehicle.

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: 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 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 practical orbit raising method and system wherein a satellite quickly escapes the Van Allen radiation belts and payload mass and mission life are maximized. A satellite is launched that contains high thrust chemical propulsion thrusters, high specific impulse electric propulsion thrusters and a solar array. The satellite quickly escapes the Van Allen radiation belts by firing the high thrust chemical propulsion thrusters at apogees of intermediate orbits, starting from the transfer orbit initiated by a launch vehicle, to successively raise the perigees until the perigee clears the Van Allen radiation belts. The payload mass and mission life are maximized by firing high specific impulse electric propulsion thrusters to raise the satellite to near synchronous orbit, while steering the thrust vector and solar array to maintain the sun's illumination on the solar array. The chemical and/or electric propulsion thrusters are then fired to achieve geosynchronous orbit.

Abstract: A system for accelerating an object that includes a towing vehicle, a tether connected to the towing vehicle, an object having at least one airfoil connected to the tether and a controller for controlling movement of the object, wherein, the towing vehicle moves in a first direction at a first speed pulling the object in the first direction and the controller controls movement of the object to move the object at an angle to the first direction and accelerate the object to a speed substantially greater than the first speed.