The PanelSat serves to launch one, better several satellites into space, whereby besides unfurling of the thin film solar cell panels off their rolls no further deployment is needed. PanelSats are small agile spacecraft thought especially for observation and communication services in LEO, which are using their thin film solar cell panels for both, harvesting electric energy as well as for fuel less station keeping, steering, pointing and propulsion. In contrast to conventional satellites with their 3-axis control design, PanelSats are not locked to only 3 axles and can tilt and point into several directions (depending on the number of panels). Besides Roller Reefing for fuel less attitude control PanelSats feature “Soso Steering” (switch on, switch off) which adds even better fuel less agility compared to prior art satellites.

Provided is a configuration construction and attitude control method for a pyramid deorbit sail. By taking into consideration environmental perturbation like atmospheric resistance perturbation and non-spherical earth perturbation, a dynamics model featuring three-dimensional orbit-and-attitude coupling based on position vectors and quaternion descriptions, the deorbit sail is taken as a rigid body, a spacecraft body is taken as a mass point, airflow obstruction is considered in the windward area, thereby improving the precision of the dynamics model; based on this model, the law of influence of the configuration parameters in the deorbit sail, such as a cone angle and a strut length, on the attitude stability and deorbiting efficiency of the spacecraft in different cases is analyzed, the configuration parameters of the pyramid deorbit sail system are analyzed and optimized according to the derived law, so as to obtain a pyramid deorbit sail achieving high attitude stability and high deorbiting efficiency.

In a method for packing a spacecraft membrane (1) which in a plane of extension comprises a longitudinal axis between opposite longitudinal corners (4) and a transverse axis running transverse to the longitudinal axis and through transverse corners into a spacecraft membrane package, two packing steps are executed: In a first packing step, the spacecraft membrane (1) is packed into a transverse package (9) along the transverse axis. In a second packing step, the transverse package (9) is packed into a longitudinal package along a longitudinal axis. The packing of the spacecraft membrane (1) in the first packing step comprises a packing of material of the spacecraft membrane (1) from or on both sides of the longitudinal axis. In the first packing step, the spacecraft membrane (1) is packed in such a way that in the created transverse package (9) the transverse corners are freely accessible. In the second packing step, the transverse package (9) is packed in such a way that in the created longitudinal package the longitudinal corners (4) are freely accessible. The longitudinal package is unpacked by pulling on the longitudinal corners (4). Subsequently, the transverse package (9) is unpacked by pulling on the transverse corners.

A space vehicle may extract and store energy, and also include a propulsion system. The space vehicle includes one or more wings connected to a body of the apparatus. Each of the one or more wings includes a propellant system configured to eject mass away from the apparatus, the ejection of the mass causes the apparatus to move from a first position to a second position.

Techniques for performing spacecraft rendezvous and/or docking include operating a first orbiting spacecraft, the first spacecraft including a sensor arrangement, a first processor and a first inter-satellite link (ISL) arrangement and performing one or both of a rendezvous operation and a docking operation with the first spacecraft and a second orbiting spacecraft, the second spacecraft including one or more actuators. The performing one or both of the rendezvous operation and the docking operation includes determining a pose and pose rate of the second spacecraft relative to the first spacecraft using observations made by the sensor arrangement and determining a desired approach trajectory for the second spacecraft.

The present invention generally is a system and method for ground, atmospheric and space based solar powered electrical energy generation and transmission of beamed microwave power. Specifically it is a system and method for generating electrical energy from a plurality of photovoltaic cells dispersed on a flexible surface each in close proximity to and functionally connected to microwave generating and transmitting means for controllably forming one or more stronger microwave beams by combining a plurality of much weaker individual microwave beams. The invention can be a microwave beam weapon for detecting and transferring microwave energy to non-cooperative targets or it clear orbital debris by momentum transfer to space object through microwave radiation pressure. Most practically it can provide electric power and microwave beam weapon defense to remote military and civilian facilities, including forward operating bases.

Exemplary embodiments described herein include innovative engagement devices. Exemplary engagement devices may include on or more tape spring systems. The tape spring system may include a continuous or segmented bi-stable tape spring. The tape spring can be stowed in a rolled up configuration, extended to a deployed configuration, and then triggered to return to a retracted configuration.

A method and apparatus for the control of the attitude of earth orbiting satellites and the orbit and attitude control of a novel gravitational wave detection satellite configuration located near the sun-earth Lagrangian points L3, L4 and L5, utilizing the control of solar radiation pressure by the use of electrically controllable variable reflection glass panels to provide the torques and forces needed.

This embodiment relates to a lifting or flying device using the energy from the radioactive decay of radioactive elements to accelerate an object. A thin radioactive coating is spread over a large surface area that allows most of the radiated particles to escape. The force from the decay of an unobstructed side A of a flattened or curved radioactive emitting device has enough energy to push an object toward its opposite side B expelling particles or waves at relativistic speeds in its exhaust if side B is covered by a shield that prevents most of the radiation from escaping that side. Trillions of small microscopic explosions per second per gram of radioactive material has enough energy from radioactive decay from alpha, beta, or gamma rays decaying to escape on A side which is much greater number of particles escaping than shielded side B imparting a force in in B direction.

An assembly for deployment and retraction of a panel for outer space environment usage, includes a member secured to the panel with the member having an axis of rotation. Mass component connected to the panel, with panel wrapped about member with mass component positioned a first radial distance from the axis of rotation. Mass component has a mass per unit area greater than a mass per unit area of the panel adjacent to the mass component. With a rate of rotation of the mass component and member being same, the mass component rotates about the axis of rotation at a second radial distance from the axis of rotation, wherein the second radial distance is greater in dimension than the first radial distance. With a rate of rotation of the mass component and the member changed to be different, radial distance of the mass component to the axis of rotation decreases.