A system and method of providing an affordable navigation, guidance and control system for arbitrary nano/micro launch vehicles by integrating commercial grade sensors with advanced estimation algorithms in a manner that provides sufficient accuracy of the resulting vehicle state estimates to inject nano/micro satellites into low earth orbits. The system and method uses commercial grade sensors and an advanced sensor-fusion estimator software that estimates and removes the estimated measurement errors and filters noise produced by the commercial grade sensors, resulting in estimated states with suitable accuracy. The filtered data are sent to a guidance and control system where actuator commands are formulated based on the filtered data. A simulated launch and flight of the launch vehicle is performed using the filtered data to validate that the GNC system and launch vehicle are ready for launch.

A vehicle, such as a satellite or other spacecraft, includes an electrodynamic tether connected thereto. A processor, contained within the vehicle and connected to the electrodynamic tether, is configured to cause a current to be directed to the electrodynamic tether to cause a change in motion of the vehicle. Sensors, such as an attitude sensor, a position sensor, a magnetometer, and an ionosphere sensor, are contained within the vehicle and are connected to the processor. The processor is configured to direct current to the electrodynamic tether based upon input received from the sensors to maintain the vehicle within a specified orbit, such as a highly-inclined eccentric orbit over the polar or other high-latitude region, or to change the vehicle's orbit. The processor may be configured in a closed-loop configuration to account for measured errors by the sensors position, attitude, ionospheric charge density, and/or the Earth's magnetic field.

A Hall sensor may include a Hall sensing element configured to produce a Hall voltage indicative of a magnetic field when traversed by an electric current, and a first pair of bias electrodes mutually opposed in a first direction across the Hall sensing element. The Hall sensor may include a second pair of bias electrodes mutually opposed in a second direction across the Hall sensing element. The Hall sensor may include a first pair of sensing electrodes mutually opposed in a third direction across the Hall sensing element, and a second pair of sensing electrodes mutually opposed in a fourth direction across the Hall sensing element. The fourth direction may be orthogonal to the third direction, each sensing electrode being between a bias electrode of the first pair and a bias electrode of the second pair.

A magnetic field sensing apparatus including a magnetic flux concentrator, a plurality of magnetoresistance units, and a plurality of magnetization direction setting elements is provided. The magnetic flux concentrator has a top surface, a bottom surface opposite to the top surface, and a plurality of side surfaces connecting the top surface and the bottom surface. The magnetoresistance units are respectively disposed beside the side surfaces. The magnetoresistance units are electrically connected to form an unchangeable Wheatstone full bridge. The magnetization direction setting elements set the magnetization directions of the magnetoresistance units into three different combinations in three different periods, respectively, so as to enable the unchangeable Wheatstone full bridge to respectively measure the magnetic field components in the three different directions in the three different periods.

The invention relates to small satellites capable to fly in formation (10), in particular nano- or picosatellites with a mass of 10 kg or less, for LEO applications, comprising a housing (12) and at least one plug-in board (14) arranged in the housing (12) with a predetermined functionality and a propulsion system (16) for generating a directed pulse in the direction of the flight trajectory T.sub.k.

It is proposed that the small satellite (10) comprises an independent and autonomously working collision avoidance system (18), which is capable of adapting a trajectory correction T.sub.kk of the trajectory T.sub.k by the propulsion system (16), when a collision with an object (30) is expected.

In a further independent aspect, the invention relates to a formation (100) composed of several small satellites capable to fly in formation (10), wherein a relative position and flight trajectory T.sub.k of each small satellite (10) is modifiable via the independently and autonomously working collision avoidance system (18).

A system utilizing an antenna generating an electromagnetic (EM) wave to interact with a solar EM wave to streamline magnetic flux in the polar cusp and to facilitate the flow of solar plasma through the Polar Cusp, resulting in an elevated plasma flux at the exit of the Polar Cusp. The elevated plasma flux intercepts and removes small space debris from Low Earth Orbit (LEO), Geosynchronous Earth Orbit (GEO) and Geosynchronous Transfer Orbits (GTO) transiting the LEO altitude regimes.

Feedback control circuitry includes rate limiter circuitry configured to generate a rate limited position command based on a position command for a controlled component and based on a speed command for the controlled component. The feedback control circuitry also includes error adjustment circuitry configured to apply a control gain to an error signal to generate an adjusted error signal. The error signal is based on position feedback and the rate limited position command, and the position feedback indicates a position of the controlled component. The feedback control circuitry further includes an output terminal configured to output a current command generated based on the adjusted error signal.

This disclosure describes techniques for providing a transformative framework to forecast physical properties of an atmosphere to predict the orbit of satellite devices. As one example, the transformative framework has two major components: (i) the development of a quasi-physical dynamic reduced-order model (ROM) that uses a linear approximation of the underlying dynamics (e.g., solar conditions or magnetic conditions) and effect of the drivers, and (ii) data assimilation and calibration of the ROM through estimation of the ROM coefficients that represent the model parameters.

A burn wire release mechanism for a spacecraft or other system having two masses initially held together by a pretensioned loop, with a spring configured to push the masses apart. The burn wire release system includes at least one burn wire held in contact with the pretensioned loop material. When an electrical current flows through the burn wire, the wire heats up and severs the pretensioned loop, and the masses are pushed apart by the spring.

A tethered spacecraft has a first endmass and a second endmass with a telescoping stacer spring and a tether arranged between the endmasses. The spring is coiled around a center rod and initially contained within a housing, the spring being biased to push the first endmass away from the second endmass. The spring housing is affixed to the first endmass, a first end of the spring being affixed to the spring housing, and tether are affixed to spring at one end and to the second endmass at the other end. A pretensioned loop holds the endmasses abuttingly together, and a burnwire release mechanism cuts the loop to deploy the spring. Upon deployment, the spring extends to its full length to form a cylindrical boom, and the endmasses continue to move outward along the spring centerline until stopped by the tether.