Provided is an apparatus for controlling an orbiting satellite by sensing a change in a yaw angle of the orbiting satellite and calculating a ground sample distance (GSD) based on the yaw angle. The apparatus may include a sensor configured to sense a yaw angle corresponding to yaw steering of the orbiting satellite, and a processor configured to calculate, based on the yaw angle, a GSD corresponding to a length of a pixel projected onto a planetary surface scanned by the orbiting satellite.

A system and method for assisted EVA self-return is provided herein. The system estimates a crewmember's navigation state relative to a fixed location, for example on an accompanying orbiting spacecraft, and computes a guidance trajectory for returning the crewmember to that fixed location. The system may account for safety and clearance requirements while computing the guidance trajectory. According to at least one embodiment, the system actuates the crewmember's safety jetpack to follow the prescribed trajectory to the fixed location. In another embodiment, the system provides the crewmember with a directional cue (e.g., a visual, auditory, or tactile cue) corresponding to the prescribed trajectory back to the fixed location. The system may be activated by the crewmember or remotely by another crewmember and/or system.

A space device includes: an adhesion part to adhere to a target existing in the space; and a propulsion part to obtain propulsion power. The space device that adheres to the target at the adhesion part moves together with the target by the propulsion part, thereby conveying the target to a predetermined target position.

Provided is an apparatus for controlling an orbiting satellite by sensing a change in a yaw angle of the orbiting satellite and calculating a ground sample distance (GSD) based on the yaw angle. The apparatus may include a sensor configured to sense a yaw angle corresponding to yaw steering of the orbiting satellite, and a processor configured to calculate, based on the yaw angle, a GSD corresponding to a length of a pixel projected onto a planetary surface scanned by the orbiting satellite.

A system, method, and computer-readable storage devices for a 6U CubeSat with a magnetometer boom. The example 6U CubeSat can include an on-board computing device connected to an electrical power system, wherein the electrical power system receives power from at least one of a battery and at least one solar panel, a first fluxgate sensor attached to an extendable boom, a release mechanism for extending the extendable boom, at least one second fluxgate sensor fixed within the satellite, an ion neutral mass spectrometer, and a relativistic electron/proton telescope. The on-board computing device can receive data from the first fluxgate sensor, the at least one second fluxgate sensor, the ion neutral mass spectrometer, and the relativistic electron/proton telescope via the bus, and can then process the data via an algorithm to deduce a geophysical signal.

An attitude control system for a satellite is presented that can determine on time values for attitude control thrusters without use of attitude rate sensors, such as those based on gyros. The attitude control systems uses the attitude values from a star tracker to determine both attitude adjustment values and attitude adjustment rate values directly from the star tracker values, where the processing is performed using quaternions. From the attitude adjustment values and attitude adjustment rate values, a set of thruster on time values are determined.

A system may include an object having a rotational axis. The system may also include a first mass movably mounted to the object and configured to adjust a moment of inertia of the object by translating relative to the object along an inertial path having a first component that is perpendicular to the rotational axis. The system may additionally include a second mass movably mounted to the object and configured to: apply to the object a first torque in a first direction along the rotational axis while the moment of inertia is adjusted above a threshold value and apply to the object a second torque in a second direction along the rotational axis while the moment of inertia is adjusted below the threshold value, by translating relative to the object along a torque path having a second component that is perpendicular to the rotational axis and the first component.

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.

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, and/or compactible structures and deployment mechanisms used to form and deploy such satellite modules and power generation tiles associated therewith are provided. Each satellite module and/or power generation tile may be formed of a compactable structure and deployment mechanism capable of reducing the payload area required to deliver the satellite module to an orbital formation within the space-based solar power station and reliably deploy it once in orbit.

The flight path of a two-staged rocket (1) is periodically calculated during flight, and the estimated impact point when a first-stage rocket body (11) or a fairing (15) is separated and discarded is periodically calculated from a second-stage rocket (13) at each passing scheduled point in the predicted flight path. As long as there is a passing scheduled point such that the estimated impact points of the first-stage rocket (11) and the fairing (15) are both within the safe area, a process is periodically performed to designate the passing scheduled point safe and nearest to ILL of the two-staged rocket (1) as the separate-and-discard point of the first-stage rocket body (11), and when the two-staged rocket (1) reaches this separate-and-discard point, the first-stage rocket body (11) or the fairing (15) is separated and discarded.