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.

Methods and apparatus to methods and apparatus for performing propulsion operations using electric propulsion system are disclosed. An example apparatus includes means to use an electric propulsion system coupled to a frame of a spacecraft, the electric propulsion system including at least a first thruster and a second thruster, the first thruster adjacent a first side of the frame, the second thruster adjacent a second side of the frame, and means to allow at least one of the first thruster or the second thruster to control the spacecraft without using a chemical propulsion system.

A wide field-of-view celestial sighting system and method are provided. The method includes orienting an imaging optic to collect light from at least one light source, such as at least one celestial body, the imaging optic being secured to a platform. The method further includes selectively collecting light from the at least one celestial body through a selective light collector secured to the platform and positioned in an imaging surface, such as an imaging plane, of the imaging optic. The method further includes combining forward scattered light from the at least one celestial body to provide a combined forward scattered light, and detecting a light intensity of the combined forward scattered light. Systems for performing the method are provided.

In an embodiment, there is provided a method for determining a spacecraft instantaneous velocity using starlight. The method includes determining, by a star pairing module, a plurality of star pairs in a selected star field image. The selected star field image includes a plurality of star images. The method includes estimating, by a line of sight estimation module, an apparent bearing direction to each star in at least some of the plurality of star pairs; determining, by the line of sight estimation module, a respective apparent inter-star angle for each star pair of the least some star pairs; and estimating, by a velocity estimation module, a spacecraft velocity to within a total velocity error based, at least in part, on the apparent inter-star angles.

Apparatus and methods for controlling a spacecraft for a transfer orbit. The spacecraft includes a propulsion subsystem with electric thrusters that are installed with two-axis gimbal assemblies. The spacecraft also includes a controller that identifies a target spin axis for the spacecraft, determines an actual spin axis for the spacecraft during the transfer orbit, determines gimbal angles for the electric thruster(s) that adjust the actual spin axis toward the target spin axis, and initiates a burn of the electric thruster(s) at the gimbal angles.

Techniques for orienting an earth-orbiting spacecraft include determining, using a star tracker on board the spacecraft, a first vector aligned between an ecliptic pole of the earth and the spacecraft, adjusting attitude of the spacecraft so as to align a first axis of the spacecraft with the first vector, and rotating the spacecraft about the first axis until presence of the sun is registered. Rotation rates may be subsequently reduced, such that the sun remains within a field of view of the sun sensor or of a solar array of the spacecraft.

A torque generation system includes: a plurality of solar array panels and/or solar array panel divisions; and a torque controller configured to control an electricity generation ratio of each of the plurality of solar array panels and/or solar array panel divisions to generate torque.

A menu-type design method based on optimized information fusion applied to a GEO satellite control system is provided, which includes: configuring four long-life inertial attitude sensor gyroscopes for a long-life GEO satellite control system; configuring sensors capable of measuring three-axis attitude according to a menu-type design requirement on hardware, where the long-life inertial attitude sensor gyroscopes and the sensors capable of measuring three-axis attitude are combined to form three types of Kalman filters; autonomously sorting, by the satellite-borne computer application software, the Kalman filters; and in a case where an FDIR module detects a fault, autonomously generating, by the FDIR module, an alarm corresponding to the fault, and autonomously performing, by a currently selected Kalman filter, reduced-order filtering, and in a case where the fault is not eliminated within a set time period, issuing, by the FDIR module, a macro instruction sequence to perform autonomous reorganization.

Methods, apparatus, and articles of manufacture to minimize command dynamics of a satellite are disclosed. An example apparatus includes a steering law module to calculate a first set of vectors to maneuver a space vehicle, and calculate a second set of vectors based on projecting the first set of vectors onto a fixed plane. The apparatus further includes an attitude controller to generate an attitude command based on the first and the second sets of vectors to prevent an unplanned rotation by the space vehicle.

A system for emergency crew return and down-mass orbit comprising a stowable, self-contained, deployable maneuvering reentry vehicle for automated, on-demand reentry to ground for cargo of 1-10 kilograms or up to single or multiple human use for evacuation of orbital facilities. The system includes a deployable aeroshell that is contiguous (a single geometric objectsurface or hollow shapethat can morph in 3D shape), modular (a collection of modular components externally acting as a contiguous shape, but morphed in 3D via actuators contained in each modular member to create a general asymmetric geometry), or discontiguous (a collection of independently controlled surfaces or bodies that morph to form desirable asymmetric drag configurations). The system contains traditional spacecraft guidance, navigation and control, propulsion, and attitude control elements, in addition to communications, power, and actuator energetics systems for controlling the vehicle aeroshell shape during reentry, thus, minimizing the landing footprint of the vehicle.