The present disclosure relates to an axial flux motor comprising a stator and a rotor. The stator comprises a first motor coil, and a second motor coil, and the rotor comprises a first and second actuator magnet array configured in an alternating axial polarity arrangement and a first rotating magnetic return path member.

The present disclosure relates to an axial flux motor comprising a stator and a rotor. The stator comprises a first motor coil, a second motor coil, a first hall sensor, and a second hall sensor, and the rotor comprises a rotor platform member, an actuator magnet array arranged in an alternating axial polarity arrangement, a trigger magnet array, and a rotating magnetic return path member.

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 satellite configuration includes a plurality of individual satellite buses each having a number of side panels that form a polygonal shape, where the individual satellite buses collectively fit together to form the satellite configuration having a regular polygon shape. A method of producing the satellite configuration includes forming a plurality of individual satellite buses each having a polygonal shape, and fitting the individual satellite buses together to form the satellite configuration in a regular polygonal shape.

Systems and methods are described for a satellite control system that exhibits improved stability and increased efficiency by implementing a non-linear dynamic inversion inner-loop control algorithm coupled with an eigen vector outer-loop control algorithm. Thus, the attitude determination and control system (ADACS) may operate using commands to rotate directly about an eigen vector. Additionally, the outer-loop control system includes a feed-forward control element to enhance pointing accuracy when tracking moving targets.

A stackable pancake satellite that is configured so that a plurality of the satellites can be stacked within a payload fairing of a launch vehicle. Each satellite includes sections that are folded or rotated together prior to launch, and unfolded or rotated away from each other when deployed. A first section is a satellite body having a first side that acts as a thermal radiator and a second side opposite the first side that includes an antenna. A second section includes one or more solar panels attached adjacent to the first side of the satellite body. A third section includes a splash plate reflector attached adjacent to the second side of the satellite body that reflects signals between Earth and the antenna. When deployed, the solar panels are pointed towards the Sun and the splash plate reflector directs the signals between the Earth and the antenna.

A system for optimizing a low-thrust trajectory of a spacecraft trajectory for orbital transfer includes an interface to receive data, a memory to store scheduled geostationary transfer orbit (GTO) data and scheduled geostationary Earth orbit (GEO) data and computer-executable programs, and a processor. The processor is configured to provide a two-dimensional (2D) averaged trajectory consisting of a predetermined revolutions by executing the optimal control program using the GTO data and GEO data, arrange N equidistant points on the 2D averaged trajectory to form segments on the 2D averaged trajectory, obtain osculating elements corresponding to the segments by solving optimization problems for the segments, estimate initial guesses of the segments under continuous thrusting conditions in tangential directions at the N equidistant points, solve a minimum energy optimization problem to obtain a minimum energy 2D osculating trajectory by using as initial guess the concatenation of segments, compute a minimum energy three-dimensional (3D) osculating trajectory by linearly decreasing an inclination of the minimum energy 2D osculating trajectory to zero, and generating a minimum fuel 3D osculating trajectory by iteratively solving a cost function while changing a parameter from one to zero.

A space vehicle element configured to be at least partially destroyed during re-entry of a space vehicle into the atmosphere comprises a heat generating part comprising a metallo-thermal composition for providing additional heat during re-entry of the space vehicle into the atmosphere by an exothermic reaction of the metallo-thermal composition. The destruction of the space vehicle element is expedited by the additional heat provided by the heat generating part. The heat generating part is at least partially integrated within the space vehicle element or at least partially surrounds a portion of the space vehicle element. The application further relates to a corresponding method of manufacturing a space vehicle element configured to be destroyed during re-entry of the space vehicle into the atmosphere.

Methods and apparatus to methods and apparatus for performing propulsion operations using electric propulsion system are disclosed. An apparatus includes a space vehicle including means for performing propulsion operations without using a chemical propulsion system.

Energy efficient satellite maneuvering is described herein. One disclosed example method includes maneuvering a satellite that is in an orbit around a space body so that a principle sensitive axis of the satellite is oriented to an orbit frame plane to reduce gravity gradient torques acting upon the satellite. The orbit frame plane is based on an orbit frame vector.