A transfer type contra-rotating geomagnetic energy storage-release delivery system is disclosed. The system includes a control system, a three-axis control moment canceller and an energy system, which are arranged on a delivery mother spacecraft, and the delivery mother spacecraft is connected, through support rod structures, with a strong magnetic moment generating device, a contra-rotating transmission mechanism and two delivery connection rod structures arranged at the two ends of the contra-rotating transmission mechanism, the strong magnetic moment generating device is arranged between the contra-rotating transmission mechanism and the delivery mother spacecraft, the two delivery connection rod structures are provided with slidable mass blocks respectively, and the strong magnetic moment generating device and the contra-rotating transmission mechanism provide energy through the energy system. The strong magnetic moment generating device is free of accelerated rotation of an attitude, thereby decoupling the dual coupling.

Described herein are systems and methods for attitude determination using infrared Earth horizon sensors (EHSs) with Gaussian response characteristics. Attitude information is acquired by detecting Earth's infrared electromagnetic radiation and, subsequently, determining the region obscured by Earth in the sensors' fields of view to compute a nadir vector estimation in the spacecraft's body frame. The method can be applied when two sensors, each with known and distinct pointing directions, detect the horizon, which is defined as having their fields of view partially obscured by Earth. The method can be implemented compactly to provide high-accuracy attitude within small spacecraft, such as CubeSat-based satellites.

A space vehicle includes one or more magnetorquers operable to change an attitude of the space vehicle in an external magnetic field, each magnetorquer comprising a coil, and a switching circuit for short-circuiting the coil of at least one of the magnetorquers so that a closed electric circuit comprising said coil is formed, for damping tumbling motion of the space vehicle in the external magnetic field. The switching circuit is configured to short-circuit the coil of the at least one magnetorquer upon occurrence of a condition indicative of end-of-life or failure of the space vehicle. The application further relates to a corresponding method of operating a space vehicle.

A space vehicle includes one or more magnetorquers operable to change an attitude of the space vehicle in an external magnetic field, each magnetorquer comprising a coil, and a switching circuit for short-circuiting the coil of at least one of the magnetorquers so that a closed electric circuit comprising said coil is formed, for damping tumbling motion of the space vehicle in the external magnetic field. The switching circuit is configured to short-circuit the coil of the at least one magnetorquer upon occurrence of a condition indicative of end-of-life or failure of the space vehicle. The application further relates to a corresponding method of operating a space vehicle.

A control system configured to execute a safing mode sequence for a spacecraft is disclosed. The control system includes one or more star trackers that each include a field of view to capture light from a plurality of space objects surrounding the celestial body. The control system also includes one or more actuators, one or more processors in electronic communication with the one or more actuators, and a memory coupled to the one or more processors. The memory stores data into a database and program code that, when executed by the one or more processors, causes the control system to determine a current attitude of the spacecraft, and re-orient the spacecraft from a current attitude into a momentum neutral attitude.

[Object] To provide an electron source that is lightweight, simple in configuration, and capable of suppressing characteristic degradation or recovering characteristics without causing an increase in power consumption.

[Solving Means] A CNT electron source includes: a CNT emitter 32 for emitting electrons; a gate electrode 33 for extracting electrons from the CNT emitter 32; and a gate power supply connection switching relay 37a and a CNT emitter grounding switching relay 37b that cause the gate electrode 33 to emit electrons to irradiate the CNT emitter with electrons.

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.

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.

A spacecraft nutation inhibition method for low-orbit geomagnetic energy storage in-orbit delivery includes: S1, enabling a delivery connection rod to be slidably connected to two mass blocks in a length direction, and adjusting the center of mass of a spacecraft system to pass through a main connecting shaft; S2, respectively measuring, calibrating and adjusting the center of mass and the principal axis of inertia of the delivery connection rod that is to deliver the space target or de-orbit debris; S3, carrying out energy storage delivery; S4, respectively adjusting the center of mass and the moment of inertia of the delivery connection rod after delivering the space target or de-orbit debris; S5, carrying out energy dissipation and unloading; and S6, enabling the spacecraft system to prepare to grab the next space target or de-orbit debris and proceeding to the next delivery work cycle.

A spacecraft nutation inhibition method for low-orbit geomagnetic energy storage in-orbit delivery includes: S1, enabling a delivery connection rod to be slidably connected to two mass blocks in a length direction, and adjusting the center of mass of a spacecraft system to pass through a main connecting shaft; S2, respectively measuring, calibrating and adjusting the center of mass and the principal axis of inertia of the delivery connection rod that is to deliver the space target or de-orbit debris; S3, carrying out energy storage delivery; S4, respectively adjusting the center of mass and the moment of inertia of the delivery connection rod after delivering the space target or de-orbit debris; S5, carrying out energy dissipation and unloading; and S6, enabling the spacecraft system to prepare to grab the next space target or de-orbit debris and proceeding to the next delivery work cycle.