A magnetically suspended control moment gyroscope comprising: a gimbal; a flywheel system, set in the gimbal; wherein the flywheel system comprises: a housing; a shaft, arranged in an inner cavity of the housing; a radial magnetic bearing, comprising: a first rotor portion and a first stator portion fixed to the shaft; an upper axial magnetic bearing and a lower axial magnetic bearing, wherein the upper axial magnetic bearing is fixed to an upper end of the first stator portion, the lower axial magnetic bearing is fixed to a lower end of the first stator portion; a wheel body, set in the radial magnetic bearing, fixed to the first rotor portion; an upper axial thrust plate and a lower axial thrust plate, wherein the upper axial thrust plate is fixed to an upper end of the wheel body, and is on an upper end of the upper axial magnetic bearing, the lower axial thrust plate is fixed to a lower end of the wheel body, and is under a lower end of the lower axial magnetic bearing.

An antenna system is configured for use in Low Earth Orbit (LEO) around Earth. The system has a plurality of antenna satellites coupled together to form a phased array. Each of the plurality of antenna satellites have an antenna body with an antenna and a solar cell. A processing device determines an orientation of the plurality of antenna satellites and position the phased array in the orientation based on an analysis of the solar cell of the antenna bodies facing the sun, the antenna of the antenna bodies facing the Earth, and maintaining a torque equilibrium of the phased array.

A space aircraft including a fuselage, two wings arranged on either side of the fuselage, and two nacelles arranged at the ends of the wings and each carrying a horizontal tail and a vertical tail, the fuselage having a cross section of variable size along the longitudinal axis with a maximum cross section being located in a longitudinal position located in front of the longitudinal position of the leading edges of the wings at the fuselage, making it possible in particular to help prevent the space aircraft from losing longitudinal static stability, the space aircraft thus having an optimized design and architecture which are suitable for the severe conditions encountered by such a space aircraft, in particular during atmospheric re-entry.

A space aircraft including a fuselage, two wings arranged on either side of the fuselage, and two nacelles arranged at the ends of the wings and each carrying a horizontal tail and a vertical tail, the fuselage having a cross section of variable size along the longitudinal axis with a maximum cross section being located in a longitudinal position located in front of the longitudinal position of the leading edges of the wings at the fuselage, making it possible in particular to help prevent the space aircraft from losing longitudinal static stability, the space aircraft thus having an optimized design and architecture which are suitable for the severe conditions encountered by such a space aircraft, in particular during atmospheric re-entry.

The present disclosure belongs to the technical field of space satellite propulsion, and particularly relates to a digital filter based method for measuring thrust response time of a satellite-borne micro-thruster. The method for measuring thrust response time in the present disclosure includes the following steps: S1: zeroing non-zero initial conditions of a torsional pendulum thrust measurement system to obtain an oscillating differential equation for a thrust measurement system after variable substitution; S2: connecting the digital filter in series behind the thrust measurement system after variable substitution to obtain an equivalent-sensitivity high-frequency thrust measurement system; S3: determining a system response of the equivalent-sensitivity high-frequency thrust measurement system; S4: determining and reading thrust response time of the satellite-borne micro-thruster from the system response; and S5: computing thrust to be measured by means of the system response inversely, and further confirming the thrust response time.

The present disclosure belongs to the technical field of space satellite propulsion, and particularly relates to a digital filter based method for measuring thrust response time of a satellite-borne micro-thruster. The method for measuring thrust response time in the present disclosure includes the following steps: S1: zeroing non-zero initial conditions of a torsional pendulum thrust measurement system to obtain an oscillating differential equation for a thrust measurement system after variable substitution; S2: connecting the digital filter in series behind the thrust measurement system after variable substitution to obtain an equivalent-sensitivity high-frequency thrust measurement system; S3: determining a system response of the equivalent-sensitivity high-frequency thrust measurement system; S4: determining and reading thrust response time of the satellite-borne micro-thruster from the system response; and S5: computing thrust to be measured by means of the system response inversely, and further confirming the thrust response time.

Methods, systems, and apparatus for designing, constructing and using instrument landers for in situ exploration of small solar system bodies, such as asteroids and comets. In one aspect, a lander includes a CubeSat-style platform; instrument packaging, wherein the CubeSat-style platform and the instrument packaging are configured and arranged for an uncontrolled descent, hopping landing on a surface of a body in a solar system, where a descending trajectory for the lander is designed based on gravitational force and solar radiation, with no lander-based propulsion; and a mobility mechanism configured and arranged to self-orient the lander on the surface of the body in the solar system.

Systems and methods for pointing photovoltaic arrays for optimal power generation. One or more methods among a plurality of methods for pointing an array may be used by a spacecraft control system to point the array. Example methods to use to point the photovoltaic array relate to analyzing current output, analyzing image data, and analyzing computational knowledge of reflective bodies or light sources. The spacecraft may be further controlled to reduce shadow by re-orienting, receiving light reflected off spacecraft, and orienting a photovoltaic array relative to incoming light sources based on topographic properties of the array such as cell grooves.

Systems and methods for pointing photovoltaic arrays for optimal power generation. One or more methods among a plurality of methods for pointing an array may be used by a spacecraft control system to point the array. Example methods to use to point the photovoltaic array relate to analyzing current output, analyzing image data, and analyzing computational knowledge of reflective bodies or light sources. The spacecraft may be further controlled to reduce shadow by re-orienting, receiving light reflected off spacecraft, and orienting a photovoltaic array relative to incoming light sources based on topographic properties of the array such as cell grooves.

A method of controlling a satellite and a computer-readable recording medium are provided. The method is for controlling a satellite moving along an orbit having an inclination angle from the equatorial plane to capture due-north images. The method includes: determining a position of the satellite; calculating a roll angle and a pitch angle of the satellite for pointing a line-of-sight vector of the satellite to a first ground surface being a photographing point; determining a compensation angle by considering effects of the inclination angle and rotation of the Earth so as to capture images in the due north direction of the photographing point; calculating a yaw angle based on the compensation angle; and rotating the satellite according to the calculated roll angle, pitch angle, and yaw angle.