Systems and methods are described herein for mounting a thruster onto a vehicle. A thruster mounting structure may comprise a first, second, and third rotational joint, a boom, and thruster pallet, and a thruster attached to the thruster pallet. The first rotational joint may be attached to the vehicle and configured to rotate in a first axis. The first rotational joint may be connected to the boom and configured to pivot the boom about the first axis. The boom may be connected to the second rotational joint, which is connected to the third rotational joint and configured to rotate the third rotational joint in the first axis. The third rotational joint may be connected to the thruster pallet and configured to pivot the thruster pallet in a second axis that is perpendicular to the first axis.

Systems and methods are described herein for mounting a thruster onto a vehicle. A thruster mounting structure may comprise a first, second, and third rotational joint, a boom, and thruster pallet, and a thruster attached to the thruster pallet. The first rotational joint may be attached to the vehicle and configured to rotate in a first axis. The first rotational joint may be connected to the boom and configured to pivot the boom about the first axis. The boom may be connected to the second rotational joint, which is connected to the third rotational joint and configured to rotate the third rotational joint in the first axis. The third rotational joint may be connected to the thruster pallet and configured to pivot the thruster pallet in a second axis that is perpendicular to the first axis.

A method for providing thermal balance of spacecraft electronics is provided. The spacecraft includes two or more electronic units wherein each electronic unit is capable of performing the same spacecraft operational task. The method for balancing the temperature of spacecraft electronics further includes providing each of the two or more electronic units with a temperature sensor for determining the temperature of that electronics unit. The electronic units and their respective temperature sensors are connected to a controller. In the event that the controller determines that the temperature of an activated first electronics unit has reached or exceeded a predetermined threshold, and the controller has determined that the temperature of a second deactivated electronics unit is below a predetermined threshold, the controller automatically deactivates the first electronics unit and activates the second electronics unit to perform the task previously being performed by the first electronics unit. This process continues automatically.

A system includes a first wire adapted to transmit a first current defined by a first waveform comprising a positive current value, 11, for a first period of time, no current for a second period of time, a negative current value, -I1, for a third period of time and no current for a fourth period of time, wherein the first period, second period, third period and fourth period are all approximately equal in duration and wherein the first waveform repeats itself, and a second wire generally parallel to the first wire and separated by a distance, d, wherein the second wire is adapted to transmit a second current defined by a second waveform comprising the first waveform time shifted by an amount of time approximately equal to the first period of time, wherein each period of time is approximately equal to an amount of time required for light to travel the distance, d.

Provided is a configuration construction and attitude control method for a pyramid deorbit sail. By taking into consideration environmental perturbation like atmospheric resistance perturbation and non-spherical earth perturbation, a dynamics model featuring three-dimensional orbit-and-attitude coupling based on position vectors and quaternion descriptions, the deorbit sail is taken as a rigid body, a spacecraft body is taken as a mass point, airflow obstruction is considered in the windward area, thereby improving the precision of the dynamics model; based on this model, the law of influence of the configuration parameters in the deorbit sail, such as a cone angle and a strut length, on the attitude stability and deorbiting efficiency of the spacecraft in different cases is analyzed, the configuration parameters of the pyramid deorbit sail system are analyzed and optimized according to the derived law, so as to obtain a pyramid deorbit sail achieving high attitude stability and high deorbiting efficiency.

A modular micro-cathode arc thruster for use in satellites. An exemplary satellite has a plurality of stacked modular arc thrusters, each having an external anode, an internal cathode, and an insulator therebetween. The arc thrusters are situated in a housing, wherein the housing has an opening to eject exhausted thrusters. Once an arc thruster is expended, the push rod ejects that arc thruster and the next arc thruster takes its place.

[Object] To provide a sheet-like structure capable of highly accurately estimating a sheet-like shape.

[Solving Means] A sheet-like structure includes a sheet-like member and a plurality of detection sensors. The sheet-like member extends along an in-plane direction orthogonal to a thickness direction and receives light incident on the sheet-like member. The plurality of detection sensors are dispersedly arranged on the sheet-like member along the in-plane direction and are for detecting an incident angle of the light with respect to the sheet-like member at each arrangement position of the plurality of detection sensors.

The present invention relates to a neutralizer (4) for an ion thruster (1) of a spacecraft (S), comprising: a cathode (5) for emission of electrons (6), a support (7) with an opening (8) inside which the cathode (5) is supported in a radially spaced manner, and an electrically conductive shielding (9) which surrounds said opening (8) and is electrically insulated from the support (7), wherein a ring (11) is mounted between the shielding (9) and the cathode (5) and is electrically insulated from the shielding (9) and radially spaced from the cathode (5).

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 system for laser-driven propulsion, system comprising a laser source and a target comprising an accelerating part and a projectile part, the accelerating part comprising a metal layer and a porous layer pressed against the metal layer; wherein the laser source is selected to emit pulse beams directed to the metal layer at a fluence below the plasma ablation threshold of the material of the metal layer.