Systems and methods for describing, simulating and/or optimizing spaceborne systems and missions. Configurations for spaceborne systems are generated and validated based on simulation output.

A lunar orbiting satellite system executes orbit planning of assigning a function (positioning, communication, and flashing) to an artificial satellite (AS) depending on a relative position of the AS to the moon at a time when the moon and the AS are observed from an input point on the earth, and correcting the relative position, which changes in accordance with the moon revolution period. The system includes: a satellite orbit planner which assigns a function to each ASs forming an AS group flying around the moon depending on a relative position of each ASs to the moon at a time when the moon and ASs are observed from an input point on the earth, and set a target orbit according to the function; and a satellite controller which causes the each ASs to execute control based on the function to implement switching of the function.

In an example, a method for deorbiting a spacecraft is described. The method includes selecting a target landing site for deorbiting the spacecraft. The method includes determining a range target and a velocity target for reaching a predicted atmospheric entry location. The method includes determining a back-propagated orbit state estimate of the spacecraft. The method includes comparing the back-propagated orbit state estimate to a known orbit state of the spacecraft to determine that the back-propagated orbit state estimate has converged with the known orbit state. The method includes calculating based on determining that the back-propagated orbit state estimate has converged with the known orbit state, (a) an estimated time of ignition for a propulsion system of the spacecraft and (b) an estimated burn velocity vector of the propulsion system using the range target and the velocity target. The method includes performing a burn pulse by the propulsion system.

The invention relates to a method for the safe release of artificial satellite in Earth orbit comprising providing an orbital transport spacecraft (1) able to move at orbital height and comprising a plurality of PODs (11) for releasing satellites (12) transported by the orbital transport spacecraft (1), housing said orbital transport spacecraft (1) in a space launcher (100) configured to reach an orbital height; generating a release signal and transmitting it to the orbital transport spacecraft (1) to release the orbital transport spacecraft (1) from the space launcher (100), in case of failure to release the orbital transport spacecraft (1) or in case of breakdown of the orbital transport spacecraft (1) after releasing from the space launcher (100), activating a safety subsystem (21) of the orbital transport spacecraft (1) to generate a POD (11) activation sequence to release the satellites (12).

The disclosure relates to a method and apparatus of rotating mass propulsion. The method and apparatus entails rotating a mass to generate thrust. Varying the speed and direction of rotation provides some control of the magnitude and direction of the thrust generated. The method and apparatus of the invention pertinent to a propulsion system for spacecrafts or astromotive vehicles under conditions of zero to low gravity and atmosphere.

A spacecraft including: an engine; a thrust vector control device controlling a thrust vector as a direction of a thrust acting on the spacecraft; and a main control device configured to acquire state quantities of the spacecraft in a powered descending in which the spacecraft is guided to a target point while the engine generates the thrust, and generate a throttling command by which combustion of the engine is controlled and an operation command by which the thrust vector control device is operated. The state quantities contain a first acceleration parameter and a second acceleration parameter. The first and second acceleration parameters are calculated as coefficients A and B obtained by fitting based on acceleration of the spacecraft detected at each time of past, supposing the following equation is satisfied between a reciprocal number 1/a of the acceleration a of the spacecraft and time t:

1/a=At+B(1).

A method for providing optimized power balanced low thrust transfer orbits utilizing split thruster execution to minimize an electric orbit raising duration of an apparatus includes monitoring an electric power balance on the apparatus. The method also includes firing a first thruster in response to the apparatus exiting an eclipse and based on the electric power balance. The method additionally includes firing a second thruster at a predetermined time delay after firing the first thruster based on the electric power balance. The method additionally includes ending firing one of the first thruster or the second thruster after a predetermined time duration based on the electric power balance. The method further includes ending firing another of the first thruster or the second thruster in response to the apparatus entering a next eclipse.

In various embodiments, while a self-driving model operates a vehicle, a user monitoring subsystem acquires sensor data associated with a user of the vehicle and a vehicle observation subsystem acquires sensor data associated with the vehicle. The user monitoring subsystem computes values for a psychological metric based on the sensor data associated with the user. Based on the values for the psychological metric, a feedback application determines a description of the user over a first time period. The feedback application generates a dataset based on the description and the sensor data associated with the vehicle. Subsequently, a training application performs machine learning operation(s) on the self-driving model based on the dataset to generate a modified self-driving model. Advantageously, the dataset enables the training application to automatically modify the self-driving model to account for the impact different driving actions have on the user.

The invention relates to small satellites capable to fly in formation (10), in particular nano- or picosatellites with a mass of 10 kg or less, for LEO applications, comprising a housing (12) and at least one plug-in board (14) arranged in the housing (12) with a predetermined functionality and a propulsion system (16) for generating a directed pulse in the direction of the flight trajectory T.sub.k.

It is proposed that the small satellite (10) comprises an independent and autonomously working collision avoidance system (18), which is capable of adapting a trajectory correction T.sub.kk of the trajectory T.sub.k by the propulsion system (16), when a collision with an object (30) is expected.

In a further independent aspect, the invention relates to a formation (100) composed of several small satellites capable to fly in formation (10), wherein a relative position and flight trajectory T.sub.k of each small satellite (10) is modifiable via the independently and autonomously working collision avoidance system (18).

Systems and methods include determining a collision probability between a first spacecraft and a second spacecraft; responsive to a determination that the collision probability exceeds a threshold, identifying a plane that is orthogonal to a relative velocity between the first spacecraft and the second spacecraft; determining a boundary on the plane that, when intersected by the first spacecraft, reduces the collision probability to a threshold level; selecting a first test point along the boundary and a second test point along the boundary; determining a first cost associated with maneuvering the first spacecraft towards the first test point and a second cost associated with maneuvering the first spacecraft towards the second test point, the first cost and the second cost based on one or more user-defined parameters; and maneuvering the first spacecraft toward a target point along the boundary that is determined based on the first cost and the second cost.