Systems and methods for satellite orbit and momentum control are disclosed herein. An example spacecraft includes a first arm having a first end pivotably coupled to the spacecraft and a second end. The example spacecraft includes a first thruster rotatably coupled to the second end of the first arm. The first arm is to selectively pivot about the spacecraft to position the first thruster to generate a change in velocity of the spacecraft to maintain the spacecraft within an orbit station.

Described are techniques for determining measures of risk through a medium fidelity, fast running flight risk analysis system, for evaluating risk of a launch or reentry. The method may include receiving trajectory information associated with a planned flight, wherein the trajectory information comprises a series of trajectory vectors, each trajectory vector comprising time information, position information and velocity information. Identification of consequences of interest, and an extent of a region of interest for potential hazard areas may also be received. The method may include determining an upper bound of expected consequence for each population center based on a number of people in each population center subject to different consequences, determining one or more collective risk metrics for each population center, and a collective risk metric for all population centers.

The present invention provides a method for refined attitude control based on output feedback for a flexible spacecraft. The control method comprises the following steps of: a) building a flexible spacecraft dynamical system Σ.sub.1, converting the flexible spacecraft dynamical system Σ.sub.1 into a flexible spacecraft dynamical system Σ.sub.2, and incorporating spacecraft rigid-flexible coupling dynamic disturbance into the flexible spacecraft dynamical system Σ.sub.2; b) constructing an external system Σ.sub.3, and describing the rigid-flexible coupling dynamic disturbance through the external system Σ.sub.3; c) configuring a disturbance observer for estimating the value of the rigid-flexible coupling dynamic disturbance; d) configuring a dynamic output feedback H.sub.∞ controller; e) compounding the disturbance observer in step c) with the dynamic output feedback H.sub.∞ controller in step d) to obtain a flexible spacecraft refined attitude control system Σ.sub.6; the flexible spacecraft refined attitude control system Σ.sub.6 compensating for the rigid-flexible coupling dynamic disturbance through the estimated value.

A constellation of Earth-orbiting spacecraft includes a first spacecraft disposed in a first orbit, a second spacecraft disposed in a second orbit, and a third spacecraft disposed in a third orbit. Each of the first orbit, the second orbit and the third orbit is substantially circular with a radius of approximately 42,164 km, and has a specified inclination with respect to the equator within a range of 5° to 20°. The first orbit has a first right ascension of ascending node RAAN1, the second orbit has a second RAAN (RAAN2) approximately equal to RAAN1+120°, and the third orbit has a third RAAN (RAAN3) approximately equal to RAAN1+240°. A fourth spacecraft is disposed in a fourth orbit that has a period of approximately one sidereal day, an inclination of less than 2°, a perigee altitude of at least 8000 km, and an eccentricity between approximately 0.4 and 0.66.

A method for thermally stabilizing a communication satellite in orbit around the Earth relies on the discrete rotational symmetry of the pattern of antenna beams of the satellite. Exploiting the symmetry, the orientation of the satellite is changed from time to time by rotating the satellite through a symmetry angle of the rotational symmetry. Because of the symmetry, the beam pattern is unchanged after the rotation; but, because the rotation angle is less than 360°, a different side of the satellite is exposed to sunlight. The use of different thermal radiators and thermal shields on different sides of the satellite means that the thermal budget of the satellite is different after the rotation. By judiciously applying rotations as needed, as the orbit's orientation relative to the Sun evolves in time, it is possible to achieve effective control on the thermal budget of the satellite.

A new approach for rapid slew and settle of small satellites is based on four single degree-of-freedom control moment gyroscopes with variable speed flywheels (or reaction wheels) in a pyramid configuration, combined with path and endpoint constraint time-optimal control. The path and endpoint constrained time-optimal control can be augmented with momentum management without the use of additional actuators.

A satellite in orbit around a planetary body includes a bus and a drag flap coupled to the bus. The drag flap is used to increase the drag torque applied to the satellite. The bus may house sensors and actuators, such as a star tracker, a gyroscope, a reaction wheel, and a global position system (GPS) receiver to monitor the attitude of the satellite in response to the applied drag torque. The measurements from the sensors and actuators may be used to determine the drag torque applied to the satellite. An estimate of the atmospheric density may be then be determined based on the drag torque. Compared to conventional approaches, the satellite and methods described herein estimates the atmospheric density at comparable, if not better, resolution and bandwidth. The atmospheric density estimates may also be acquired in real-time using a cheaper, lighter, and smaller satellite.

A satellite in orbit around a planetary body includes a bus and a drag flap coupled to the bus. The drag flap is used to increase the drag torque applied to the satellite. The bus may house sensors and actuators, such as a star tracker, a gyroscope, a reaction wheel, and a global position system (GPS) receiver to monitor the attitude of the satellite in response to the applied drag torque. The measurements from the sensors and actuators may be used to determine the drag torque applied to the satellite. An estimate of the atmospheric density may be then be determined based on the drag torque. Compared to conventional approaches, the satellite and methods described herein estimates the atmospheric density at comparable, if not better, resolution and bandwidth. The atmospheric density estimates may also be acquired in real-time using a cheaper, lighter, and smaller satellite.

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 reusable staging system comprising: a processor-based device configured to monitor one or more rocket stages of a launch vehicle having a payload, wherein the processor-based device has at least one interface communicating with the one or more rocket stages of the launch vehicle; and a memory device for storing data and executing software routines, and wherein the reusable staging system is disposed within the payload of the launch vehicle, and wherein the reusable staging system is configured to actively monitor flight-related data to detect one or more detach requirements; and further configured to release the one or more rocket stages when the one or more detach requirements is met.