A method of guidance for placing a satellite on station comprises the following steps carried out during a predefined current cycle: A) determining on the ground a law of orientation of the thrust vector, and a history of state variables and of adjoint state variables of the satellite for the transfer from a starting orbit to a target orbit using optimal control theory, B) determining on the ground a law of evolution of the rotation of the satellite about the thrust vector, on the basis of the orientation law and of the history, C) representing according to a predetermined format the evolution of the state variables and adjoint state variables so as to obtain first parameters, D) representing according to a predetermined format a law of evolution of the rotation so as to obtain second parameters, E) concatenating the first and second parameters so as to obtain a guidance plan for the satellite, F) downloading onboard the guidance plan, G) periodically repeating according to a predefined period which is smaller than the duration of the guidance cycle: g1) reconstructing onboard the satellite a guidance instruction, g2) executing onboard the satellite the instruction by applying a closed control loop, H) measuring on the ground the real orbital trajectory of the satellite, I) repeating steps A) to H) with the trajectory measured at the end of the cycle as starting orbit of the following cycle, until the target orbit is attained.

An orbit transfer method for a spacecraft using a continuous or quasi-continuous thrust propulsion, the method comprises: the acquisition, at least once in each half-revolution of the spacecraft, of measurements of its position and of its velocity; the computation of a thrust control function as a function of the measurements; and the driving of the thrust in accordance with the control law; wherein the control law is obtained from a Control-Lyapunov function using orbital parameters, preferably equinoctial, of the spacecraft, averaged over at least one half-revolution. An embedded driving system for a spacecraft for implementing such a method and a spacecraft equipped with the driving system are provided.

A metal encapsulated ceramic tile thermal insulation system for rockets and associated methods is disclosed. A representative system includes a launch vehicle having a first end and a second end generally opposite the first end and includes a heat shield positioned at the second end. The heat shield includes a plurality of thermal protection apparatuses, where individual of the thermal protection apparatuses include ceramic tiles encapsulated by inner and outer metal layers, which are positioned on opposing top and bottom surfaces of the ceramic tiles. The plurality of thermal protection apparatuses includes a plurality of pins positioned within corresponding holes drilled through the ceramic tiles and are secured to the metal layers. The outer metal layer can protect the ceramic tile from tool strikes and debris and can also prevent water from reaching and being absorbed by the ceramic tile.

Embodiments of the present disclosure are directed to techniques for autonomously transitioning a spacecraft from a power-saving state to a power-consuming state at a time after launch of the spacecraft on a launch vehicle. Because the spacecraft can autonomously detect conditions for transitioning to the power-consuming state, commands received via an umbilical connection to the launch vehicle, or detecting the presence or absence of such a connection, is unnecessary, thereby removing several technical barriers to eliminating such umbilical connections altogether. In some embodiments, low-cost vacuum detection devices that use very small amounts of power may be used by the spacecraft to detect when the spacecraft has reached an altitude suitable for transitioning to the power-consuming state.

Techniques for deploying a plurality of smallsats from a common launch vehicle are disclosed where a structural arrangement provides a load path between an upper stage of the launch and the plurality of spacecraft. Each spacecraft is mechanically coupled with the launch vehicle upper stage only by the structural arrangement. The structural arrangement includes at least one trunk member that is approximately aligned with the longitudinal axis of the launch vehicle upper stage, a plurality of branch members, each branch member being attached to the trunk member and having at least a first end portion that is substantially outboard from the longitudinal axis; and a plurality of mechanical linkages, each linkage coupled at a first end with a first respective spacecraft and coupled at a second end with one of the plurality of branch members, the trunk member or a second respective spacecraft.

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 system for air-launching a liquid fueled rocket launch vehicle using a tubular rocket support structure for holding the launch vehicle and for supplying the launch vehicle with make-up cryogenics, electrical power, and control signals, and for providing coupling to a launch-assist aircraft. The tubular rocket support structure contains cryogenic fluids, in addition to fuel and oxidizer, to cool the fuel and oxidizer during the pre-launch phase. The tubular rocket support structure has features that keep the liquid fuel and oxidizer from sloshing away from the tank outlet ports to the launch vehicle's rocket engine after release from the aircraft but before rocket launch. In operation, the aircraft controllably releases the tubular rocket support structure containing the launch vehicle, and the launch vehicle then launches from the rocket support structure.

A rocket launch system includes a tubular rocket launcher carriage with electromotive cableway traction drives conveyed beneath a two axis pivot anchored to the earth, elevated into a co-axial transfer tube leading to three primary tether cables whose weight is offset by balloons. The carriage is conveyed to a docking station supported into the stratosphere by a pair of secondary cables suspended under an attachment frame for tensioning balloons. The carriage is engaged by a carriage end gripper guided by two sets of secondary cables and two sets of tertiary cables and lifted by a lower hoist guided by the secondary cables to a lift ring assembly. This lower hoist is supported by an upper hoist suspended from the tensioning balloons attachment frame. The carriage, which engages a lift ring guided by two secondary cables, is elevated further, rotated in azimuth and elevation, and rocket ejection occurs from a launch tube during freefall of the carriage, with engine ignition occurring at a safe distance. The carriages have traction drives which grip cables from which they derive power and rotate to drive the carriage from the low altitude to the high altitude. The traction drives rotate in the opposite direction as the carriage descends the cable following the launch of a rocket under gravitational force. The kinetic energy of the traction drive is converted to electrical energy which is fed back to the cables during descent of the carriage.

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.

Apparatus and methods for determining orbital parameters for spacecraft are provided. A computing device can receive a first plurality of orbital parameters defining a first orbit by a first spacecraft of a particular object. The first orbit can have a corresponding first ground track over the particular object. The computing device can receive a following time for a second spacecraft. The computing device can determine a second plurality of orbital parameters for a second orbit by the second spacecraft of the particular object based on the first plurality of orbital parameters and the following time. The second orbit can have a corresponding second ground track over the particular object that follows the first ground track. The computing device can generate an output that includes the second plurality of orbital parameters.