The present invention relates to a single-board processor card configured for use in a 1U CubeSat payload form-factor multi-purpose architecture, including: a field-programmable-gate-array (FPGA) which is reconfigurable in flight; wherein a configuration memory of the FPGA can be scrubbed in flight to correct errors or upsets; and a radiation-hardened monitor (RHM) which provides radiation mitigation and system monitoring of the single-board processor card, and which reconfigures said FPGA during flight, scrubs the configuration memory, and monitors a health of the FPGA. The 1U CubeSat payload form-factor multi-purpose architecture includes a backplane having a plurality of slots, one of the plurality of slots which accommodates the single-board processor card, wherein the backplane routes signals to a plurality of standard-sized processor cards, interchangeably disposed in any of the plurality of slots.

Aspects of wireless communication are described, including a radiofrequency (RF) amplifier chip, configured for transmitting or receiving data, comprising a first substrate comprising a first material and a second substrate comprising a second material that is different from the first material. The first substrate and the second substrate may be lattice-matched such that an interface region between the first substrate and the second substrate exhibits an sp3 carbon peak at about 1332 cm.Math..sup.1 having a full width half maximum of no more than 5.0 cm.Math..sup.1 as measured by Raman spectroscopy. In some aspects, the first substrate and said second substrate permit said chip to transmit or receive data at a transfer rate of at least 500 megabits per second and a frequency of at least 8 GHz. In some aspects, the RF amplifier chip is part of a satellite transmitter.

Aspects of wireless communication are described, including a radiofrequency (RF) amplifier chip, configured for transmitting or receiving data, comprising a first substrate comprising a first material and a second substrate comprising a second material that is different from the first material. The first substrate and the second substrate may be lattice-matched such that an interface region between the first substrate and the second substrate exhibits an sp3 carbon peak at about 1332 cm.Math..sup.1 having a full width half maximum of no more than 5.0 cm.Math..sup.1 as measured by Raman spectroscopy. In some aspects, the first substrate and said second substrate permit said chip to transmit or receive data at a transfer rate of at least 500 megabits per second and a frequency of at least 8 GHz. In some aspects, the RF amplifier chip is part of a satellite transmitter.

A modular and scalable system to transfer space articles between space orbits. In one embodiment, the system employs a rendezvous vehicle which docks with a space article in an initial orbit, the connected stack then docking with a locomotive vehicle which maneuvers to a targeted orbit where the space article is detached. In one feature, the rendezvous vehicle and locomotive vehicle use a common propellant and the space article is a satellite.

A radiator for a geostationary satellite is disclosed having a radiative panel perpendicular to a radiation axis, and pivoting relative to the radiation axis, a mounting foot for the panel, a motor which rotates the mounting foot about a rotation axis, the radiation axis and the rotation axis being tilted relative to each other by an angle corresponding to the angle of the satellite's orbital plane relative to the ecliptic plane of the planet, and a guidance system for the panel, limiting rotation of the panel about the rotation axis, including a connecting arm pivoting relative to the satellite about a first axis and relative to the panel about a second axis concurrent with the first axis at a point of intersection of all the axes.

A radiator for a geostationary satellite is disclosed having a radiative panel perpendicular to a radiation axis, and pivoting relative to the radiation axis, a mounting foot for the panel, a motor which rotates the mounting foot about a rotation axis, the radiation axis and the rotation axis being tilted relative to each other by an angle corresponding to the angle of the satellite's orbital plane relative to the ecliptic plane of the planet, and a guidance system for the panel, limiting rotation of the panel about the rotation axis, including a connecting arm pivoting relative to the satellite about a first axis and relative to the panel about a second axis concurrent with the first axis at a point of intersection of all the axes.

A system and method for installing, deploying, and recovering a plurality of spacecraft that provides an ease of use and structural stability, and facilitates a standardization of spacecraft design. In embodiments of this invention, threaded rods are arranged orthogonal to a surface of a baseplate, and each spacecraft includes a coupling mechanism that selectively engages or disengages each threaded rod. Each spacecraft is added to the stack by engaging its coupling mechanism and rotating the threaded rods while the preceding spacecraft on the stack disengage their coupling mechanisms, thereby enabling the spacecraft to travel along the threaded rods toward the baseplate. When all of the spacecraft are added to the stack, the stack is preloaded by rotating the treaded rods into a terminator component at the top of the stack while the coupling mechanisms in all of the spacecraft are disengaged. Spacecraft are deployed by reversing the process.

A system and method for installing, deploying, and recovering a plurality of spacecraft that provides an ease of use and structural stability, and facilitates a standardization of spacecraft design. In embodiments of this invention, threaded rods are arranged orthogonal to a surface of a baseplate, and each spacecraft includes a coupling mechanism that selectively engages or disengages each threaded rod. Each spacecraft is added to the stack by engaging its coupling mechanism and rotating the threaded rods while the preceding spacecraft on the stack disengage their coupling mechanisms, thereby enabling the spacecraft to travel along the threaded rods toward the baseplate. When all of the spacecraft are added to the stack, the stack is preloaded by rotating the treaded rods into a terminator component at the top of the stack while the coupling mechanisms in all of the spacecraft are disengaged. Spacecraft are deployed by reversing the process.

An object is to prevent satellites constituting a mega-constellation from remaining in outer space in large numbers after completing their missions. An artificial satellite includes a propulsion device. The artificial satellite has propellant to be used by the propulsion device, and the propellant is in an amount required for the artificial satellite to operate in orbit for a first period of L1 years, which is a satellite design life, and then enter the atmosphere within a period less than the first period of years after deorbit. A ground facility controls the artificial satellite so that the artificial satellite has the amount of propellant required to operate in orbit for the first period of L1 years, which is the satellite design life, and then enter the atmosphere within a period of less than the first period of L1 years after deorbit.

An object is to prevent satellites constituting a mega-constellation from remaining in outer space in large numbers after completing their missions. An artificial satellite includes a propulsion device. The artificial satellite has propellant to be used by the propulsion device, and the propellant is in an amount required for the artificial satellite to operate in orbit for a first period of L1 years, which is a satellite design life, and then enter the atmosphere within a period less than the first period of years after deorbit. A ground facility controls the artificial satellite so that the artificial satellite has the amount of propellant required to operate in orbit for the first period of L1 years, which is the satellite design life, and then enter the atmosphere within a period of less than the first period of L1 years after deorbit.