A test article described herein enables ground-based arc jet testing to investigate RF blackout mitigation using electrophilic gas injection upstream of an antenna. The article can be scaled up to actual flight vehicles, thereby allowing reentry vehicles to be in constant, or near constant, communication during atmospheric reentry. Plasma blackout mitigation is an enabling technology that is required to advance hypersonic flight. Example articles include an integral structure that supports a nozzle, piping for gas connected to the nozzle and an RF window. An ablator can be attached to the structure. The ablator can include a graphite ablator and an insulator. A flight vehicle can include an antenna and such an article.

A test article described herein enables ground-based arc jet testing to investigate RF blackout mitigation using electrophilic gas injection upstream of an antenna. The article can be scaled up to actual flight vehicles, thereby allowing reentry vehicles to be in constant, or near constant, communication during atmospheric reentry. Plasma blackout mitigation is an enabling technology that is required to advance hypersonic flight. Example articles include an integral structure that supports a nozzle, piping for gas connected to the nozzle and an RF window. An ablator can be attached to the structure. The ablator can include a graphite ablator and an insulator. A flight vehicle can include an antenna and such an article.

Cell-based systems may interlock in a reconfigurable configuration to support a mission. Space systems, for example, of a relatively large size may be assembled using an ensemble of individual “cells”, which are individual space vehicles. The cells may be held together via magnets, electromagnets, mechanical interlocks, etc. The topology or shape of the joined cells may be altered by cells hopping, rotating, or “rolling” along the joint ensemble. The cells may be multifunctional, mass producible units. Rotation of cell faces, or of components within cells, may change the functionality of the cell. The cell maybe collapsible for stowage or during launch.

A satellite rescue system (SRS) (1) for rescue and recertification of dormant satellites, said SRS having a thruster end (13) with a primary propulsion nozzle (11) and maneuvering thrusters (12) and a satellite connection end (8) with a body (15) between both ends. The satellite connection end of the SRS has an interface ring (14) with clinch clamps (4) that securely attach to a ring (3) on the rescued satellite. An umbilical connector (7) on the satellite connecting end of the SRS provides power and data to the rescued satellite.

A satellite rescue system (SRS) (1) for rescue and recertification of dormant satellites, said SRS having a thruster end (13) with a primary propulsion nozzle (11) and maneuvering thrusters (12) and a satellite connection end (8) with a body (15) between both ends. The satellite connection end of the SRS has an interface ring (14) with clinch clamps (4) that securely attach to a ring (3) on the rescued satellite. An umbilical connector (7) on the satellite connecting end of the SRS provides power and data to the rescued satellite.

A space oven operates in microgravity environments by forcing convection towards the center through a unique heating element and airflow design. The space oven includes a tubular chamber, a heating rack, a heating system, a cooling system, a hatch, a user interface, a microcontroller, an enclosure, at least one first vent, at least one second vent and at least one temperature sensor. The tubular chamber is the cooking area. The heating rack holds consumables in place. The heating system heats up consumables. The cooling system prevents any overheating. The hatch closes off and allows access to the inside of the tubular chamber. The user interface allows a user to input commands. The microcontroller manages the electronic components. The enclosure protects the tubular chamber. The at least one first vent and the at least one second vent reduce pressure buildup. The at least one temperature sensor monitors the internal temperature.

A space oven operates in microgravity environments by forcing convection towards the center through a unique heating element and airflow design. The space oven includes a tubular chamber, a heating rack, a heating system, a cooling system, a hatch, a user interface, a microcontroller, an enclosure, at least one first vent, at least one second vent and at least one temperature sensor. The tubular chamber is the cooking area. The heating rack holds consumables in place. The heating system heats up consumables. The cooling system prevents any overheating. The hatch closes off and allows access to the inside of the tubular chamber. The user interface allows a user to input commands. The microcontroller manages the electronic components. The enclosure protects the tubular chamber. The at least one first vent and the at least one second vent reduce pressure buildup. The at least one temperature sensor monitors the internal temperature.

A method for estimating collision between a satellite in orbit and at least one piece of space debris having a time of closest approach to the satellite is disclosed including: obtaining the reference orbit of the satellite; determining an ephemeris of state transition data representative of the trajectory of the reference orbit; communicating the reference orbit and the ephemeris of state transition data to the satellite. The method includes the steps on board the satellite of: determining the true orbital position of the satellite; propagating the true orbit; calculating a probability of collision between the satellite and the piece of debris.

A trussed collapsible tubular mast includes a deformable beam having an extended state, a flattened state, and a rolled state, where a stiffness and strength of the deformable beam in the extended state is greater than a different stiffness and a different strength of the deformable beam in the flattened state. At least one collapsible tubular mast wall has a plurality of truss members of a first material having a first material thickness. At least one truss member is disposed substantially perpendicular to a longitudinal axis of the trussed collapsible tubular mast. Disposed between the truss members is a wall area of a second material thickness less thick than the first material thickness.

The disclosed method for a communication satellite may include (1) simultaneously generating a first transmission beam to a first ground station and a second transmission beam to each of a plurality of second ground stations in sequence according to a schedule, (2) simultaneously receiving a third transmission beam from the first ground station and a fourth transmission beam from each of the second ground stations in sequence according to the schedule, (3) forwarding first data received via the third transmission beam to each of the second ground stations via the second transmission beam, and (4) forwarding second data received via the fourth transmission beam from each of the second ground stations to the first ground station via the first transmission beam. Various other methods and systems are also disclosed.