Establishing and growth of a lunar or planetary surface base involves continuing to use landing spacecraft as docked modules of the base for habitation and work. A first spacecraft is landed at a specified surface site then doubles as first module of the base. A second (and later third and subsequent) spacecraft is landed at the site a safe distance from the existing base modules then moved over the surface into a side-by-side position to dock with selected base modules. At least some of the landing, surface transport, and operational electric power is supplied by micro-fusion using ambient cosmic rays and muons interacting with deuterium-containing particle fuel material to generate energetic reaction products.

A matrix thruster that may be used to reposition and/or stabilize a CubeSAT satellite. The matrix thruster includes a conductive plate with an opening, a plurality of wires within the opening, a power supply electrically connected to the conductive plate or each of the plurality of wires via an inductor, and an electrical switch. The electrical switch creates a current change that creates an electric potential spike across the inductor. The electric potential spike across the inductor initiates an arc discharge between one of the wires and the conductive plate, which forms plasma that ejects cathode particles from the matrix thruster. Using multiple wires (e.g., four titanium wires) extends the lifetime of the thruster, as each wire restores an inter-electrode film needed for the other wires to continue generating plasma.

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.

Systems and methods for payload attachment are disclosed. An example payload tie-down system includes a retaining stud assembly and a payload tie-down assembly. The payload tie-down assembly includes a tie-down housing and a tie-down fork. The tie-down for is movably enclosed within the tie-down housing. Additionally, the tie-down fork is configured to move linearly along an axis of the tie-down housing. The linear movement is generated by a rotational input to the payload tie-down assembly. Additionally, a first end of the tie-down fork is angled along at least a first axis.

Methods and systems for implementing a Jupiter aerospace mission can enable delivery of a science payload to a Jupiter orbit on a direct Earth-to-Jupiter trajectory. Solar power and use of avionics also allow a fast assembly, integration, and test process compared to past outer Solar System missions. The spacecraft can include an aerodynamic forebody, a shell, and a thermal protection system. The spacecraft can manage the radiation environment, generate solar power, and return data to Earth with a robust radio-frequency (RF) amplification and antenna gain.

A system for detecting objects in space comprises an array of satellite nodes. The array of satellite nodes comprise at least one transmitter module for transmitting an electromagnetic signal, and a plurality of receiver modules for receiving diffractions from electromagnetic waves scattered from objects in space. The system comprises a control module for focussing the plurality of receiver modules to receive diffractions from a focussed virtual aperture in space.

A communications system includes cellular devices and cellular base stations in communication with the cellular devices in a first frequency band. A non-geostationary satellite may include sensing circuitry operable in a second frequency band susceptible to interference from the first frequency band. Each cellular base station may include a controller and a transceiver cooperating therewith. The controller may be configured to store satellite path data for the non-geostationary satellite, determine when the satellite path data indicates interference would otherwise be experienced by the non-geostationary satellite, and implement an interference mitigation action in cooperation with associated cellular devices based upon the satellite path data indicating interference would otherwise be experienced by the non-geostationary satellite.

A non-geostationary satellite system and method for weather and climate monitoring, communications applications, scientific research, and similar tasks. The satellite system provides global coverage using a constellation of six satellites in two orthogonal, 24 sidereal hour orbits (geosynchronous) with inclinations of 70 to 90, and eccentricities of 0.275-0.45. By placing three of the satellites in a first orbit with an apogee over the north pole, and three of the satellites in a second, orthogonal orbit with an apogee over the south pole, global coverage may be obtained. As well, the satellites in these orbits avoid most of the Van Allen Belts.

In the system of the system of the present invention, the detection processing device detects whether there is any target orbiting craft; the acquisition processing device acquires a surface image of said solar panels after detecting the target orbiting craft; the shadow feature data processing device analyzes the acquired surface image of solar panels to detect whether there is any shadow feature data; the trigger processing device triggers starting power ensuring control for the environment control and life support system after detecting the shadow feature data; the matching processing device matches said shadow feature data with the above-mentioned preset shadow feature data cast by the target orbiting craft on said solar panels, and if they match each other, continue to power said experiment cabin to prevent data loss of the experiment cabin, which can effectively avoid unnecessary experimental data loss of the space station.

A collision-avoidance propulsion system and method for orbiting satellites and other spacecraft takes advantage of ambient cosmic rays in space to catalyze micro-fusion events via particle-target fusion and muon-catalyzed fusion processes, using the reaction products to produce thrust upon orbiting satellites and other spacecraft. A supply of deuterium-containing particle fuel material is propelled in a specified direction of the spacecraft in response to indication of a potential collision with another space object (e.g. orbiting debris). In one embodiment, this may be performed by propellant gas expelling the fuel material through conduits to specified ports on the exterior of the spacecraft. The propelled material interacts with the ambient cosmic rays and muon generated from those cosmic rays to induce micro-fusion. A portion of the energetic reaction products (e.g. alpha particles) are received upon the spacecraft to alter its trajectory in a manner that avoids the potential collision.