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.

Space vehicles with paraglider re-entry, and associated systems and methods are disclosed. A representative system includes a re-useable space vehicle, a collapsible, deployable and re-stowable re-entry heat shield carried by the space vehicle, and a collapsible, deployable and re-stowable flexible paraglider wing also carried by the space vehicle. The space vehicle can accordingly carry out repeated space-based missions, and can be refurbished and replenished on Earth and/or at an orbiting dock between missions.

A method and a device for deorbiting artificial satellites with multilayer insulation arranges a deorbiting device beneath it. The device includes a spring element, a bonding agent, and a heating element designed to melt the bonding agent, which at least partially detaches at least one layer of the multilayer insulation from the satellite when a signal is transmitted to the satellite by rolling up or splaying out in the manner of a flap, enlarging its cross-sectional area. The partial detachment exposes the underlying satellite structure to environmental influences. This exposure accelerates the self-disintegration of the artificial satellite, reduces the mass, and increases the ballistic coefficient throughout the deorbiting period. This area enlargement and mass reduction reduce energy, resulting in re-entry into Earth's atmosphere. A number of layers can be arranged so that, regardless of the satellite rotation, at least one surface is always directed against the aerodynamic flow.

A non-tubular rocket stage landing apparatus for more reliably landing and returning a non-tubular rocket stage for reuse in subsequent main missions. The non-tubular rocket stage landing apparatus comprises a platform, a dampening and cushioning device, an exhaust diffuser device, a hollow shell, and a rotational device for angular aligning the hollow shell with the pass-through non-circular circumference of a vertically landing non-tubular rocket stage.

A stage of a rocket is disclosed. The rocket stage may include: a body, and a plurality of foldable propulsion units spaced around a circumference of the body, where each propulsion unit comprises: a folding beam; at least one motor mounted to the folding beam, and at least one propeller mounted to the at least one motor, configured to generate a thrust to propel the rocket.

A vehicle configured for in-space and atmospheric reentry operations includes a pressure-fed propulsion engine that uses cryogenic fuel and cryogenic oxidizer as propellants. In some embodiments, the vehicle includes a first conduit configured to provide a first propellant to the pressure-fed propulsion engine, a second conduit configured to provide a second propellant to a heat shield heat exchanger disposed relative to a heat shield wall defining an outer surface of the vehicle, and a pump located along the second conduit. In such embodiments, the first conduit is configured to provide the first propellant to the pressure-fed propulsion engine at a first mass flow rate, the second conduit and the pump are configured to provide the second propellant to the heat shield heat exchanger at a second mass flow rate, and the second mass flow rate is substantially less than the first mass flow rate. A related method is also disclosed.

A non-tubular rocket stage landing apparatus for more reliably landing and returning a non-tubular rocket stage for reuse in subsequent main missions. The non-tubular rocket stage landing apparatus comprises a platform, a dampening and cushioning device, an exhaust diffuser device, a hollow shell, and a rotational device for angular aligning the hollow shell with the pass-through non-circular circumference of a vertically landing non-tubular rocket stage.

A launch vehicle with a rocket body having a longitudinal axis which has at least one propulsion stage which can be driven by a reaction propulsion system acting predominantly parallel to the longitudinal axis, wherein the launch vehicle is provided with a plurality of rotors which can be driven by means of a respective rotor drive and whose respective rotor axis is aligned substantially parallel to the longitudinal axis of the rocket body, is characterized in that a separate takeoff stage is provided which is coupled or can be coupled to the rocket body and/or the propulsion stage, which is coupled to or can be coupled to and decoupled from the separate takeoff stage, which has the plurality of outer rotors, in that the outer rotors are arranged in the manner of a multi-copter radially outside the rocket body and surrounding the rocket body.

Systems and methods for a parachute or parachute system are described. The parachute or parachute system includes two or more panels and includes at least one band which may be in a lateral direction relative to the two or more panels. The at least one band may be secured at different points which are lateral mid-points in individual contiguous ones of the two or more panels. Further, a first distance between an individual set of points of the different points may be shorter than twice a second distance. The second distance is a lateral distance from at least one point of the individual set of points to a seam of one panel of the two or more panels having the at least one point.

A carrier rocket system, with a carrier rocket, together with a launch assistance unit, which is detachably connected, or connectable, to the carrier rocket. The carrier rocket system is set up to take off in a horizontal launch via the launch assistance unit when thrust is generated by at least one propulsion unit of the carrier rocket. The launch assistance unit can be detached from the onward-flying carrier rocket after supersonic velocity has been reached, and is set up to fly back to the earth's surface after its detachment from the carrier rocket. The launch assistance unit preferably provides all the necessary structural supports to enable a carrier rocket that has been designed for a vertical launch to use one or more of its essential components without structural modifications. Also disclosed are a take-off method for a carrier rocket system, and a manufacturing method for a carrier rocket system.