A system for launching aerospace payloads includes a wingless, unmanned modified lifting body spacecraft (100), with a payload compartment in the forward section of the spacecraft. The spacecraft is propelled by hybrid rockets clustered in the aft section of the spacecraft. Reaction control system (RCS) modules control the flight path and its associated avionics hardware and software. This system also includes a carrier aircraft (200) configured to air-launch the spacecraft. The carrier aircraft includes a flight operations control system, which monitors the spacecraft's payload and monitors and controls launch and flight operations of the spacecraft. A ground-based mission control system monitors and controls the spacecraft's payload and monitors and controls the launch and flight operations of the spacecraft.

Electrically operated propellant thrust assist supplements an airplane's takeoff, landing or inflight maneuvers. Unlike conventional SRM propellants, the burn rate of the electrically operated propellant can be varied via an electrical input and even extinguished by interrupting the electrical to control a secondary thrust profile (e.g., amplitude, transition rates) to fulfill the needs of a given takeoff, inflight or landing maneuver and provide a smooth transition in and out of the maneuver. Multiple pairs of fixed thrusters (opposite sides of the fuselage), a single pair of gimbaled thrusters or a hybrid of fixed and gimbaled thrusters may be configured to provide all such maneuvers. Flight control inputs are passed back and forth through an interface to enable the thrust assist.

A fuel grain for a rocket, the fuel grain having a plurality of layers of fuel grain material, each layer comprising a plurality of concentric circular structures of different diameter fused together to form a central opening therein, wherein the fuel grain material comprises an ignitable substance. The plurality of layers are stacked and joined securely to form a cylindrical fuel grain with the central opening of each one of the plurality of layers aligned to form a combustion unit extending axially through the fuel grain and bounded by a combustion surface, and wherein the fuel grain is configured to permit mixing of heterogenous materials to enhance thrust performance.

A cylindrically-shaped hybrid rocket engine solid fuel grain defines an axial combustion port. A fuel grain material comprises a compounded blend of thermoplastic fuel and aluminum. The fuel grain comprises fused stack layers, each layer comprising a plurality of fused abutting concentric beaded structures arrayed to define the combustion port; the port exhibits a rifling pattern or rifling inducing geometry along the port wall. When an oxidizer is introduced into the combustion port combustion occurs along the exposed port wall. Each beaded structure defines a geometry that increases the combustion surface area while inducing a vortex flow of oxidizer and fuel gas. As each layer ablates, an abutting layer exhibiting a similar geometry, is revealed, undergoes a gas phase change, and ablates. This process repeats and persists until oxidizer flow is terminated or the fuel grain material is exhausted. The fuel grain may be manufactured by an additive manufacturing process.

A hybrid rocket engine system has: an oxidizer tank containing a liquid oxidizer; a rocket engine having a combustion chamber operatively connected to the oxidizer tank; a solid propellant fuel within the combustion chamber; a nozzle fluidly connected to the combustion chamber, the nozzle having a convergent section and a divergent section downstream of the convergent section; and a thrust vector control device operatively connected to the divergent section of the nozzle and operable to inject a fluid through at least one aperture defined through the divergent section for controlling a direction of a thrust generated by the rocket engine.

A hybrid rocket engine system has: an oxidizer tank containing a liquid oxidizer; a rocket engine having a combustion chamber operatively connected to the oxidizer tank; a solid propellant fuel within the combustion chamber; a nozzle fluidly connected to the combustion chamber, the nozzle having a convergent section and a divergent section downstream of the convergent section; and a thrust vector control device operatively connected to the divergent section of the nozzle and operable to inject a fluid through at least one aperture defined through the divergent section for controlling a direction of a thrust generated by the rocket engine.

The invention relates to a method for deflecting space debris comprising steps of: launching (E2) a thruster (2) by means of a sounding rocket (1) at a target altitude close to that of the one or more debris to be deflected. generating (E3) by the thruster a gas cloud (G) above the sounding rocket.

A hybrid-like liquid fuel motor (the motor) may include a port surrounded by a wall. Surrounding the wall are a plurality of chambers and segmented walls to separate the chambers. In some instances, a single helix chamber may surround the wall, and may operate similar to that of a segmental chamber. During operation of the motor, gas flows from one end of the port to another end of the port. As the walls surrounding the port begin to disintegrate, liquid fuel within chambers begins to begin to mix with the flow of gas. As the segmented walls between the chambers begin to disintegrate, liquid from the other chambers begin to mix with the flow of gas, creating a metering of the liquid fuel.

The present invention, LB-1A, is a system for launching satellites, including small-sats, mini-sats, nano-sats, medical and scientific experiments, sub-orbital, orbital, and other aerospace payloads, which includes a modified existing carrier aircraft and a streamlined, unmanned rocket propelled lifting body spacecraft, air-launched from said carrier aircraft.

A method of making a multi-grained fuel grain for a rocket is disclosed, the method comprising the steps of using at least one nozzle to extrude a first propellant in an additive manufacturing process, the first propellant comprising a multi-grained fuel grain, the multi-grained fuel grain forming the at least one void, the at least one void facilitating variation in internal ballistics, forming sensors, said sensors permitting continuous monitoring and continuous modification such that a user controls the ballistics profile of a rocket motor, forming an electrically-controlled second propellant in contact with and operatively coupled to the sensors; and wherein the additive manufacturing process uses at least at least one nozzle to extrude raw materials.