Passive reactivity control technologies that enable reactivity control of a nuclear thermal propulsion (NTP) system with little to no active mechanical movement of circumferential control drums. By minimizing or eliminating the need for mechanical movement of the circumferential control drums during an NTP burn, the reactivity control technologies simplify controlling an NTP reactor and increase the overall performance of the NTP system. The reactivity control technologies mitigate and counteract the effects of xenon, the dominant fission product contributing to reactivity transients. Examples of reactivity control technologies include, employing burnable neutron poisons, tuning hydrogen pressure, adjusting wait time between burn cycles or merging burn cycles, and enhancement of temperature feedback mechanisms. The reactivity control technologies are applicable to low-enriched uranium NTP systems, including graphite composite fueled and tungsten ceramic and metal matrix (CERMET), or any moderated NTP system, such as highly-enriched uranium graphite composite NTP systems.

A rocket engine having a co-axial, bidirectional flow arrangement is described herein. The rocket engine receives fuel and an oxidizer into the rocket engine in a first direction, whereby a portion of the fuel is combusted in a pre-burner. The flow direction of the partially combusted fuel/oxidizer mixture is reversed, whereby the mixture is introduced into a combustion chamber. The fuel and oxidizer are combusted in the combustion chamber. The combustion products exit a throat and an expansion plenum in a direction similar to the first direction, whereby the combustion products exit a nozzle of the rocket engine, providing thrust.

According to one contemplated embodiment of the rocket engine invention, water is first pumped from a water tank through a rocket nozzle cooling heat exchanger wherein it is evaporated into said superheated steam. A generator supplies electricity to an electrolyzer that electrolyzes superheated steam into gaseous hydrogen and gaseous oxygen. The gaseous hydrogen and gaseous oxygen is employed for forming an annular curtain of secondary combustion in a divergent rocket engine. The secondary combustion gas surrounds a central thrust of combustion gas produced in an upstream combustion chamber by a primary injection of hydrogen/oxygen supplied from a liquid hydrogen tank and liquid oxygen tank. The rocket liquid hydrogen tank and liquid oxygen tank are pressurized by gaseous hydrogen and gaseous oxygen generated by the electrolyzer.

A propulsion assembly for a rocket includes a propellant tank configured to contain a propellant and an engine comprising a combustion chamber configured to subject the propellant to combustion and generate exhaust gases. The propulsion assembly further includes a supply circuit and an exhaust gas circuit. The supply circuit is disposed between the propellant tank and the combustion chamber, and the supply circuit is configured to supply the combustion chamber with the propellant. The exhaust gas circuit is disposed between the combustion chamber and the propellant tank, and the exhaust gas circuit is configured to convey at least part of the exhaust gases from the combustion chamber to the propellant tank to provide pressurization of the propellant tank.

Devices and methods of rocket propulsion are disclosed. In one aspect, a staged combustion liquid rocket engine with preburner and turbopump unit (TPU) integrated into the structure of the combustion chamber is described. An initial propellant mixture is combusted in a preburner combustion chamber formed as an annulus around a main combustion chamber, the combustion products from the preburner driving the turbine of the TPU and subsequently injected into the main combustion chamber for secondary combustion along with additional propellants, generating thrust through a supersonic nozzle. The preburner inner cylindrical wall is shared with the outer cylindrical wall of the engine's main combustion chamber and the turbine is axially aligned with the main combustion chamber. Liquid propellants supplied to the engine are utilized for regenerative cooling of the combustion chamber and preburner, where the liquid propellants are gasified in cooling manifolds before injection into the preburner and main combustion chamber.

Devices and methods of rocket propulsion are disclosed. In one aspect, a staged combustion liquid rocket engine with preburner and turbopump unit (TPU) integrated into the structure of the combustion chamber is described. An initial propellant mixture is combusted in a preburner combustion chamber formed as an annulus around a main combustion chamber, the combustion products from the preburner driving the turbine of the TPU and subsequently injected into the main combustion chamber for secondary combustion along with additional propellants, generating thrust through a supersonic nozzle. The preburner inner cylindrical wall is shared with the outer cylindrical wall of the engine's main combustion chamber and the turbine is axially aligned with the main combustion chamber. Liquid propellants supplied to the engine are utilized for regenerative cooling of the combustion chamber and preburner, where the liquid propellants are gasified in cooling manifolds before injection into the preburner and main combustion chamber.

The present disclosure relates to a rocket engine, and more particularly, a rocket engine with an integrated combustor head and turbopump in which a turbopump of the rocket engine is formed integrally with a combustor head.

A pyrotechnic device comprising a main pyrotechnic charge, a firing device for firing the main pyrotechnic charge, a discharge passage for discharging the gas generated by firing the main pyrotechnic charge, and an injector device configured to inject a cooling fluid into said gas discharge passage, so as to deliver gas, specifically for driving turbines, at temperatures that are relatively low, and a method of cooling gas generated by firing the main pyrotechnic charge by injecting the cooling fluid.

A pyrotechnic device comprising a main pyrotechnic charge, a firing device for firing the main pyrotechnic charge, a discharge passage for discharging the gas generated by firing the main pyrotechnic charge, and an injector device configured to inject a cooling fluid into said gas discharge passage, so as to deliver gas, specifically for driving turbines, at temperatures that are relatively low, and a method of cooling gas generated by firing the main pyrotechnic charge by injecting the cooling fluid.

A liquid propellant rocket engine includes a pump that is disposed along a central axis. The pump includes a purge system, a collection annulus in fluid communication with the purge system, and a drain. The collection annulus has an outer diameter wall, an inner diameter wall, and an end wall. The end wall defines an annular channel that has a channel depth that varies circumferentially. The drain opens to the collection annulus. At the drain, the annular channel has a lowest point at which the channel depth is maximum depth.