Abstract: A turbine engine has: a compressor; a combustor; a turbine; a gaspath passing downstream from the compressor through the combustor and then through the turbine; a fuel source; a fuel flowpath from the fuel source to the combustor; and a heat exchanger for transferring heat from the gaspath to the fuel flowpath The heat exchanger has: an inner wall in heat transfer relation with the gaspath; an outer wall; tubes between the inner wall and the outer wall bounding respective segments of the fuel flowpath; and a heat transfer fluid between the inner wall and the outer wall and in heat transfer relation with the tubes and the inner wall.

Abstract: Various embodiments that pertain to fuel processing are described. A fuel processor can produce an endothermic reaction that cools a substance and produces a processed fuel from a raw fuel. A generator can employ the processed fuel to produce an electricity. The generator can supply the electricity to a load that uses the electricity to function. The load can become hot due to its functioning and can benefit from being cooled. The substance cooled by the fuel processor can cool load and in the process the substance can rise in temperature. This warmer substance can be transferred to the fuel processor to be cooled again and this cycle can continue. Further, the fuel processor can use the warmer substance to achieve the endothermic reaction.

Abstract: The present disclosure discloses a propulsion method based on liquid carbon dioxide phase change and a propulsion device. The method includes the following steps of: accommodating carbon dioxide in a thermally insulated container in a liquid phase form; transiently heating to convert the carbon dioxide from a liquid phase to a gas phase; and jetting carbon dioxide gas after the phase change in a predetermined direction by a predetermined jet-out amount so as to obtain a propulsion force.

Abstract: An integrated propulsion system for hybrid rockets includes an oxidizer tank, a rocket engine, a pressurization device, a pressurization device and an oxidizer pipe and valve unit. The rocket engine is disposed within the oxidizer tank partially and located on a first side of the oxidizer tank. The pressurization device is disposed, at least in part, within the oxidizer tank, is located on a second side of the oxidizer tank opposite to the first side of the oxidizer tank, and is configured to regulate an overall pressure level within the oxidizer tank. The oxidizer pipe and valve unit is connected to the oxidizer tank and the rocket engine, and is configured to control feeding of an oxidizer from the oxidizer tank into the rocket engine.

Abstract: A device includes a heat exchanger having one end connected to an air line through which air flows, and the other end connected to a hydrogen line through which liquid-state hydrogen flows. The heat exchanger is configured to produce liquid-state air as the air and the liquid-state hydrogen exchange heat with each other. The device also includes an air storage container connected to the heat exchanger via the air line and configured to store the liquid-state air discharged from the heat exchanger, and an evaporator connected to the air storage container via the air line and configured to evaporate the liquid-state air, supplied from the air storage container, through heat exchange. The device additionally includes a power generator configured to receive the air, discharged from the evaporator, via the air line, thereby producing electrical power.

Abstract: Nozzle of a spacecraft engine, comprising a nozzle body extending along a main direction, from a proximal end secured to the engine of the spacecraft, and a free distal end, said nozzle body having a plurality of stiffeners extending from an outer surface of the nozzle body, radially with respect to the main direction, said nozzle comprising a thermal insulation system, comprising at least one strip of insulating material disposed so as to surround the outer surface of the nozzle body over at least part of the height of said nozzle body, said thermal insulation system comprising a plurality of holding elements, each holding element being positioned so as to surround said strips of insulating material, and being disposed between two stiffeners of the nozzle body.

Abstract: In accordance with at least on aspect of this disclosure, there is provided a hydrogen fuel system for aircraft. The hydrogen fuel system includes a gas turbine engine and a fuel feed conduit. The fuel feed conduit is defined at least in part by, in fluid series, a liquid hydrogen tank fluidly connected to a combustor of the gas turbine engine, a liquid hydrogen pump to drive fuel to the combustor of the gas turbine engine, an evaporator, and an electric heat source in thermal communication with the evaporator to add heat into a flow of hydrogen passing through the evaporator. In embodiments, the electric energy source associated with the electric heat source to power the electric heat source.

Abstract: 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.

Abstract: An air-breathing turbojet engine for a hypersonic vehicle is shown. The engine comprises a pump for pumping a cryogenic fuel, an inlet configured to compress inlet air by one or more shocks, a cooler to cool the compressed inlet air using the cryogenic fuel, and a turbo-compressor to compress the air further. A precooler cools the compressed inlet air using compressed cooled air from the turbo-compressor. A combustor receives compressed cooled air from the turbo-compressor and a first portion of the cryogenic fuel for combustion. A first turbine expands and is driven by combustion products, and a second turbine expands and is driven by a second portion of the cryogenic fuel. The first turbine and the second turbine drive the turbo-compressor via a shaft. An afterburner receives combustion products from the first turbine and the second portion of the cryogenic fuel from the second turbine for combustion therein.

Abstract: An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.

Abstract: 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.

Abstract: The invention relates to a thrust chamber device comprising a thrust chamber with a thrust space having a first portion, a second portion adjacent thereto, and a third portion adjacent to the second portion, the thrust space being delimited in all three portions by an outer nozzle wall having an outer thrust space surface, which outer thrust space surface tapers in the first and second portion toward the third portion, widens in the third portion away from the second portion, and has a narrowest point at the transition from the second portion to the third portion, the first portion being delimited by an inner nozzle wall with an inner thrust space surface, which tapers toward the second portion, an annular combustion chamber being formed between the inner thrust space surface and the outer thrust space surface and extending over the first portion.

Abstract: A combustion chamber structure for a rocket engine includes a hot gas wall (12) that surrounds a combustion chamber (40) and has a plurality of first coolant channels (50) and a plurality of second coolant channels (52). The plurality of first (50) and second (52) coolant channels extend from a first longitudinal end (16) of the hot gas wall (12) to a second longitudinal end (18) of the hot gas wall (12) opposite to the first longitudinal end (16). The combustion chamber structure (10) further comprises a first manifold (20) forming a first coolant chamber (30) and a second manifold (22) forming a second coolant chamber (32) being fluidly separated from the first coolant chamber (30). The first (20) and second (22) manifolds are provided at the first longitudinal end (16) of the hot gas wall (12) and extend in a circumferential direction of the hot gas wall (12).

Abstract: A liquid rocket engine integrates tap-off openings at a combustion chamber wall to direct exhaust from the combustion chamber to a tap-off manifold that provides the exhaust to one or more auxiliary systems, such as a turbopump that pumps oxygen and/or fuel into the combustion chamber. The tap-off opening passes through a fuel channel formed in that combustion chamber exterior wall and receives fuel through a fuel opening that interfaces the fuel channel and tap-off opening. The tap-off manifold nests within a fuel manifold for thermal management. The fuel channel directs fuel into the combustion chamber through fuel port openings formed in the combustion chamber, the fuel port openings located closer to a headend of the combustion chamber than the tap-off openings.

Abstract: A swirl preburner having a first swirl core defining a first swirl chamber having a first swirl chamber first end and a first swirl chamber second end. The swirl preburner further includes a second swirl core defining a second swirl chamber having a second swirl chamber first end and a second swirl chamber second end. The swirl preburner also includes a mixing element defining a mixing chamber, the mixing chamber surrounding a portion of the first and second chamber at least including the first chamber second end and the second chamber second end.

Abstract: An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.

Abstract: An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.

Abstract: 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.

Abstract: Embodiments of the present invention generally relate to a vapor retention device and methods of using a vapor retention device to manage propellant for upper stage space vehicles. The use of a vapor retention device, in combination with controlled acceleration, drives liquid propellant from a propellant supply line communicating with an upper stage main engine back into a propellant tank and establishes an insulating liquid/gas propellant interface that prevents the exchange of gaseous propellant across the interface.

Abstract: An air-breathing rocket engine in certain embodiments comprises an hourglass-shaped outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake that terminates in a floor and an inner front wall that forms a first circumferential gap between the inner front wall and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector polis and one or more ignition ports are situated at the front end of the second circumferential gap.

Abstract: An air-breathing rocket engine with an hourglass-shaped outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake that terminates in a floor and an inner front wall that forms a first circumferential gap between the inner front wall and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.

Abstract: A combustor and a method for reducing a temperature gradient of a combustor component are provided. The combustor includes a coating applied to at least a portion thereof with the coating serving to alter the emissivity of the at least a portion to which it is applied. The method includes applying a coating on at least one of a combustor liner and a flow sleeve, wherein the coating alters the emissivity exhibited where applied.

Abstract: The present invention relates to a reactor for the decomposition of ammonium dinitramide-based liquid monopropellants into hot, combustible gases for combustion in a combustion chamber, and a rocket engine or thruster comprising such reactor, which reactor further comprises an inner reactor housing accommodating a heat bed and a catalyst bed, and separating the heat bed and catalyst bed from contact with the inner surface of the reactor housing.

Abstract: A system and methods are provided for combining systems of an upper stage space launch vehicle for enhancing the operation of the space vehicle. Hydrogen and oxygen already on board as propellant for the upper stage rockets is also used for other upper stage functions to include propellant tank pressurization, attitude control, vehicle settling, and electrical requirements. Specifically, gases from the propellant tanks, instead of being dumped overboard, are used as fuel and oxidizer to power an internal combustion engine that produces mechanical power for driving other elements including a starter/generator for generation of electrical current, mechanical power for fluid pumps, and other uses. The exhaust gas from the internal combustion engine is also used directly in one or more vehicle settling thrusters. Accumulators which store the waste ullage gases are pressurized and provide pressurization control for the propellant tanks.

Abstract: A propellant tank for storing a liquid propellant A and supplying vapor produced by evaporation of part of the liquid propellant A to an external location comprises a tank body for storing the liquid propellant A, a mesh member arranged inside the tank body to cover a liquid surface of the liquid propellant A to divide an interior of the tank body into a liquid propellant storing area LA and a gas storing area GA by utilizing surface tension of the liquid propellant, and a heater arranged to a gas storing area GA side of the tank body to keep the gas storing area GA at higher temperature than temperature in the liquid propellant storing area LA. The tank body has a propellant inlet open into the liquid propellant storing area LA and a gas outlet open into the gas storing area GA.

Abstract: A device heating a fluid and usable in a rocket launcher to pressurize a liquefied propellant. The device includes a first burner performing first combustion between a limiting propellant and an excess propellant; a first heat exchanger in which first burnt gas from the first combustion transfers heat to the fluid; at least one second burner into which both the first burnt gas and some limiting propellant are injected to perform second combustion between the limiting propellant and at least a portion of unburnt excess propellant present in the first burnt gas. The second burnt gas from the second combustion flows through a second heat exchanger to transfer heat to the fluid. Burnt gas from each combustion flows in respective burnt gas tubes within a common overall heat exchanger including the heat exchange units, the gas transferring heat to the fluid, the fluid flowing between the burnt gas tubes.

Abstract: A system and method of cooling a rocket motor component includes injecting a high pressure liquid coolant through an injector nozzle into a cooling chamber. The cooling chamber having a pressure lower than the high pressure liquid coolant. The liquid coolant flashes into a saturated liquid-vapor coolant mixture in the cooling chamber. The saturated liquid-vapor coolant mixture is at equilibrium at the lower pressure of the cooling chamber. Heat from the rocket motor component to be cooled is absorbed by the coolant. A portion of the liquid portion of the saturated liquid-vapor coolant mixture is converted into gas phase, the converted portion being less than 100% of the coolant. A portion of the coolant is released from the cooling chamber and the coolant in the cooling chamber is dynamically maintained at less than 100% gas phase of the coolant as the thrust and heat generated by the rocket motor varies.

Abstract: Disclosed is a liquid-fuel storage vessel for use in a vapor jet system to store liquid fuel, wherein the vapor jet system is adapted to jet the fuel in a state after being vaporized inside the liquid-fuel storage vessel, outside the liquid-fuel storage vessel, to obtain a thrust. The liquid-fuel storage vessel comprises: a hollow tank for storing the liquid fuel, wherein the tank has an ejection port for ejecting the vaporized fuel, outside the liquid-fuel storage vessel therethrough; a heating device for heating the tank; and a porous metal formed to have a plurality of interconnected cells and provided inside the tank, wherein at least a part of the liquid fuel is held in the cells of the porous metal, and heat energy given from the heating device to the tank is transferred to the liquid fuel through the porous metal to cause vaporization of at least a part of the liquid fuel. The liquid-fuel storage vessel of the present invention can obtain a stable thrust level and ensure spacecraft attitude control.

Abstract: In an aspect of the invention, a system for rocket propulsion includes a heater operable to generate thermal energy from energy supplied from a non-chemical energy source, and to supply the thermal energy to a non-cryogenic fuel to thermally decompose the fuel into components that include at least a first component and a second component. The rocket propulsion system also includes a combustion chamber and a nozzle. The combustion chamber is operable to receive an oxidizer and at least a portion of the thermally decomposed fuel, and allow the two to combust. The nozzle generates thrust by directing the products of the combustion out of the system.

Abstract: A component configured for being subjected to a high thermal load during operation includes a wall structure with cooling channels adapted for handling a coolant flow. At least one first cooling channel is adapted to convey the coolant from a first portion of the component to a second portion of the component. At least one second cooling channel in the second portion is closed so that the coolant is at least substantially-prevented from entering the closed second cooling channel from a cooling channel in the first portion.

Abstract: A dual fuel propulsion system comprising a gas turbine engine configured to generate a propulsive thrust using a cryogenic liquid fuel.

Abstract: The subject of the invention is a device for driving a pump of a rocket engine of a space vehicle that comprises an air-breathing internal combustion engine running on an oxidizer/fuel mixture of the air/hydrocarbon type, wherein the supply of oxidant and fuel is provided by tanks and a circuit that are separate from the propellant tanks of the rocket engine. The invention applies to a device for fueling a rocket engine that includes at least two pumps, each driven by the device of the invention, and means for controlling the internal combustion engines for driving the pumps, adapted for independently varying the operating parameters of these engines so as to independently adjust the rotation speeds of the pumps.

Abstract: An expander cycle rocket engine includes secondary turbopump to further pressurize a gaseous fuel discharged from a primary turbine prior to entering the combustion chamber. The secondary turbopump is driven by fuel bled off from the primary fuel pump. A gaseous fuel that is heated from passing around the nozzle that is passed through the primary turbine to drive the primary fuel and oxidizer pumps is then passed through the secondary turbine to drive the secondary fuel compressor. With the secondary turbopump used in the Johnson-Sexton cycle engine, a thrust produced by the expander cycle rocket engine is greater than those obtained by prior art expander cycle rocket engines due to the square-cube rule.

Abstract: A method of cooling a rocket engine that includes removal of a gaseous propellant from a tank containing liquid propellant, thereby cooling the propellant in the tank, cooling a rocket engine with a coolant, thereby heating the coolant, and transferring heat from the heated coolant into the propellant in the tank, thereby heating the propellant in the tank and maintaining the pressure in the tank. In certain embodiments, the coolant is injected into the rocket engine after transferring heat from the rocket engine to the propellant in the tank.

Abstract: An apparatus and method to electrocatalytically induce the decomposition of a liquid propellant for propulsion. This is accomplished by providing a reaction chamber filled with catalyst that has a first electrode, such as near the center, and the other electrode disposed away from the first electrode. A charge source that is capable of applying voltage is connected to the electrodes. When the reaction chamber is dosed with a propellant such as an aqueous based amine propellant, such as HAN, an electric current flows between the electrodes and the propellant. This initiates the decomposition of the propellant which is driven to completion by the catalyst bed in order to be used for power generation.

Abstract: One embodiment is directed to a method for testing an aircraft prior to flight. The method includes receiving a user signal from a pre-flight test input source, the user signal indicating that a pilot of the aircraft has directed engine control circuitry, which is arranged to electronically control operation of a set of piston engines of the aircraft during flight, to begin testing the aircraft in an automated manner. The method includes, in response to the user signal, conducting a pre-flight test of the aircraft from the engine control circuitry. The method includes, upon completion of the pre-flight test, outputting a result of the pre-flight test from the engine control circuitry.

Abstract: In some embodiments a propulsion system includes a thrust chamber having an inside wall, an expansion nozzle mounted to the thrust chamber and having an interior and having an exterior, a main propellant injector mounted to the thrust chamber to inject a fluid in the interior of the thrust chamber, the fluid comprising oxidizer, fuel and internal film coolant, the internal film coolant ranging from about 1% to about 5% of the fluid, limited coolant tubing circumscribing the exterior of the expansion nozzle to circulate an external coolant, and an injector mounted to the expansion nozzle to inject the external coolant in the interior of the expansion nozzle, the external convective coolant about 2.5% of the fluid. The system operates at lower temperatures while having conventional amounts of thrust, in which the thrust chamber can be made of thin walls of lower cost conventional metals with simple coolant tube construction.

Abstract: An expander heat exchanger cycle system provides a fuel or oxidizer routed through a heat exchanger to cool and condense a coolant/turbine drive fluid.

Abstract: An augmented expander cycle rocket engine includes first and second turbopumps for respectively pumping fuel and oxidizer. A gas-generator receives a first portion of fuel output from the first turbopump and a first portion of oxidizer output from the second turbopump to ignite and discharge heated gas. A heat exchanger close-coupled to the gas-generator receives in a first conduit the discharged heated gas, and transfers heat to an adjacent second conduit carrying fuel exiting the cooling passages of a primary combustion chamber. Heat is transferred to the fuel passing through the cooling passages. The heated fuel enters the second conduit of the heat exchanger to absorb more heat from the first conduit, and then flows to drive a turbine of one or both of the turbopumps. The arrangement prevents the turbopumps exposure to combusted gas that could freeze in the turbomachinery and cause catastrophic failure upon attempted engine restart.

Abstract: A propulsion system for a rocket engine and a method of cooling a rocket engine includes a propellant tank fluidically coupled to the rocket engine to hold a pressurized propellant, a coolant tank to hold a coolant, a first heat exchanger thermally coupled to the rocket engine and fluidically coupled to the coolant tank, a second heat exchanger thermally coupled to the propellant tank and fluidically coupled to the first heat exchanger, and a third heat exchanger disposed inside the propellant tank to thermally couple a propellant withdrawn from the tank for combustion to a propellant disposed inside the tank. The coolant flows from the coolant tank to the first heat exchanger and through the first heat exchanger to cool the rocket engine. The propellant withdrawn from the propellant tank receives heat from the propellant disposed inside the tank through the third heat exchanger to convert to a gaseous propellant when withdrawn from the propellant tank as a liquid propellant.

Abstract: A system and method of driving a fuel pump and/or an oxidizer pump in a propulsion system includes pumping an oxidizer from an oxidizer supply to a rocket engine with the oxidizer pump. A first portion of the pumped oxidizer is used for combustion in the rocket engine. The heat from the combustion of the first portion of the oxidizer is then transferred to a second portion of the pumped oxidizer to convert the second portion of the oxidizer to a super-heated gaseous oxidizer. The super-heated gaseous oxidizer operates a motor, which drives the oxidizer pump and/or the fuel pump.

Abstract: A propulsion system for a rocket engine and a method of cooling a rocket engine includes a propellant tank fluidically coupled to the rocket engine to hold a pressurized propellant, a coolant tank to hold a coolant, a first heat exchanger thermally coupled to the rocket engine and fluidically coupled to the coolant tank, a second heat exchanger thermally coupled to the propellant tank and fluidically coupled to the first heat exchanger, and a third heat exchanger disposed inside the propellant tank to thermally couple a propellant withdrawn from the tank for combustion to a propellant disposed inside the tank. The coolant flows from the coolant tank to the first heat exchanger and through the first heat exchanger to cool the rocket engine. The propellant withdrawn from the propellant tank receives heat from the propellant disposed inside the tank through the third heat exchanger to convert to a gaseous propellant when withdrawn from the propellant tank as a liquid propellant.

Abstract: An augmented expander cycle rocket engine includes first and second turbopumps for respectively pumping fuel and oxidizer. A gas-generator receives a first portion of fuel output from the first turbopump and a first portion of oxidizer output from the second turbopump to ignite and discharge heated gas. A heat exchanger close-coupled to the gas-generator receives in a first conduit the discharged heated gas, and transfers heat to an adjacent second conduit carrying fuel exiting the cooling passages of a primary combustion chamber. Heat is transferred to the fuel passing through the cooling passages. The heated fuel enters the second conduit of the heat exchanger to absorb more heat from the first conduit, and then flows to drive a turbine of one or both of the turbopumps. The arrangement prevents the turbopumps exposure to combusted gas that could freeze in the turbomachinery and cause catastrophic failure upon attempted engine restart.

Abstract: A dual expander cycle (DEC) rocket engine with an intermediate closed-cycle heat exchanger is provided. A conventional DEC rocket engine has a closed-cycle heat exchanger terhamlly coupled thereto. The heat exchanger utilizes heat extracted from the engine's fuel circuit to drive the engine's oxidizer turbomachinery.

Abstract: A process and a system for delivering a propellant combination to a rocket engine is described. The process comprises the steps of providing a flow of a hydrocarbon propellant fuel, raising the pressure of the hydrocarbon propellant fuel, cracking the hydrocarbon propellant fuel in a cracker, introducing the cracked hydrocarbon propellant fuel into a combustion chamber of the rocket engine and introducing an oxidizer into the combustion chamber. A system for performing the process is also described.

Abstract: A system for providing an oxidizer and a fuel to a rocket engine is provided. The system includes a fuel supply system. The fuel supply system includes a fuel pump that pumps fuel to the rocket engine. The system includes a coolant supply system that supplies a coolant to the rocket engine, and a power plant that powers at least one of the fuel supply system and the coolant supply system. The power plant is powered by energy received from the coolant system.

Abstract: A system for cooling a portion of a rocket engine with an inert compound and transferring the thermal energy from the inert compound to the propellants. The energy absorbed by the coolant is used also to power the turbines which powers the pumps that pump the fuel, the oxidizer, and the coolant. Additionally, the coolant, which is an inert compound, is used to separate the oxidizer and the fuel before the oxidizer and the fuel enter the combustion chamber eliminating the need for a complex inert turbo pump seal package. The systems which pump or comprise the coolant physically separate the propellants before the propellants enter the rocket engine. The coolant remains substantially unconsumed in this cycle.

Abstract: The present invention relates to a reactor for the decomposition of ammonium dinitramide-based liquid monopropellants into hot, combustible gases for combustion in a combustion chamber, and more particularly a rocket engine or thruster comprising such reactor and a combustion chamber. The invention also relates to a process for the decompostion of ammonium dinitramide-based liquid monopropellants.

Abstract: A method for forming a coolant system for a rocket engine combustion chamber is provided. The method comprises the steps of providing a plurality of tubes formed and shaped into the profile of a nozzle with each of the tubes having a constantly expanding cross section in an upper chamber area, providing an inlet manifold and an exit manifold with a plurality of holes for receiving an end of each tube, inserting a brazing preform into each hole, inserting a first end of each tube into the inlet manifold and a second end of each tube into the outlet manifold so that the first end is surrounded by a first brazing preform and the second end is surrounded by a second brazing preform, and brazing the inlet and outlet manifolds to the tubes. The brazing step forms a series of brazed joints between the tubes and the manifolds. The method further includes the steps of forming a layer of coating material on exposed portions of the tubes and forming a single piece jacket construction around the tubes.

Abstract: A gas generator or rocket engine includes: a first storage tank configured to contain a high-pressure liquid propellant; a nozzle connected to the first storage tank; a heat source connected between the first storage tank and the nozzle and configured to add heat to the high-pressure liquid propellant at a heat transfer rate to substantially gasify the high-pressure liquid propellant, where the nozzle is configured to expel and expand the substantially gasified high-pressure liquid propellant, and where the gas generator is configured so that an expanded temperature of the substantially gasified high-pressure liquid propellant after being expanded by the nozzle is in the range ?50° C. to 100° C., preferably 0° C. to 50° C.; and a controller connected to the heat source and configured to adjust the heat transfer rate.