Abstract: An oxygen-hydrogen pressurization system includes a cryogenic oxygen tank, cryogenic hydrogen tank, thermal switch, supercritical oxygen bottle, supercritical hydrogen bottle, and pressure management system and a thermodynamic vent system. The thermal switch permits heat to flow between hot and cool areas within the space vehicle to help facilitate pressure management within the cryogenic liquid oxygen tank and the cryogenic liquid hydrogen tank in conjunction with the higher pressure fluid from the supercritical oxygen tank and the fluid from the supercritical hydrogen tank and the added cooling from the pressure management system.

Abstract: Disclosed are materials of variable density or tiered porosity micro-fluidic porous media structures of sintered metal or other materials, and methods of making same. An embodiment discloses an aluminum porous media element of variable density having a tiered porosity micro-fluidic media structure. A method of making the aluminum porous media element disclosed herein includes mixing a binding agent with a metal powder to generate a first mixture, heating the first mixture to a sub metal sintering temperature to get a homogeneous composite of the metal powder and heating the homogeneous composite to a metal sintering temperature to sinter-bond the metal powder to get a porous media of first porosity.

Abstract: An injection head for the combustion chamber of a rock propulsion unit has a base area with a plurality of injection elements distributed on it. Propellant constituents can be injected from each outlet of the injection elements into the combustion chamber, generating hot gases by a mixing and combustion of the propellant constituents in the combustion chamber, which hot gas can be accelerated to a high velocity for generating thrust. The base area of the injection head had a contoured shape which prevents or reduces the formation of zones of a reduced pressure, particularly between directly adjacent outlets.

Abstract: A liquid propelled rocket engine comprising a unique pintle injector, radial injector plate and ablative insert to improve combustion stability and reduce harmonic disturbance within the chamber. The injector plate is generally in the shape of a ring surrounding a combination spud and pintle injector, the two allowing at least one liquid propellant and one liquid oxidizer to be injected into the chamber. The injector plate has an array of injectors for injecting liquid propellant into the chamber and a series of cavities within allowing the harmonic disturbance to be damped and providing a more stable combustion environment. The number of cavities and the unique pintle injector's length and width are chosen to reduce harmonic disturbance in the longitudinal, radial and lateral modes. A method for designing engine components and tuning harmonic frequencies within a combustion chamber is also disclosed.

Abstract: The present invention provides a heretofore-unknown use for zirconium nitride as a hydrogen peroxide compatible protective coating that was discovered to be useful to protect components that catalyze the decomposition of hydrogen peroxide or corrode when exposed to hydrogen peroxide. A zirconium nitride coating of the invention may be applied to a variety of substrates (e.g., metals) using art-recognized techniques, such as plasma vapor deposition. The present invention further provides components and articles of manufacture having hydrogen peroxide compatibility, particularly components for use in aerospace and industrial manufacturing applications. The zirconium nitride barrier coating of the invention provides protection from corrosion by reaction with hydrogen peroxide, as well as prevention of hydrogen peroxide decomposition.

Abstract: A propulsion system and method are described configured to exert a force upon a vehicle. The system includes a concentrated mass which may be discharged from the vehicle by a propellant typically via an ejection tube. The system is optimized so that the discharged concentrated-mass imparts a large impulse upon the vehicle. The system and method may be used to alter the momentum of vehicles for propulsion, attitude correction, vehicle separation and the like.

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 catalytically activated transient decomposition propulsion system provides thrust by decomposing flow controlled propellant in contact with a catalyzing agent using a fixed volume of liquid propellant that is placed in contact with the catalyst within the decomposition chamber by a calibrated flow control valve. After injecting the liquid propellant into the decomposition chamber, the valve returns to the closed position while surface tension holds the liquid within the decomposition chamber until complete decomposition and exhaust of the warm gaseous products through a converging and diverging nozzle occurs. The increasing and decreasing transient pressure in the decomposition chamber changes each cycle in response to flow control valve actuation as the decomposition process is repeated.

Abstract: Control and/or drive device for a flying body for ejecting hot gas streams of a combusted fuel combination of at least a first and second component. Device includes a first hollow chamber body structured and arranged to contain first component, a second hollow chamber body structured and arranged to contain second component, a controllable fuel valve arranged between first hollow chamber body and second hollow chamber body to control feed of first component to second hollow chamber body, and a plurality of outlets structured and arranged to eject respective hot gas streams for influencing a flight path of flying body. Second hollow chamber body is formed as a combustion chamber for combusting the at least first and second components within second hollow chamber body to generate respective hot gas streams, and plurality of outlets are connected to the second hollow chamber body.

Abstract: An engine assembly includes a gas-turbine engine having a tailcone portion and a bypass duct, a rocket engine combustion assembly located at the tailcone portion of the gas-turbine engine, and a movable nozzle segment subassembly that is selectively engageable with the gas-turbine engine bypass duct in an open position and with the rocket engine combustion assembly in a closed position.

Abstract: The electrolytic ignitor comprises an injector (2) constituting a first electrode, a second electrode (5) that is electrically insulated from the injector (2) by an insulator (4) and that extends downstream beyond the injector (2), a tank (12) for the first mono-propellant, a solenoid valve (11) interposed between the tank (12) and a distribution channel (3) serving to deliver the first mono-propellant to the first mono-propellant injector device constituted by at least one injection hole (1) opening out in the vicinity of the second electrode (5), and an electrical power supply circuit adapted to raise the second electrode (5) to a potential lying in the range 50 V to 1000 V relative to the potential of the first electrode (2).

Abstract: A rocket engine includes a combustor assembly for carrying out a combustion process of fuel and oxidizer rocket propellants to produce thrust. A swirl generator is positioned within the combustor assembly to produce a turbulent flowfield of the fuel and oxidizer rocket propellants within the combustor assembly.

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 sliding connection between components of dissimilar materials includes a first mating surface formed on a first component and a second mating surface formed on a second component and configured to mate with the first mating surface. The first component is connected to the second component at the first and the second mating surfaces by a fastener including a shank, a sleeve received by the shank, and a resilient member received by the shank and contained by the sleeve. The resilient member is configured to press the first mating surface and the second mating surface against one another and to permit the first mating surface to slide against the second mating surface.

Abstract: A heat exchange injector assembly includes a heat exchange element comprising a fuel sleeve, a liquid oxidizer post disposed in the fuel sleeve, and a multi-passage swirl member such as a double helix member, disposed in the liquid oxidizer post.

Abstract: A combustion chamber for expelling a hot gas stream from a rocket engine includes a combustion chamber wall and a plurality of cooling channels arranged inside the combustion chamber wall and configured to receive a flow of a cooling medium. Each cooling channel defines a longitudinal axis, a depth direction, and a substantially rectangular cross section. At least one cooling channel includes a plurality of depressions configured to prevent a stratification of the cooling medium in the at least one cooling channel. The plurality of depressions are arrange transverse to a direction of the flow direction and transverse to the longitudinal axis in the depth direction.

Abstract: A method of fabrication of a rocket engine nozzle assembly using pressure brazing generally includes initially assembling a rocket engine nozzle liner into a rocket engine nozzle jacket for a rocket engine nozzle assembly. The rocket engine nozzle assembly may then be sealed. Prior to pressure brazing the rocket engine nozzle assembly, pressure brazing parameters may be determined. The pressure brazing may be performed with the determined pressure brazing parameters to complete the fabrication of the rocket engine nozzle. The rocket engine nozzle assembly may include a rocket engine nozzle jacket and a rocket engine nozzle liner having a plurality of channels, with the space between each channel defining a land, and the rocket engine nozzle liner having at least a pair of endlands disposed at each end thereof. The rocket engine nozzle liner is bonded to the rocket engine nozzle jacket by the endlands being bonded to the nozzle jacket and the lands being pressure brazed to the nozzle jacket.

Abstract: A catalytically activated transient decomposition propulsion system provides thrust by decomposing flow controlled propellant in contact with a catalyzing agent using a fixed volume of liquid propellant that is placed in contact with the catalyst within the decomposition chamber by a calibrated flow control valve. After injecting the liquid propellant into the decomposition chamber, the valve returns to the closed position while surface tension holds the liquid within the decomposition chamber until complete decomposition and exhaust of the warm gaseous products through a converging and diverging nozzle occurs. The increasing and decreasing transient pressure in the decomposition chamber changes each cycle in response to flow control valve actuation as the decomposition process is repeated.

Abstract: Propellant management systems and methods are provided for controlling the delivery of liquid propellants in a space launch vehicle utilizing multiple rockets. The propellant management systems and methods may be configured to enable substantial simultaneous depletion of liquid propellants in each of a plurality of active rockets during operation of various booster stages of the launch vehicle.

Abstract: A pump for pressurizing a fluid includes an engine portion including a first pressure vessel, a first piston movable inside the first pressure vessel, at least two pressurant entrance valves connected to the first pressure vessel, and at least two pressurant exit valves connected to the first pressure vessel. The valves are configured to be opened and closed automatically and directly as a function of a position of the first piston inside the first pressure vessel. The valves are also configured to be automatically opened and closed out of phase with each other. The pump also includes a pump portion including a second pressure vessel, a second piston connected to the first piston and movable inside the second pressure vessel, at least two fluid entrance valves connected to the second pressure vessel, and at least two fluid exit valves connected to the second pressure vessel.

Abstract: Methods and apparatus for providing a Two-Stage Ignition System are disclosed. In one embodiment of the invention, a pilot stage (16) is employed to ignite a plurality of propellants (12, 14) and to create a pilot flame (22). The plurality of propellants (12, 14) are ignited in the main combustion stage (24) using the pilot flame (22), and a flow of an elevated temperature combustion product (30) is produced.

Abstract: A combustion chamber/supersonic nozzle assembly is cooled by an array of coolant channels in the wall of the assembly with coolant being fed to the assembly at the throat plane between the subsonic (convergent) and supersonic (divergent) sections. A minor portion of the coolant entering at the throat plane is directed to coolant channels in the supersonic section wall, while the remainder is directed to a complex arrangement of channels in the subsonic section wall. The latter arrangement includes an outer layer of wide axially oriented channels for delivery of coolant to, and removal from, an inner layer of smaller, circumferentially oriented channels that are adjacent to the hot gas and carry the bulk of the coolant load. The path of coolant travel through each inner channel is relatively short, lessening the pressure drop through those channels relative to coolant channel arrangements of the prior art.

Abstract: Method and arrangement for providing a component (1) for being subjected to high thermal load during operation. The component includes a wall structure, which defines an inner space for gas flow. The component is formed by at least a first part (5) that includes an inner wall (8), an outer wall (9) and at least one cooling channel (11) between the walls. An end portion of said inner wall of the first part of the component is joined to a second part (6). The joint (18) is located at a distance from the interior of the component.

Abstract: A method of pumping a fluid includes: providing a pressurizer for pressurizing the fluid, the pressurizer including at least two storage tanks, where, for each storage tank, the pressurizer includes a propellant entrance valve, a propellant exit valve, a pressurant entrance valve, and a pressurant exit valve, where each of the storage tanks is configured to be filled with the fluid under a low pressure and drained of the fluid under a high pressure by the force of a pressurant; and for each storage tank, opening and closing its associated valves in cycles to sequentially fill and drain the storage tank of the fluid, the cycles each having a cycle time between approximately 1 and 500 milliseconds.

Abstract: A propulsion system for creating a propulsive force has a combustion chamber, a pair of electrodes within the combustion chamber and a power supply attached to the electrodes to create a high voltage field within an initiation zone of the combustion chamber, and an injector for introducing a propellant, preferably in atomized form, into the high voltage field for creating the propulsive force. In a preferred embodiment of the present invention, the propellant is hydrogen peroxide. In another embodiment of the present invention, a second propellant is introduced into the combustion chamber for increasing the propulsive force.

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 system to provide a two piece robust fluid injector. According to various embodiments, the fluid injector is a fuel injector for a combustion engine. The injector includes two coaxially formed annuluses. One annulus is formed in a face plate and the second annulus or hole is defined by a tube extending through the face plate. The tube extends through the face plate in a portion of a through bore which also is used to define the second annulus. The second annulus is formed using a throughbore through which the tube extends. This allows the second annulus to always be formed inherently and precisely substantially coaxial with the first annulus. Moreover, the second annulus can be formed with a much greater tolerance than if other independent components needed to be added.

Abstract: A pressurizer for pressurizing a fluid includes: at least two storage tanks, where, for each storage tank, the pressurizer further includes: a propellant entrance valve, a propellant exit valve, a pressurant entrance valve, and a pressurant exit valve, where each of the storage tanks is configured to be filled with the fluid under a low pressure when its propellant entrance and pressurant exit valves are open and its propellant exit and pressurant entrance valves are closed, and to be drained of the fluid under a high pressure by the force of a pressurant when its valves are reversed, where its valves are configured to be opened and closed in a cycle to sequentially fill and drain the storage tank of the fluid, the cycle having a cycle time of between 1 and 500 milliseconds, and where the cycles of the valves of the storage tanks are out of phase with each other such that at some time in which one storage tank is being filled with the fluid, at least one other storage tank is being drained of the fluid.

Abstract: Disclosed is a propulsion system for a spacecraft. The propulsion system includes a supply of oxidizer and at least one nozzle. A conduit fluidly couples the supply of oxidizer and the nozzle. The conduit provides a pathway for oxidizer to flow in a downstream direction from the supply of oxidizer toward and into the nozzle. A pressure regulator is coupled to the conduit and is interposed between the supply of oxidizer and the nozzle, wherein the pressure regulator regulates the pressure of oxidizer flowing through the conduit and downstream of the pressure regulator to a pressure at or below the pressure required to maintain the oxidizer in a gas state to ensure that the any oxidizer flowing through the conduit is in a gas state prior to entering the nozzle. The conduit supplies oxidizer from the supply of oxidizer to a hybrid rocket motor.

Abstract: Multilayer composite for combustion chambers or nozzles of missiles, comprising an interior layer in contact with the combustion gases and an outer layer, wherein the interior layer is a fiber-reinforced ceramic whose matrix comprises phases of carbon and/or phases of silicon carbide and the outer layer is a polymer reinforced with carbon fibers, process for producing the same and combustion chambers and nozzles for missiles made of this composite.

Abstract: An ablative baffle for a liquid rocket engine thrust chamber that includes a hub member having a hollow structure, of which both top and bottom parts are open; a plurality of blade rib members, each of which is connected removably at one end to the outer surface of the hub member; and a blade-connecting member having a hollow structure, of which both top and bottom parts are open, to the inner part of which each of blade rib members is connected at the other end. An ablation cooling method is utilized which includes the composite heat resistant material structure with metal core, so that a liquid rocket engine system may be simplified, and the reliability of liquid rocket engine increases, and the manufacturing cost is cut down.

Abstract: A method and apparatus for augmenting thrust in a rocket traveling through atmospheric gas. Rocket motor designs are provided where a throat(s) from one or more rocket motors eject high-speed primary exhaust gas in a configuration which peripherally surrounds an outlet for induced, secondary gas. The secondary gas is mixed with the jet of primary exhaust gas to add momentum, and therefore thrust. Either expansion deflection or plug type rocket discharge nozzles can be utilized. In one embodiment, a thrust augmentation of over one hundred percent is achieved. In another embodiment, a plurality of rocket motor assemblies each containing a thrust augmenting rocket motor design is affixed to a rocket body. Such rocket motors enhance rocket thrust performance, and enables more efficient payload to rocket motor selection, or, alternatively, allows higher loads to be carried with the same amount of thrust.

Abstract: While commercial jet aircraft are capable of transporting large pay load economically, the same is not true for rockets. To solve this problem, a hybrid propulsion system is proposed for use in Horizontal Take-Off and Land (HOTOL) craft. The purpose is to maximized the use of ambient oxygen and minimize the use of cryogenic oxygen. At the core of the system is a jet engine whose after burner is modified to act as an evaporator. Liquid oxygen is pumped in, vaporized and ducted forward to the air intake through a gate. The gate is designed to control the mix of ambient oxygen and cryogenic oxygen going into the engine. It is also used as a hear shield to protect the engine from reentry temperatures.

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.

Abstract: A rocket engine assembly is provided for a vertically launched rocket vehicle. A rocket engine housing of the assembly includes two or more combustion chambers each including an outlet end defining a sonic throat area. A propellant supply for the combustion chambers includes a throttling injector, associated with each of the combustion chambers and located opposite to sonic throat area, which injects the propellant into the associated combustion chamber. A modulator, which may form part of the injector, and which is controlled by a controller, modulates the flow rate of the propellant to the combustion chambers so that the chambers provide a vectorable net thrust. An expansion nozzle or body located downstream of the throat area provides expansion of the combustion gases produced by the combustion chambers so as to increase the net thrust.

Abstract: The pointed end of a central spike, used in an engine of a space launcher, may have a deformable structure that can adopt either a folded position or a deployed position.

Abstract: An injector for use with the rocket thruster has a plurality of fuel ports separated from a plurality of oxidizer ports. The oxidizer and fuel ports are paired together directing their respective fluids along a path with radial and tangential components so that the two fluids impinge at a predetermined spaced apart distance from the chamber wall of the combustion chamber at an impingement track. By providing the fuel at a steeper angle relative to the chamber walls than the oxidizer, the fuel can be utilized to provide a fuel rich zone near the chamber walls to assist in cooling the chamber walls during operation.

Abstract: In a supersonic nozzle incorporating injectors and a combustion chamber as part of an expander cycle rocket engine, the oxidizer is injected in two streams. One of the streams, preferably a small fraction of the total, is injected into an upstream or preburner section of the combustion chamber and the other to a downstream or main section of the chamber. The preburner combustion gas is cooled in a substantially uniform manner to a moderate temperature by cooling the bulk of the gas rather than cooling only the gas in a boundary layer adjacent to the chamber wall. The combustion gas produced in the downstream section is hotter, and heat from that gas is drawn through the chamber wall into a jacket. The limited combustion in the preburner permits the use of a cooling element with highly intimate heat exchange construction, extracting a high level of energy from the preburner gas without damage to the cooling element and an overall improvement in the regenerative cooling.

Abstract: An improved engine construction, such as a rocket engine construction is provided. The engine construction comprises a combustion chamber, a smooth wall nozzle, and a transition zone between the chamber and the smooth wall nozzle. The transition zone has a coolant system which includes a manifold formed from a non-copper material through which a coolant flows. In a high heat transfer embodiment, the transition zone includes an additional manifold formed from a copper based material.

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 hypergolic fuel system comprises hydrogen peroxide, silane and a liquid fuel. The hypergolic fuel system is employed in a method for producing thrust, for example in a rocket, by contacting hydrogen peroxide with silane to obtain initial ignition and thereafter feeding a liquid fuel for combustion.

Abstract: In a supersonic nozzle incorporating injectors and a combustion chamber as part of an expander cycle rocket engine, the oxidizer is injected in two streams. One of the streams, preferably a small fraction of the total, is injected into an upstream or preburner section of the combustion chamber and the other to a downstream or main section of the chamber. The preburner combustion gas is cooled in a substantially uniform manner to a moderate temperature by cooling the bulk of the gas rather than cooling only the gas in a boundary layer adjacent to the chamber wall. The combustion gas produced in the downstream section is hotter, and heat from that gas is drawn through the chamber wall into a jacket. The limited combustion in the preburner permits the use of a cooling element with highly intimate heat exchange construction, extracting a high level of energy from the preburner gas without damage to the cooling element and an overall improvement in the regenerative cooling.

Abstract: A hypergolic fuel system comprises hydrogen peroxide, silane and a liquid fuel. The hypergolic fuel system is employed in a method for producing thrust, for example in a rocket, by contacting hydrogen peroxide with silane to obtain initial ignition and thereafter feeding a liquid fuel for combustion.

Abstract: An actively-cooled, fiber-reinforced ceramic matrix composite thrust chamber for liquid rocket propulsion systems is designed and produced with internal cooling channels. The monocoque tubular structure consists of an inner wall, which is fully integrated to an outer wall via radial coupling webs. Segmented annular void spaces between the inner wall, outer wall and adjoining radial webs form the internal trapezoidal-shaped cooling channel passages of the tubular heat exchanger. The manufacturing method enables producing any general tubular shell geometry ranging from simple cylindrical heat exchanger tubes to complex converging-diverging, Delaval-type nozzle structures with an annular array of internal cooling channels. The manufacturing method allows for transitioning the tubular shell structure from a two-dimensional circular geometry to a three-dimensional rectangular geometry.

Abstract: An apparatus is provided for burning a fuel and nitrous oxide. The apparatus has a combustor, a catalyst, a nitrous oxide supply passage for directing the nitrous oxide to a contact position with the catalyst, and a fuel supply passage for supplying the fuel to the combustor. The catalyst is for facilitating decomposition of the nitrous oxide, and the combustor is for burning the fuel, the decomposed nitrous oxide and/or further nitrous oxide decomposed in the reaction.

Abstract: A combustion chamber structure for a rocket engine or the like, has a combustion chamber liner with cooling channels and at least one manifold for feeding and removing a coolant, particularly a cryogenic fuel. The at least one manifold is brazed together with the combustion chamber liner and the area of the combustion chamber liner that is not covered by the at least one manifold is coated with an electroplated structural jacket. The invention also includes a method for manufacturing the combustion chamber structure.

Abstract: Heat is extracted from the combustion gas in a rocket engine combustion chamber by diverting portions of the gas through channels in the nozzle wall. The channels are layered between channels of coolant, which in expander cycle rocket engines is uncombusted fuel, to achieve intimate heat exchange between the combustion gas and the fuel. The combustion gas channels are relatively short, returning combustion gas thus cooled to the chamber interior. By drawing combustion gas from the chamber interior into the chamber wall, the cooling process no longer relies on the combustion gas boundary layer for heat transfer as in the prior art.

Abstract: The invention is a coolant system for a rocket engine. The rocket engine includes an injector, a fuel supply, an oxidizer supply, a pump for feeding fuel from the fuel supply to the injector, a pump for feeding oxidizer from the oxidizer supply to the injector, a combustor, and a nozzle, the combustor and nozzle forming a combustor and nozzle assembly. The coolant system includes a cooling jacket surrounding the combustor and nozzle assembly. A coolant fluid is supplied to the combustor jacket by a coolant pump that circulates the coolant fluid through the jacket. The coolant fluid vaporizes to cool the combustor and nozzle assembly. A turbine is operatively connected to the coolant pump and is driven by the vaporized coolant from the jacket. A heat exchanger transfers heat from the vaporized fluid to fuel or oxidizer from the supply thereof prior to the feeding of the fuel or oxidizer into the injector.

Abstract: The invention concerns a rocket engine wherein the combustion chamber consists of at least one first monolithic component made of a thermostructural composite material comprising a porous wall (5) through which the fuel is introduced in the core (C) of said combustion chamber. A small part of said fuel is directed towards the neck (3) for it to be cooled.

Abstract: A pressurizer for pressurizing a fluid includes: a pressurant entrance configured for the introduction of a pressurant; a fluid entrance configured for the introduction of the fluid; a fluid exit configured for the expulsion of the fluid; and at least one transfer chamber movable in a cycle with respect to at least one of the pressurant entrance, the fluid entrance, and the fluid exit, where the pressurizer is configured so that for a portion of a cycle the pressurant exerts a force on the fluid inside the transfer chamber, and where the transfer chamber is configured to receive the pressurant via the pressurant entrance, receive the fluid via the fluid entrance, and expel the fluid via the fluid exit by the force exerted by the pressurant upon the fluid inside the transfer chamber.