Abstract: An injector of a propellant combustion member, the injector including a first feed and a second feed; the first feed being connected to a plurality of feed chimneys arranged around a longitudinal axis; the second feed being configured so as to feed an injection chamber; and the injection chamber being connected to a plurality of feed sheaths with injection orifices, each feed chimney being surrounded by a coaxial feed sheath; the injector including a takeoff chamber connected to a takeoff duct for taking off the second propellant, the takeoff chamber being arranged in such a manner as to take off the second liquid propellant from the injection chamber; and it is configured to achieve uniform injection of the second propellant via the injection orifices.

Abstract: A gas turbine engine having a core engine. The core engine includes an inlet, a compressor section, a combustion section, a turbine section, and an exhaust. The gas turbine engine also includes a bearing sump arranged in the core engine for containing lubrication, the bearing sump and lubrication having an operational range between at least about 0° F. and about 550° F.

Abstract: A thruster nozzle (24) includes a throat (28) and a nozzle end portion (30) that is integral with the throat. The nozzle end portion includes an outer wall (32) that has a plurality of circumferentially-disposed nozzle features or connectors (34). A brace (26) may be disposed at least partially around the thruster nozzle and mated with the nozzle features to restrict relative movement between the thruster nozzle and the brace. The brace may be attached with a vehicle body (22) of a vehicle.

Abstract: One of the objects of the invention is to provide a water-saving type advanced humid air gas turbine system (AHAT) that can decrease the amount of makeup water to be supplied from the outside, by reducing the amount of water consumed when the gas turbine system is starting up, shut down, or subjected to load rejection. The gas turbine system includes a compressor, the compressed air header for generating humidified combustion air, a combustor for generating combustion gas, and the turbine. When the gas turbine system is starting up, shut down or subjected to load rejection, steam coming from the heat recovery steam generator is recovered by blocking the first steam system and making the second steam system communicate with the heat recovery steam generator.

Abstract: Methods, systems and apparatuses are disclosed for delivering high triple point propellant to a rocket engine and maintaining the desired phase of the propellant during engine ignition.

Abstract: An apparatus and method to enhance the performance of rockets engines which utilize liquid methane/oxygen propellants by injecting optimized amounts of pressurized hot helium gas into the combustion chamber with the propellants. In one embodiment, the pressurized helium gas is stored at low temperatures near those of the cryogenic propellants and is used for regenerative cooling of the combustion chamber and nozzle during rocket operation in order to raise the temperature of the helium gas before being injected into the combustion chamber.

Abstract: A fragmenting nozzle system includes a first nozzle at least partially disposed within a second nozzle. The first nozzle includes an ablative shell, a syntactic foam support disposed between the ablative shell and the second nozzle, and an ignition system disposed at least partially within the syntactic foam support. For example, the ignition system is operable to generate a controlled-energy deflagration pressure wave that fragments the first nozzle but not the second nozzle.

Abstract: A thrust nozzle, in particular thrust nozzle for a rocket engine, with a convergent wall section, a throat section and a divergent wall section. The divergent wall section has a first region adjacent to the throat section, the wall contour of which region corresponds to a truncated ideal nozzle, and the divergent wall section has a second region facing away from the throat section, which region has a wall contour deviating from the first region.

Abstract: A nozzle (1) for a space vehicle engine (M), the nozzle comprising a stationary portion (2) and a movable portion (3), the nozzle (1) including a pneumatic deployment system (4) comprising: a deployment actuator (5) for deploying the movable portion (3) of the nozzle (1); a high unlocking actuator (6); a low unlocking actuator (7); and an ejector (41); the deployment system (4) including a feed system (8) configured so as to, sequentially: move the deployment actuator (5) from its support position towards its deployment position; move the high and low unlocking actuators (6, 7) into their high and low unlocking positions; and actuate the ejector so as to eject the deployment system (4) from the nozzle (1).

Abstract: Electrically operated propellant is used to supplement the thrust provided by solid rocket motor (SRM) propellant to manage thrust produced by a SRM. The gas produced by burning the electrically operated propellant may be injected upstream of the nozzle to add mass and increase chamber pressure Pc, injected at the throat of the nozzle to reduce the effect throat area At to increase chamber pressure Pc or injected downstream of the throat to provide thrust vector control or a combination thereof. Certain types of electrically operated propellants can be turned on and off provided the chamber pressure Pc does not exceed a self-sustaining threshold pressure eliminating the requirement for physical control valves.

Abstract: An injection device, a combustor, and a rocket engine include a device main body partitioned into a fuel manifold and an oxidant manifold, and a plurality of injectors arranged at predetermined intervals in the device main body to inject fuel and oxidant into a combustion chamber, each of the injectors includes a LOx channel including a proximal end portion communicating with the oxidant manifold and a distal end portion communicating with the combustion chamber, a restrictor provided on the proximal end portion of the LOx channel and a GH2 channel including a proximal end portion communicating with the fuel manifold and a distal end portion communicating with the combustion chamber, and the restrictors have different shapes.

Abstract: A heat storage mechanism, of a heat exchanger, which does not shorten the service life of the heat exchanger even when a facility using the heat exchanger is intermittently operated and which suppresses a decrease in efficiency at the time of restart of the facility, is provided. The heat storage mechanism for storing heat of a heat exchanger during stop of operation of a facility provided with the heat exchanger includes an outflow prevention unit configured to prevent outflow of an exhaust gas to the outside, which is a heating medium of the heat exchanger, during stop of operation of the facility is provided in an exhaust passage through which the exhaust gas is discharged to the outside.

Abstract: A gas turbine engine includes an outer case structure around a central longitudinal engine axis. An intermediate case structure is included inboard of the outer case structure. An inner case structure is included inboard of the intermediate case structure. A variable area exhaust mixer is included, which is movable between a closed position adjacent to the intermediate case structure and an open position adjacent to the inner case structure.

Abstract: A system for providing ignition and pressurization of hypersonic vehicles is disclosed. The system combines the pressurization, barbotage and ignition functions into a single system saving mass and volume and simplifying the hypersonic vehicle plumbing. A monopropellant fuel such as Tridyne is used to pressurize a fuel tank, warm the fuel as it enters fuel injectors, and provide barbotage of the fuel just prior to its injection into a combustion chamber.

Abstract: A system for increasing the efficiency of heating cryogenic fluid flowing in a downstream direction through a fluid conduit includes a heating mechanism, an upstream valve, a downstream valve, and a controller. The heating mechanism heats the cryogenic fluid, resulting in conversion of a portion of the cryogenic fluid into a buoyant flow moving in an upstream direction. The upstream valve is located upstream of the heating mechanism and controls an upstream-valve mass flow rate of the cryogenic fluid. The downstream valve is located downstream of the heating mechanism and controls a downstream-valve mass flow rate of the cryogenic fluid. The controller adjusts the upstream valve to a choked position at which: an upstream-valve non-buoyant mass flow rate substantially matches the downstream-valve mass flow rate, and the upstream valve at least partially blocks the buoyant flow from flowing in the upstream direction past the upstream valve.

Abstract: A method and fuel supply system for supply of a combustion chamber with at least one combustible fluid are provided. The fuel supply system includes a combustion chamber, at least one supply circuit, and at least one purge circuit, the purge circuit coupled to the at least one supply circuit, the purge circuit including at least two isolation valves defining a cavity between, a source of relatively high temperature purge air coupled in flow communication to the cavity through one of the at least two isolation valves, a source of relatively low temperature ventilation air coupled in flow communication to the cavity, and a vent coupled in flow communication to the cavity, the at least one purge circuit configured to channel a flow of relatively low temperature ventilation air from the source relatively low temperature ventilation air through the cavity to the vent during operation of the combustion chamber.

Abstract: A lobe lance is disclosed for a gas turbine combustor which includes a plurality of N (N?4) lobe fingers, each configured as a streamlined body with two lateral surfaces. A plurality of nozzles for injecting a gaseous and/or liquid fuel mixed with air are provided whereby lobes running between the nozzles are provided for improving the mixing quality and reducing pressure loss in said combustor. The lobes of each lobe finger have one of two opposite orientations with respect to said flow direction, and the lobes of all lobe fingers follow a predetermined pattern of orientation across the lobe fingers at least one pair of neighboring lobe fingers has the same lobe orientation resulting in a grouped lobe arrangement ( . . . LL . . . or . . . RR . . . ) such that at least two of the vortices generated by the lobe shape downstream of the lobe fingers combine.

Abstract: A method of regulating pressure within a first propellant tank of a rocket engine having a first propellant tank containing a first propellant and a second propellant tank containing a second propellant, and a regulator device for regulating pressure within the first tank, the regulator device comprising a gas generator and a heat exchanger co-operating with the gas generator so as to vaporize at least part of the first propellant prior to reintroducing it into the first tank (16), the gas generator and the heat exchanger both being fed with the first propellant by a single first motor-driven pump, while the gas generator is fed with the second propellant by a single second motor-driven pump, wherein the flow rate of the first motor-driven pump is controlled as a function of a first parameter, while the flow rate of the second motor-driven pump is controlled as a function of a second parameter.

Abstract: The present invention relates to an igniter assembly, including a blast tube provided with a supporting part formed in a cylindrical shape, and an aft-closure part having front and rear sides through which igniter combustion gas is emitted, an igniter disposed adjacent to the rear side of the aft-closure part and storing ignition pellet therein, and a blast tube stopper formed in a shape surrounding the supporting part, accommodating the igniter, and coupled to the rear side of the aft-closure part, wherein the aft-closure part includes a plurality of holes through which igniter combustion gas is emitted.

Abstract: A solid rocket propellant grain having rocket propellant and a membrane in contact with the rocket propellant. The membrane includes a highly heat conductive pattern which affects the propellant burning rate through localized conductive heat transfer from the combustion zone and into the uncombusted propellant. Different geometries for the thermally conductive pattern produce different combustion results.