Abstract: A premixer for a combustor includes: a centerbody having a hollow interior cavity; a swirler assembly radially outward of the centerbody; a peripheral wall disposed radially outward of the centerbody and the swirler assembly such that a mixing duct is defined between the peripheral wall and the centerbody, downstream from the swirler assembly; an annular splitter radially inward of the swirler assembly and radially outward of the centerbody such that a radial gap is defined between the splitter and an outer surface of the centerbody, wherein the splitter includes a trailing edge which extends axially aft of the swirler assembly; a fuel gallery disposed inside the interior cavity of the centerbody; and at least one fuel injector extending outward from the fuel gallery and passing through an injector port communicating with the outer surface of the splitter.

Abstract: Proposed is a hybrid rocket engine using an electric motor-driven oxidizer pump, the hybrid rocket engine including: an oxidizer tank configured to store the oxidizer; an oxidizer pump configured to pressurize the oxidizer by being connected to the oxidizer tank through a first oxidizer supply line; a drive unit including an electric motor configured to drive the oxidizer pump and a battery configured to supply power to the electric motor; an auxiliary oxidizer line configured to guide the oxidizer from the oxidizer tank to the electric motor to cool the electric motor; an oxidizer recirculation line configured to recharge oxidizer vapor, generated through heat exchange between the electric motor and the oxidizer, to the oxidizer tank, thereby pressurizing an inner side of the oxidizer tank; and a combustion chamber configured to combust the oxidizer and fuel by being connected to the oxidizer pump through a second oxidizer supply line.

Abstract: Injector assemblies (12) and ducting arrangement (20) including such injector assemblies are provided. The injector assembly may include a reactant-guiding structure (16) arranged to convey a flow of reactants (19) into the combustion stage and means for injecting (24, 25, 26) a flow of air (22) into the combustion stage. The flow of air injected into the combustion stage may be arranged to condition interaction of the flow of reactants injected into the combustion stage with a cross-flow of combustion products (21), as the flow of reactants is admitted into the combustion stage.

Abstract: A spacecraft includes a propulsion system that includes one or more pressurant tanks configured to store an inert gas at a high pressure, one or more propellant tanks configured to store liquid propellant at an intermediate pressure, electric thrusters operable with the inert gas at a low pressure and pneumatically coupled with the one or more pressurant tanks by way of a first pressure regulator, and chemical thrusters operable with the liquid propellant. The inert gas is one or a mixture of two or more of xenon, argon and krypton. At least a portion of the liquid propellant is stored in at least one of the propellant tanks, the propellant tank including an ullage volume pneumatically coupled with at least one of the pressurant tanks by way of a second pressure regulator having an output set to the intermediate pressure and the ullage volume is pressurized by the inert gas.

Abstract: A rocket stage for a spacecraft includes an engine, a tank for storing a fuel and an oxidizer for combustion in the engine, a rocket stage primary structure, and an engine thrust frame that connects and transmits forces between the engine and the primary structure. The engine thrust frame includes at least an internal part thereof arranged internally within the tank. This internal part of the engine thrust frame forms an imperforate partition that divides an interior space of the tank into at least two chambers for storing the fuel and the oxidizer separately from one another. Propellant management devices including liquid guide vanes and refillable liquid reservoirs may be provided in connection with the partition in the tank interior space.

Abstract: An airblast fuel injector is provided for a fuel spray nozzle of a gas turbine engine. The injector has an annular air passage for the passage of a swirling air flow therethrough. The swirling air flow is used by the injector to produce an atomised fuel spray. The air passage contains a swirler for producing the swirling air flow, the swirler comprising a circumferential row of vanes which span inner and outer side walls of the air passage. Viewed on a longitudinal section through the injector, the air passage has a bend downstream of the swirler, the bend changing the direction of the air passage. The vanes are configured to introduce a radial component to the air flow exiting the swirler, the radial component guiding the air flow around the bend.

Abstract: The invention relates to a burner arrangement for using in a single combustion chamber or in a can-combustor comprising a center body burner located upstream of a combustion zone, an annular duct with a cross section area, intermediate lobes which are arranged in circumferential direction and in longitudinal direction of the center body. The lobes being actively connected to the cross section area of the annular duct, wherein a cooling air is guided through a number of pipes within the lobes to the center body and cools beforehand at least the front section of the center body based on impingement cooling. Subsequently, the impingement cooling air cools the middle and back face of the center body based on convective and/or effusion cooling. At least the back face of the center body includes on the inside at least one damper.

Abstract: A design for an effectively cooling a liner of a gas turbine combustor by means of convective cooling is disclosed.

Abstract: Variable speed drive controlled discharge pumps are used for pumping chilled water from a chilled water storage tank to coils operatively associated with the air inlet of a gas turbine, wherein the gas turbine is used to generate electricity. Use of the variable speed drive controlled discharge pumps aids is protecting the temperature distribution in the chilled water storage tank so that a thermocline rises with the addition of chilled water during charging of the tank with chilled water, and lowers during discharge of the tank when using the chilled water to cool inlet air at one or more gas turbines operated at a facility.

Abstract: The invention relates to a method for operating a gas turbine, in which an oxygen-reduced gas and fresh air are delivered to a compressor of the gas turbine in a radially staged manner, the fresh air being delivered via an outer sector of the inlet cross section in relation to the axis of rotation of the compressor, and the oxygen-reduced gas being delivered via an inner sector of the inlet cross section in relation to the axis of rotation of the compressor. The invention relates, further, to a gas turbine power plant with a gas turbine having a compressor inlet which is followed by the flow duct of the compressor and which is divided into an inner sector and an outer sector, a feed for an oxygen-reduced gas being connected to the inner sector of the compressor inlet, and a fresh air feed being connected to the outer sector of the compressor inlet.

Abstract: A gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a fan. A geared architecture is configured for driving the fan. A turbine section is configured for driving the geared architecture. A thermal management system that includes a first fluid circuit and a second fluid circuit that manage heat generated in at least a portion of the gas turbine engine. A first heat exchanger is incorporated into each of the first fluid circuit and the second fluid circuit. A second heat exchanger is incorporated into the first fluid circuit. A valve controls an amount of a first fluid that is communicated to the first heat exchanger and the second heat exchanger. A controller is configured to control a positioning of the valve.

Abstract: A steam reheat process is provided to enhance a thermal power cycle, and particularly a renewable steam thermal cycle. An oxyfuel combustion gas generator is provided which combusts a hydrogen and/or carbon containing fuel with an oxidizer of primarily oxygen to generate products of combustion including steam and/or carbon dioxide. Water from the thermal cycle is directed to the reheater for mixing with the products of combustion within the reheater to generate a working fluid containing steam. This steam is routed through a turbine or other expander and power is outputted from the system. The water is optionally thereafter condensed and at least partially routed back to the thermal cycle. Any carbon dioxide within the working fluid can be separated in a condenser downstream of the expander for capture of the carbon dioxide, such that increased power output for the thermal power cycle is achieved without atmospheric emissions.

Abstract: A fuel supply system for a gas turbine includes a combustion section, a transition duct downstream from the combustion section, a turbine section downstream from the transition duct, and a first stage of stationary vanes circumferentially arranged inside the turbine section. A hot gas path is between the transition duct and the stationary vanes, and a fuel injector provides fluid communication into the hot gas path. A method of supplying fuel to a gas turbine includes combusting a first fuel in a combustion chamber to produce combustion gases, flowing the combustion gases through a transition duct to a hot gas path, and flowing the combustion gases through the hot gas path to a first stage of stationary vanes in a turbine section. The method further includes injecting a second fuel into the hot gas path downstream from the transition duct and upstream from the first stage of stationary vanes.

Abstract: A fuel purging system for a turbine assembly includes a fuel delivery system. The fuel delivery system includes a fuel source for providing a fuel to the turbine assembly, a control valve for regulating a fuel flow of the fuel, a flow divider for selectively distributing the fuel to at least one combustor, and a combustor valve located upstream of the at least one combustor. The fuel purging system also includes a steam source for distributing a steam to the fuel delivery system at a location upstream of the combustor valve.

Abstract: A combustion apparatus includes a first device for mixing a fuel with an oxidant, a combustion chamber in which combustion of the fuel/oxidant mix takes place, a pre-chamber located between the first device and the combustion chamber and a second device for outputting a gas flow into the pre-chamber. The first device is adapted for outputting the oxidant or the fuel or the fuel/oxidant mix to the pre-chamber. The second device is located upstream of the first device and wherein the second device is arranged for mixing fuel with a gas and the gas flow is a fuel/gas mix.

Abstract: The invention proposes a system for supplying (10) fuel to an aircraft turbomachine, which includes a low-pressure circuit (14) in which a low-pressure pump (12) is installed, and a high-pressure circuit (44) supplying an injection device (20) of the turbomachine and actuators (22) of components of the turbomachine, which includes several high-pressure pumps (30, 32), characterised in that each of the said high-pressure pumps (30, 32) consists of a stage of a high-pressure pump with several stages, each stage of the said pump with several stages being associated with each of the said segments (24, 26).

Abstract: In some implementations, a system may include a compressor, a heat exchanger and an ITM. The compressor is configured to receive an air stream and compress the air stream to generate a pressurized stream. The heat exchanger is configured to receive the pressured stream and indirectly heat the pressurized stream using heat from an oxygen stream from an Ion Transport Membrane (ITM). The ITM is configured to receive the heated pressurized stream and generate an oxygen stream and the non-permeate stream, wherein the non-permeate stream is passed to a gas turbine burner and the oxygen stream is passed to the heat exchanger.

Abstract: A method of fabricating a rolling bladder propellant tank assembly includes providing a pair of half-domes comprising a nanophase metallic material and a bladder formed of a ductile metallic material. A bladder support ring extends from at least one of the pair of half-domes. The pair of half-domes are positioned together to form a cavity therein with a portion of the bladder trapped between the pair of half-domes adjacent to the bladder support ring, wherein the bladder is positioned within the cavity. A seal is formed between the pair of half-domes and the bladder by friction stir welding. A friction stir welding pin used to form the seal is aligned with the bladder support ring during the friction stir welding operation.

Abstract: An engine unit, particularly for urban transport, comprising an engine (3) supplied with a compressed gas and having a expansion chamber (9), a liquid gas tank (28) in communication with the engine (3), and means (M, M?) for gasifying the liquid gas, which are interposed between the liquid gas tank (28) and the engine (3) for obtaining compressed gas. The gasifying means (M) comprise a gasification chamber (22) in communication with the liquid gas tank (28) and a liquid fuel tank (39) which is connected to the gasification chamber (22). The gasification chamber (22) is in fluid communication with both the liquid fuel tank (39) for the combustion of the liquid fuel with the oxygen of the liquid gas (22), and the expansion chamber (9) so that the compressed liquid gas and gaseous products of combustion process are used to do useful work.

Abstract: A method of assembling a combustion system includes providing a gas turbine engine that includes a gas turbine section coupled downstream from a combustion chamber and coupling a source of oxygen to the gas turbine engine such that a stream of oxygen discharged from the source facilitates displacing nitrogen in the working fluid of the gas turbine engine and facilitates decreasing emissions generated within the gas turbine engine.

Abstract: A propellant supply device (10) is provided for a vehicle (11) having a main propulsion motor (12) and having an attitude control system (11) including a plurality of thrusters (14A, 14B, 14C, . . . , 14F). The improved device comprises: a pressure vessel (15); first and second movable walls (20, 21) operatively arranged within the pressure vessel and dividing the interior space therewithin into three separate sealed chambers (22, 23, 24) from each of which fluid may be supplied; a first fluid (e.g., a first bipropellant) in one of the chambers; a second fluid (e.g., a second bipropellant) in a second of the chambers; and a third fluid (e.g., ammonia) in a third of the chambers, the third fluid being a volatile liquid having a liquid phase and a gaseous phase, and wherein all three chambers are pressurized to the vapor pressure of the third fluid.

Abstract: The present invention provides a widely variable and yet precisely controlled system for metering the flow of fuel to an engine component. The present invention utilizes a plurality of solenoid activated injector valves to meter fuel into an output plenum where the fuel is accumulated to avoid pressure pulses and pressurized if necessary for introduction to the combustion chamber of a gas turbine engine. This system is particularly well suited to high pressure engine systems such as helicopter jet or turbo shaft engines.

Abstract: An air intake includes a spike and a cylindrical cowl. The spike is composed of a plurality of plates which are arranged axial symmetrically around a central axis thereof, and the cowl includes an inner wall parallel to the central axis and is formed so as to enclose a rear portion of the spike via a given space. The adjacent ones of the plates form respective aerodynamic compressive surfaces in spaces formed by the adjacent ones. The distances between the adjacent ones of the plates are variable, and the distance between the forefront of the spike along the central axis and the cowl is variable.

Abstract: A propellant supply device (10) is provided for a vehicle (11) having a main propulsion motor (12) and having an attitude control system (11) including a plurality of thrusters (14A, 14B, 14C, . . . , 14F). The improved device comprises: a pressure vessel (15); first and second movable walls (20, 21) operatively arranged within the pressure vessel and dividing the interior space therewithin into three separate sealed chambers (22, 23, 24) from each of which fluid may be supplied; a first fluid (e.g., a first bipropellant)in one of the chambers; a second fluid (e.g., a second bipropellant) in a second of the chambers; and a third fluid (e.g., ammonia) in a third of the chambers, the third fluid being a volatile liquid having a liquid phase and a gaseous phase, and wherein all three chambers are pressurized to the vapor pressure of the third fluid.

Abstract: The cryogenic propulsion module comprises a main cryogenic thruster (10), two attitude-controlling secondary thrusters (21, 22), tanks (31, 32, 33, 34) for feeding cryogenic propellants, a device for intermittently pressurizing the tanks (31, 32, 33, 34), and a device for initiating firing of the main cryogenic thruster (10) in intermittent manner while the tanks (31, 32, 33, 34) are intermittently pressurized. The device for intermittently pressurizing a tank (31, 32, 33, 34) comprises a heat exchange circuit associated with a heat accumulator (61, 62) and a device (71, 72) for putting a predetermined quantity of a propellant into circulation in the heat exchanger circuit. The module also comprises a device for heating the heat accumulator (61, 62) in the periods between two consecutive firings.