Abstract: Gas turbines with multiple gas flow paths are provided. In this regard, a representative gas turbine includes: a spool; a compressor; a turbine mechanically coupled to the spool; the compressor having a first set of blades and a second set of blades, the second set of blades being located downstream of the first set of blades, the first set of blades and the second set of blades being driven by the spool; and means for enabling the first set of blades to rotate at a lower rotational speed than the second set of blades.

Abstract: Described herein are embodiments of systems and methods for oxidizing gases. In some embodiments, a reaction chamber is configured to receive a fuel gas and maintain the gas at a temperature within the reaction chamber that is above an autoignition temperature of the gas. The reaction chamber may also be configured to maintain a reaction temperature within the reaction chamber below a flameout temperature. In some embodiments, heat and product gases from the oxidation process can be used, for example, to drive a turbine, reciprocating engine, and injected back into the reaction chamber.

Abstract: A flow conditioning device (30, 50, 70, 100, 150) for a can annular gas turbine engine, including a plurality of flow elements (32, 34, 52, 54, 72, 74, 102) disposed in a compressed air flow path (42, 60, 80, 114, 122) leading to a combustor (12), configured such that relative adjustment of at least one flow directing element (32, 52, 72, 110) with respect to an adjacent flow directing element (34, 54, 74, 112, 120) during operation of the gas turbine engine is effective to adjust a level of choking of the compressed air flow path (42, 60, 80, 114, 122).

Abstract: A compressed air supply system for routing compressed air from a compressor to at least one combustor of a gas turbine engine is disclosed. The compressed air supply system may be formed from one or more plenums having an upstream end in fluid communication with an inner chamber of the compressor in which air is compressed and having a downstream end in fluid communication with the at least one combustor. Channeling compressed air through the plenum reduces damage to other components by confining the compressed air within the plenum. An upstream end of the plenum may be sealed to at least a portion of the compressor, and a downstream end of the plenum may be sealed to at least a portion of one or more combustors.

Abstract: The present invention discloses a novel apparatus and method for operating a gas turbine combustor having a structural configuration proximate a pilot region of the combustor which seeks to minimize the onset of thermo acoustic dynamics. The pilot region of the combustor includes a generally cylindrical extension having an outlet end with an irregular profile which incorporates asymmetries into the system so as to destroy any coherent structures.

Abstract: A system includes a bleed system configured to direct a bleed flow from a high pressure region to a low pressure region. The bleed system includes a valve configured to control the bleed flow through the bleed system and a staged bleed conduit configured to incrementally depressurize the bleed flow. The staged bleed conduit includes an inlet coupled to the valve, a first stage configured to depressurize the bleed flow that is coupled to the inlet, a second stage configured to depressurize the bleed flow that is coupled to the first stage, and an outlet coupled to the second stage. The outlet is configured to direct the bleed flow to the low pressure region. The inlet, the first stage, the second stage, and the outlet are disposed along parallel axes.

Abstract: Described herein are embodiments of systems and methods for oxidizing gases. In some embodiments, a reaction chamber is configured to receive a fuel gas and maintain the gas at a temperature within the reaction chamber that is above an autoignition temperature of the gas. The reaction chamber may also be configured to maintain a reaction temperature within the reaction chamber below a flameout temperature. In some embodiments, heat and product gases from the oxidation process can be used, for example, to drive a turbine, reciprocating engine, and injected back into the reaction chamber.

Abstract: Described herein are embodiments of systems and methods for oxidizing gases. In some embodiments, a reaction chamber is configured to receive a fuel gas and maintain the gas at a temperature within the reaction chamber that is above an autoignition temperature of the gas. The reaction chamber may also be configured to maintain a reaction temperature within the reaction chamber below a flameout temperature. In some embodiments, heat and product gases from the oxidation process can be used, for example, to drive a turbine, reciprocating engine, and injected back into the reaction chamber.

Abstract: A valve includes a body with bleed ports and a ring. The ring surrounds the body and includes two adjacent ring segments. The ring is movable between an open position and a closed position, the latter of which prevents flow through the ports.

Abstract: A first orifice on a first inner surface of a first annular zone of a forward-bounded annular combustor provides for injecting air forward and radially outwards thereinto, from a location aft of a fuel slinger/injector. A radial dimension of a first outer surface of the first annular zone exceeds that of a corresponding first inner surface. A radial dimension of a second outer surface of a second annular zone of the annular combustor—aft of the first annular zone—exceeds that of a corresponding second inner surface. The first and second outer surfaces, and the first and second inner surfaces, are respectively coupled by transitional outer and inner surfaces of an annular transition zone located therebetween. The radial dimensions of the transitional outer and inner surfaces at the junctions with the corresponding second surfaces exceed the corresponding radial dimensions at the junctions with the corresponding first surfaces.

Abstract: The present invention relates to a gas-turbine combustion chamber with mixing air orifices, with a combustion chamber wall, with tiles arranged on the inside of the combustion chamber wall at a certain distance from said combustion chamber wall as well as with mixing air orifices passing through the combustion chamber wall and the tiles, characterized in that the mixing air orifices are formed by chutes which are designed tube-like and pass through the tiles, and that several chutes are formed on a mixing air wall element extending at least around part of the circumference of the combustion chamber.

Abstract: Described herein are embodiments of systems and methods for oxidizing gases. In some embodiments, a reaction chamber is configured to receive a fuel gas and maintain the gas at a temperature within the reaction chamber that is above an autoignition temperature of the gas. The reaction chamber may also be configured to maintain a reaction temperature within the reaction chamber below a flameout temperature. In some embodiments, heat and product gases from the oxidation process can be used, for example, to drive a turbine, reciprocating engine, and injected back into the reaction chamber.

Abstract: The present invention relates to an aircraft gas turbine thrust-reversing device with an engine having an engine cowling, the rear area of which can be displaced in the axial direction of the engine from a closed forward thrust position into a rearwardly displaced thrust reversal position, resulting in an essentially annular free space towards a forward stationary area of the engine cowling, with the rear area of the engine cowling being functionally coupled to deflecting elements arranged in the forward thrust position within the front area of the engine cowling, with the deflecting elements during displacement of the rear area of the engine cowling being moveable on a partial-circular path facing the central axis of the engine and contactable by their rear end areas with a cowling of the core engine.

Abstract: A hydraulic circuit for a jet engine thrust reverser is disclosed. The hydraulic circuit may comprise a primary hydraulic actuator cylinder assembly and a secondary hydraulic actuator cylinder assembly. The primary hydraulic actuator cylinder assembly may be hydraulically coupled to the secondary hydraulic actuator cylinder assembly, in a master-slave relationship, and the secondary hydraulic actuator cylinder assembly may drive a thrust reverser blocker door. The primary hydraulic actuator cylinder assembly may comprise a rod, and the rod may comprise a channel through which hydraulic fluid is capable of flowing. The hydraulic circuit may further comprise a group of secondary hydraulic actuator cylinder assemblies, wherein the group of secondary hydraulic actuator cylinder assemblies may be coupled to the primary hydraulic actuator cylinder assembly.

Abstract: A combustion chamber for a gas turbine is proposed. The combustion chamber has a wall section and a brim element. The wall section has an inlet aperture for injecting a cooling medium into the combustion chamber. The brim element is mounted to an inner face of the wall section. The brim element is formed in such a way that a projected area of the brim element onto the inner face along a direction of a normal of the inner face at least partially covers the inlet channel.

Abstract: A system including a multi-tube fuel nozzle, including a first plate having a first plurality of openings, a plurality of tubes extending through the first plurality of openings in the first plate, wherein each tube of the plurality of tubes includes an air inlet, a fuel inlet, and a fuel-air mixture outlet, and a resilient metallic seal disposed along the first plate about the plurality of tubes.

Abstract: A liner for a gas turbine engine includes a liner defining an inner surface exposed to exhaust gases and a duct spaced radially outward of the liner. A plurality of hanger assemblies is disposed within the radial space between the liner and the duct for supporting the liner relative to the duct. Each of the hanger assemblies includes a cable having a first end attached to the duct and a second end attached to the liner.

Abstract: A fuel nozzle arrangement includes an annular combustor having four quadrants. Fuel injectors are located in the four quadrants and each fuel injector has either a first or second type nozzle attached to it. The first and second type fuel nozzles are located in an array so that at least two of one nozzle type are located together to allow for non-uniform injector spacing with uniform downstream temperatures.

Abstract: A rocket having a motor with a user adjustable thrust is provided. The rocket includes a main cylinder containing a rocket propellant, the propellant configured to generate gas during operation and one or more nozzles arranged to direct the gas in a first direction. A thrust adjustment device is arranged to receive a portion of the gas, the thrust adjustment device configured to change the direction of flow of at least a portion of the gas.

Abstract: A gas turbine is configured to operate with a high temperature combustion gas stream. The gas turbine may include a combustor that provides a combustion gas stream including charged particles and at least one turbine stage including at least one high temperature surface that may be driven with a voltage selected to repel the charged particles. The at least one high temperature surface may output a film-cooling layer including cool air, the film-cooling layer being stabilized by Coulombic forces between the voltage and the charged particles.