Abstract: A gas turbine combustor of the present invention comprises a fuel injector for injecting a fuel toward a combustion chamber; a swirler which takes-in compressed air generated in a compressor and swirls the compressed air, in the vicinity of the fuel injector; a tubular guide member for guiding the compressed air taken-in from the swirler, to the combustion chamber; and a heat shield having a cylindrical portion located outward relative to the guide member; wherein the cylindrical portion has a purge hole; and air is introduced through the purge hole and is supplied to a space formed between the guide member and the cylindrical portion.

Abstract: A fuel delivery system for a combustion turbine engine, comprising: a fuel line having a fuel compressor and parallel branches downstream of the fuel compressor: a cold branch that includes an after-cooler; and a hot branch that bypasses the after-cooler; a rapid heating value meter configured to measure the heating value of the fuel from the fuel source and transmit heating value data relating to the measurements; means for controlling the amount of fuel being directed through the cold branch and the amount of fuel being directed through the hot branch; and a fuel-mixing junction at which the cold branch and the hot branch converge; wherein the fuel-mixing junction resides in close proximity to a combustor gas control valve.

Abstract: A liquefied natural gas (LNG) transport vessel for transporting liquefied natural gas (LNG) is disclosed which is capable of storing excess boil off gas BOG until needed for combustion in one or more combustion apparatus on the vessel. A method for managing the delivery of the BOG to the combustion apparatus is also described. The LNG vessel includes at least one insulated LNG storage tank which stores LNG. A first stage LNG receiver receives and stores BOG from the at least one LNG storage tank. A second stage or high pressure BOG storage tank receives compressed BOG from the receiver and stores the BOG as needed for combustion by one or more combustion apparatus of the vessel. A pressure regulator allows BOG gas to be delivered to the combustion apparatus if there is sufficient pressure in the high pressure storage tank to passively deliver the BOG at a predetermined delivery pressure.

Abstract: A method is provided of reducing lockout time of a gas turbine engine which includes: an inlet, a compressor, a combustor, a turbine, and an exhaust duct, where the compressor and the turbine are carried on a turbomachinery rotor and each include an array of blades mounted for rotation inside a casing of the engine. The method includes: operating the engine at a first power output; shutting down operation of the engine without substantially reducing the power output beforehand, wherein thermomechanical changes occur in the engine subsequent to shutdown that tend to reduce a radial clearance between at least one of the blades and the casing; and subsequent to shutting down the engine, (1) heating the casing and/or (2) pumping an airflow of ambient air into the inlet and through the casing, past the rotor, and out the exhaust duct, so as to reverse at least partially the thermomechanical changes.

Abstract: The present invention's nozzle is especially suitable for use in a shipboard air-discharge maneuverability device known as a “bow thruster.” As typically embodied, the inventive nozzle includes a nozzle wall and an arc-shaped vane. The nozzle wall has a circular nozzle inlet and an elliptical nozzle outlet, and is configured so that the geometric major axis of the nozzle outlet is horizontal. The arc-shaped vane: is upwardly bowed in the nozzle wall's transverse direction; joins the nozzle wall along its two horizontally opposite sides in the nozzle wall's longitudinal direction; extends most or all of the longitudinal distance between the nozzle inlet and the nozzle outlet; has a transversely intermediate section that, taken in the transverse direction, is generally uniform in thickness; has two transversely lateral sections that, taken in the transverse direction, gradually thicken toward the two respective joints of the arc-shaped vane with respect to the nozzle wall.

Abstract: Assembly for supporting a gas turbine engine comprises a support member including mounting sites for attachment to an associated aircraft. The support member includes integral slider tracks disposed on opposed sidewalls for selective engagement with a translatable cowl. The translatable cowl is translatable along the slider tracks to selectively cover and expose a cascade thrust-reversing structure. The cascade structure includes arcuate segments mounted in hinged relationship to the support member. Rearward translation of the translatable cowl to a service position permits the cascade structure to be opened for access to the core engine.

Abstract: A system includes a multi-tube fuel nozzle having an inlet plate and a plurality of tubes adjacent the inlet plate. The inlet plate includes a plurality of apertures, and each aperture includes an inlet feature. Each tube of the plurality of tubes is coupled to an aperture of the plurality of apertures. The multi-tube fuel nozzle includes a differential configuration of inlet features among the plurality of tubes.

Abstract: A fossil-fueled power station including a steam turbine is provided. A steam generator is mounted downstream of the steam turbine via a steam return line and a carbon dioxide separation device. The carbon dioxide separation device is connected to the steam return line via a process steam line and a backpressure steam turbine is mounted into the process steam line.

Abstract: One embodiment of the present invention is a unique pulse detonation combustion system. Another embodiment is a unique gas turbine engine including a unique pulse detonation combustion system. In some embodiments, the pulse detonation combustion system includes an inlet section, a vortex generator and at least one flame accelerator. The inlet section, vortex generator and the at least one flame accelerator may be operative to initiate a deflagration to detonation transition. In some embodiments, the pulse detonation combustion system may include a flame accelerator configured with a directionally-dependent drag coefficient. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for pulse detonation combustion systems, gas turbine engines, and other machines and engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: An injector seal for a gas turbomachine includes a seal body having a first surface, an opposing second surface, a central opening configured and disposed to receive a gas turbomachine injector, an outer edge, and a plurality of passages extending through the seal body. The plurality of passages are configured and disposed to guide a fluid at the first surface through the second surface.

Abstract: A fossil-fueled power station including a steam generator is provided. A steam turbine is mounted downstream of the steam generator via a hot intermediate superheater line and a carbon dioxide separation device. The carbon dioxide separation device is connected to the hot intermediate superheater line via a process steam line and a backpressure steam turbine is mounted into the process steam line.

Abstract: Gas-turbine combustion chamber with a combustion chamber head made from a metallic material and mounting at least one burner, and with a combustion chamber wall made from a ceramic material, where at least one igniter plug or other combustion chamber attachment is arranged in a recess of the combustion chamber wall, and where in the area of the recess a seal is arranged that is mounted by a metallic holding mechanism from another component than the combustion chamber wall.

Abstract: In one aspect, an embodiment of the present disclosure provides an Integrated Gasification Combined Cycle (IGCC) apparatus. The apparatus includes a saturator configured to saturate NPG with water vapor, and a heat recovery steam generator (HRSG), a low pressure steam loop through the saturator, wherein the HRSG is configured to heat the low pressure steam loop. The apparatus further includes a compressor and a heat exchanger configured to heat the NPG using waste process heat and extraction air from the compressor, wherein the heated NPG thereby becomes diluent nitrogen.

Abstract: An adaptive vane separator system is provided for use with an inlet filter house of a gas turbine engine system. The vane separator system is coupled to the inlet filter house. The vane separator system includes a slide rail and a plurality of vanes rotatably coupled to the slide rail. The inlet filter house is configured to channel air to an air inlet of a turbine engine. A drive motor is coupled to the vane separator system. The drive motor is operable to selectively move at least one of the vanes to facilitate reducing an amount of moisture channeled through the air inlet.

Abstract: Disclosed is a gas turbine for a thermal power plant, especially a gas turbine, comprising a rotatable rotor and a duct which can be penetrated by a gas and in which turbine blades for driving the rotor are disposed. The inventive gas turbine is characterized by a blocking device which at least partially prevents the gas from penetrating the duct in a closed position.

Abstract: A burner system is provided that includes at least two adjacent burners that are separate from each other. Each of the two burners has at least one combustion chamber and a head end. The head end includes at least a fuel injection and a fuel-air premix. Each burner has a cap with a cap side and cap upper side, wherein at least the cap upper side is arranged ahead of the head end, seen in the direction of flow. In this manner, a burner plenum is formed between the cap upper side and the head end. The at least two burner plenums thus formed have an acoustic connection. A method for damping such a burner system is also provided.

Abstract: A trans cowl for a jet engine includes a chevron coupled with a linear actuators. The chevron is movable by the linear actuator forward or aft to change a gas flow path formed by an core cowl and thrust reverser translating cowl. In a first position, the chevrons are disposed substantially parallel the gas flow path to attenuate drag and/or loss of engine thrust. In a second position, the chevrons are moved aft to project, or further project, into the gas flow path. In one embodiment, the linear actuator comprises a first component that is coupled with the outer cowl. A second component of the linear actuator is coupled with the chevron. When installed, the linear actuator can be coupled with a controller and an electrical power source. A position sensor coupled with the controller senses a position of the linear actuator and/or the chevron.

Abstract: An exhaust nozzle flap for a turbine engine may include an exhaust nozzle flap linkage, an exhaust nozzle flap panel and a mounting pin. The linkage may include a first linkage segment that extends longitudinally from a linkage end to a second linkage segment, and a mounting aperture that extends transversely through the second linkage segment. The panel may include a first panel segment that extends longitudinally from a panel end to a second panel segment. The first panel segment may be pivotally engaged with the first linkage segment. The mounting pin may extend through and move transversely within the mounting aperture, and may be connected to the second panel segment.

Abstract: The burner (1) of a gas turbine includes a swirl generator (2) and, downstream of it, a mixing tube (3). The swirl generator (2) is defined by at least two walls facing one another to define a conical swirl chamber (5) and is provided with nozzles (6) arranged to inject a fuel and apertures (7) arranged to feed an oxidizer into the swirl chamber (5). The burner (1) includes a lance (9) which extends along a longitudinal axis of the swirl generator (2) and is provided with side nozzles (11) for ejecting a fuel within the burner (1). The side nozzles (11) have their axes (12) inclined with respect to the axis of the lance (9) and can be positioned along the axis of the burner.

Abstract: A fan duct (5) is defined between an inner skin (2) and an outer skin into which the inner skin of the sliding cover (7) is incorporated in the direct-thrust configuration. The inner skin of the sliding cover (7) comprises an annular part extending radially inwards and capable, in the reverse-thrust configuration, of coming to face a larger-diameter annular part of the inner skin (2) of the fan duct (5). The annular part of the sliding cover (7) comprises at least one flap (8) mounted such that it can pivot between the retracted position and a deployed position in which a front part (8a) of the flap (8), upstream of its middle axis (9) about which it is articulated, projects into the fan duct (5). Translational stop means (11) are provided in order, in the reverse thrust position, to act on a rear part (8b) of the flap (8) in such a way as to cause the flap (8) to pivot into the deployed position to close off the fan duct (5).