22677 results about "Gas turbines" patented technology

A shaft for a gas turbine engine is provided that includes a first shaft section, a second shaft section, a first flexible linkage, and a second flexible linkage. The first shaft section extends between a forward axial end and an aft axial end along a first axial centerline. The second shaft section extends between a forward axial end and an aft axial end along a second axial centerline. The first flexible linkage includes a bridge section connected between a first diaphragm and a second diaphragm. The first diaphragm is connected to the aft axial end of the first shaft section. The second diaphragm is connected to the forward axial end of the second shaft section. The second flexible linkage includes a diaphragm and a hub. The second flexible linkage diaphragm cantilevers radially outwardly from an inner radial end to an outer radial end, and is connected to the aft axial end of the second shaft section. The hub is connected to the outer radial end of the second flexible linkage diaphragm, and includes an engine shaft coupling connected to the hub.

A wall in a gas turbine engine includes inner and outer surfaces having a row of compound chevron film cooling holes extending therethrough. The chevron holes diverge both longitudinally and laterally between an inlet at the wall inner surface and a chevron outlet at the wall outer surface.

A method of assembling a gas turbine assembly includes providing a core gas turbine engine including a high-pressure compressor, a combustor, and a turbine, coupling a low-pressure turbine axially aft from the core gas turbine engine, coupling a fan assembly axially forward from the core gas turbine engine, and coupling a booster compressor to the low-pressure turbine such that the booster compressor and the low-pressure turbine rotate at a first rotational speed.

A hybrid structure that combines characteristics of the DACRS and Swozzle burners to provide the high mixing ability of an axial flowing counter rotating vane swirler with good dynamic flame stability characteristics of a bluff center body.

A method and apparatus for cooling a combustor liner and transitions piece of a gas turbine include a combustor liner with a plurality of circular ring turbulators arranged in an array axially along a length defining a length of the combustor liner and located on an outer surface thereof; a first flow sleeve surrounding the combustor liner with a first flow annulus therebetween including a plurality of axial channels (C) extending over a portion of an aft end portion of the liner parallel to each other, the cross-sectional area of each channel either constant or varying along the length of the channel, the first flow sleeve having a plurality of rows of cooling holes formed about a circumference of the first flow sleeve for directing cooling air from the compressor discharge into the first flow annulus; a transition piece connected to the combustor liner and adapted to carry hot combustion gases to a stage of the turbine; a second flow sleeve surrounding the transition piece a second plurality of rows of cooling apertures for directing cooling air into a second flow annulus between the second flow sleeve and the transition piece; wherein the first plurality of cooling holes and second plurality of cooling apertures are each configured with an effective area to distribute less than 50% of compressor discharge air to the first flow sleeve and mix with cooling air from the second flow annulus.

A recuperated gas turbine engine system and associated method employing catalytic combustion, wherein the combustor inlet temperature can be controlled to remain above the minimum required catalyst operating temperature at a wide range of operating conditions from full-load to part-load and from hot-day to cold-day conditions. The fuel is passed through the compressor along with the air and a portion of the exhaust gases from the turbine. The recirculated exhaust gas flow rate is controlled to control combustor inlet temperature.

A gas turbine engine includes an epicyclic gear train that drives a turbo fan. The epicyclic gear train employs a one-piece carrier in which the spaced side walls are interconnected with circumferentially spaced apart mounts to form a unitary structure. Baffles are secured to the carrier near the mounts and provide lubrication passages that spray oil onto the sun gear and / or intermediate gears arranged between the baffles. The baffles can be constructed from a different material than the carrier since the baffles are not structural components in the gear train.

Embodiments of systems and methods for tuning a turbine are provided. In one embodiment, a method may include receiving at least one of a measured operating parameter or a modeled operating parameter of a turbine during operation; and tuning the turbine during operation. The turbine may be tuned during operation by applying the measured operating parameter or modeled operating parameter or parameters to at least one operational boundary model, applying the measured operating parameter or modeled operating parameter or parameters to at least one scheduling algorithm, comparing the output of the operational boundary model or models to the at output of the scheduling algorithm or algorithms to determine at least one error term, and closing loop on the one error term or terms by adjusting at least one turbine control effector during operation of the turbine.

A composite thermal barrier coating system includes a honeycomb metallic structure filled with high thermal expansion ceramic hollow spheres in a phosphate bonded matrix. The composite thermal barrier coating system may be manufactured to thicknesses in excess of current thermal barrier coating systems, thereby imparting greater thermal protection. Superior erosion resistance and abrasion properties are also achieved. The composite thermal barrier coating is useful on combustion turbine components such as ring seal segments, vane segment shrouds, transitions and combustors.

A variable geometry inlet guide vane for a gas turbine aircraft engine includes an aerodynamic shell for turning inlet flow to a turbine or compressor and an internal spar spaced from the airfoil shell by an air gap. A number of actuation mechanisms grounded to the spar and connected to the inner surface of the aerodynamic shell adjust the shape of the shell in response to varying operating conditions of the engine, imbalanced aerodynamic loading of the shell or vibration or other transient loads on the shell.

Provided is an accessory gearbox for a gas turbine engine that contributes minimally to engine weight and complexity. An accessory gearbox comprises a stationary frame that is circumferentially disposed about a central longitudinal axis of the engine. The frame includes a forward-directed axial face and a rearward-directed axial face for fastening the frame to the engine. A central bore accepts a rotational driving means and an outermost rim circumscribes the outer diameter of the frame. A compartment is located between the bore and the rim. At least one primary accessory is mounted to a face and proximate the compartment, remotely engaging the driving means and able to be driven concurrently therewith.