Patent application number | Description | Published |
20090293488 | Combustor - A gas turbine engine combustor has forward bulkhead extending between inboard and outboard walls and cooperating therewith to define a combustor interior volume or combustion chamber. At least one of the walls has an exterior shell and an interior shell including a number of panels. Each panel has interior and exterior surfaces and a perimeter having leading and trailing edges and first and second lateral edges. A number of cooling passageways have inlets on the panel exterior surface and outlets on the panel interior surface. The shell has a plurality of holes for directing air to a space between the shell and heat shield and adapted for preferentially directing said air toward leading edge portions of first stage vanes of a turbine section. | 12-03-2009 |
20110052381 | COMBUSTOR TURBINE INTERFACE FOR A GAS TURBINE ENGINE - A turbine vane downstream of a combustor section includes an arcuate outer vane platform defined about an axis, the arcuate outer vane platform includes a segment of the arcuate outer vane platform along the axis which follows an outer combustor liner panel structure and an arcuate inner vane platform defined about the axis, the arcuate inner vane platform includes a segment of the arcuate inner vane platform along the axis which follows an inner combustor liner panel structure. | 03-03-2011 |
20110120134 | GAS TURBINE COMBUSTOR - An annular combustor for a gas turbine has a combustion chamber having an interior volume that, in longitudinal section, includes a forward volume, an intermediate volume and an aft volume. The forward volume represents from about 30% to about 40% of the combustor interior volume, the intermediate volume represents from about 10% to about 20% of the combustor interior volume, and the aft volume represents from about 40% to about 60% of the combustor interior volume. | 05-26-2011 |
20110126543 | COMBUSTOR PANEL ARRANGEMENT - A combustor module for a gas turbine engine is provided that includes a first annular liner assembly extending along a longitudinal axis of the engine. The first annular liner assembly includes a first annular support shell and a plurality of first heat shield panels coupled to the first annular support shell. The first heat shield panels form a segmented ring defining a plurality of first axial seams therebetween. The combustor module further includes a bulkhead coupled to the first annular liner assembly. The bulkhead provides a plurality of fuel nozzles for passing a first mass flow comprising fuel and air. The combustor module further includes a second annular liner assembly coupled to the bulkhead. The second annular liner assembly is in spaced-apart generally coaxial relationship from the first annular liner assembly by a channel height H. The second annular liner assembly includes an air admittance hole having a mean diameter D extending along a hole axis. The hole axis is offset from the first axial seam defined by the first heat shield panels. | 06-02-2011 |
20120186259 | FUEL INJECTOR ASSEMBLY - A fuel injector assembly for a combustor is provided, including a fuel nozzle having an axial inflow swirler and one or more radial inflow swirlers spaced radially outward of the downstream end of the fuel nozzle and mounted to the combustor, wherein the airstreams produced by the swirlers airblast atomize fuel films produced by the fuel nozzle. | 07-26-2012 |
20120186262 | GAS TURBINE COMBUSTOR AND METHOD FOR OPERATING - A combustor for a gas turbine engine includes a radially inboard liner, a radially outboard liner, and a bulkhead that cooperatively define an annular combustion chamber, a plurality of first fuel injectors that are disposed in the bulkhead, and a plurality of second fuel injectors that are disposed in at least one of the inboard liner and the outboard liner aftward of the bulkhead. A method is also provided for operating the combustor of the gas turbine engine wherein fuel distribution between the forward combustion zone and the downstream combustion zone is selectively varied in response to the power operating mode of the gas turbine engine with an objective to control NOx formation. | 07-26-2012 |
20120186264 | GAS TURBINE COMBUSTOR - A combustor for a gas turbine engine having an annular combustion chamber includes a plurality of main fuel injection and air swirler assemblies and a plurality of pilot fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of a forward bulkhead. The plurality of pilot fuel injection and air swirler assemblies are interspersed amongst and disposed in the circumferential ring of main fuel injection and air swirler assemblies. Fuel being supplied to the combustor is selectively distributed between the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies in response to the level of power demand on the gas turbine engine. | 07-26-2012 |
20130025293 | TEMPERATURE MIXING ENHANCEMENT WITH LOCALLY CO-SWIRLING QUENCH JET PATTERN FOR GAS TURBINE ENGINE COMBUSTOR - A combustor for a turbine engine includes a first liner defined about an axis with a first row of first combustion air holes, one of the first combustion air holes is defined along each of a multiple of fuel injector zero pitch lines. A second liner defined about the axis with a second row of second combustion air holes, each of the second combustion air holes circumferentially offset relative to each of the multiple of fuel injector zero pitch lines. | 01-31-2013 |
20130125556 | FUEL-AIR MIXTURE DISTRIBUTION FOR GAS TURBINE ENGINE COMBUSTORS - A method of noise control from a combustor of a gas turbine engine includes selectively forming a plurality of local circumferential zones with different fuel-air ratios within the combustor. | 05-23-2013 |
20130186099 | Gas Turbine Combustor - A combustor for a gas turbine engine having an annular combustion chamber includes a plurality of main fuel injection and air swirler assemblies and a plurality of pilot fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of a forward bulkhead. The plurality of pilot fuel injection and air swirler assemblies are interspersed amongst and disposed in the circumferential ring of main fuel injection and air swirler assemblies. Fuel being supplied to the combustor is selectively distributed between the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies in response to the level of power demand on the gas turbine engine. | 07-25-2013 |
20130199200 | FUEL DISTRIBUTION WITHIN A GAS TURBINE ENGINE COMBUSTOR - A fuel system for a gas turbine engine includes a plurality of duplex nozzles arranged on each side of top dead center and a plurality of simplex nozzles. A primary manifold is operable to communicate fuel to a primary flow jet in each of the plurality of duplex nozzles and a secondary manifold is operable to communicate fuel to a secondary flow jet in each of the plurality of duplex nozzles and a secondary flow jet in each of the plurality of simplex nozzles. An equalizer valve that is in communication with both the primary manifold and the secondary manifold distributes fuel at various pressures to both the primary and secondary manifolds. | 08-08-2013 |
20130340437 | TURBINE ENGINE COMBUSTOR WALL WITH NON-UNIFORM DISTRIBUTION OF EFFUSION APERTURES - A turbine engine combustor wall includes support shell and a heat shield. The support shell includes shell quench apertures, first impingement apertures, and second impingement apertures. The combustor heat shield includes shield quench apertures fluidly coupled with the shell quench apertures, first effusion apertures fluidly coupled with the first impingement apertures, and second effusion apertures fluidly coupled with the second impingement apertures. The shield quench apertures and the first effusion apertures are configured in a first axial region of the heat shield, and the second effusion apertures are configured in a second axial region of the heat shield located axially between the first axial region and a downstream end of the heat shield. A density of the first effusion apertures in the first axial region is greater than a density of the second effusion apertures in the second axial region. | 12-26-2013 |
20140083112 | Cooled Combustor Liner Grommet - A combustor liner grommet is disclosed. The grommet may include a peripheral wall defining a hole in a combustor liner and further including at least one cooling air flow channel. The cooling air flow channel in the grommet wall may be a slot or a hole. The channel may increase cooling flow to the grommet and the combustor liner around the grommet to prevent cracking from heat stress. | 03-27-2014 |
20140090394 | FLOW MODIFIER FOR COMBUSTOR FUEL NOZZLE TIP - A fuel injector nozzle assembly includes a body extending along an axis and a core swirl plug positioned at least partially within the body. The core swirl plug has a flow modifying structure configured to swirl fuel at a location upstream from a distal end of the nozzle assembly. | 04-03-2014 |
20140216044 | GAS TURBINE ENGINE COMBUSTOR HEAT SHIELD WITH INCREASED FILM COOLING EFFECTIVENESS - A heat shield for a gas turbine engine includes a hot side with one or more raised features that extend therefrom. | 08-07-2014 |
20140338351 | GAS TURBINE ENGINE COMBUSTOR WITH TAILORED TEMPERATURE PROFILE - A method of tailoring a combustor flow for a gas turbine engine includes controlling an airflow into a swirler to be generally uniform and controlling an airflow into a quench zone to provide a desired pattern factor. | 11-20-2014 |