| Patent application number | Description | Published |
|---|
| 20090263243 | Combustion Turbine Including a Diffuser Section with Cooling Fluid Passageways and Associated Methods - A combustion turbine includes a compressor section, a combustion section downstream from the compressor section, and a turbine section downstream from the combustion section. A diffuser section is downstream from the turbine section and has an outer wall, an inner wall, and at least one strut member extending therebetween. The outer wall has at least one first gas passageway therein, the inner wall has at least one second gas passageway therein, and the at least one strut member has at least one third gas passageway therein. The at least one first, second and third gas passageways deliver gas therethrough to assist attachment of a boundary layer to adjacent surfaces of the outer wall, the inner wall, and the at least one strut, respectively. | 10-22-2009 |
| 20100074733 | Ingestion Resistant Seal Assembly - A seal assembly limits gas leakage from a hot gas path to one or more disc cavities in a gas turbine engine. The seal assembly includes a seal apparatus associated with a blade structure including a row of airfoils. The seal apparatus includes an annular inner shroud associated with adjacent stationary components, a wing member, and a first wing flange. The wing member extends axially from the blade structure toward the annular inner shroud. The first wing flange extends radially outwardly from the wing member toward the annular inner shroud. A plurality of regions including one or more recirculation zones are defined between the blade structure and the annular inner shroud that recirculate working gas therein back toward the hot gas path. | 03-25-2010 |
| 20100074734 | Turbine Seal Assembly - A seal assembly that limits gas leakage from a hot gas path to one or more disc cavities in a turbine engine. The seal assembly includes a seal apparatus that limits gas leakage from the hot gas path to a respective one of the disc cavities. The seal apparatus comprises a plurality of blade members rotatable with a blade structure. The blade members are associated with the blade structure and extend toward adjacent stationary components. Each blade member includes a leading edge and a trailing edge, the leading edge of each blade member being located circumferentially in front of the blade member's corresponding trailing edge in a direction of rotation of the turbine rotor. The blade members are arranged such that a space having a component in a circumferential direction is defined between adjacent circumferentially spaced blade members. | 03-25-2010 |
| 20100175387 | Cooling of Turbine Components Using Combustor Shell Air - A turbine engine assembly for a generator including a turbine engine having a compressor section, a combustor section and a turbine section. An air bleed line is in communication with the combustor section for receiving combustor shell air from the combustor section and conveying the combustor shell air as bleed air to a plurality of stages of the turbine section. Bleed air is controlled to flow through the air bleed line when an operating load of the turbine engine assembly is less than a base load of the engine to bypass air exiting the compressor section around a combustor in the combustor section and effect a flow of high pressure combustor shell air to the stages of the turbine section. | 07-15-2010 |
| 20100251689 | MULTIPLE STAGE GAS TURBINE ENGINE - A gas turbine engine ( | 10-07-2010 |
| 20100251727 | Engine brake for part load CO reduction - Aspects of the invention relate to a system and method for operating a turbine engine assembly. The turbine engine assembly has a turbine engine having a compressor section, a combustor section and a turbine section. The combustor section has a lower T_PZ limit and the turbine engine has a design load. The assembly further includes at least one air bleed line from the compressor and at least one valve for controlling air flow through the bleed line. Control structure is provided for opening the valve to allow bleed air to flow through the bleed line when an operating load is less than the design load. The flow rate through the bleed line is increased as the operating load is decreased, reducing the power delivered by the turbine assembly while maintaining the T_PZ above a lower T_PZ limit. A method for operating a turbine engine assembly is also disclosed. | 10-07-2010 |
| 20110302924 | COOLED CONDUIT FOR CONVEYING COMBUSTION GASES - A conduit through which hot combustion gases pass in a gas turbine engine. The conduit includes a wall structure having a central axis and defining an inner volume of the conduit for permitting hot combustion gases to pass through the conduit. The wall structure includes a forward end, an aft end axially spaced from the forward end, the aft end defining a combustion gas outlet for the hot combustion gases passing through the conduit, and a plurality of generally radially outwardly extending protuberances formed in the wall structure. The protuberances each include at least one cooling fluid passage formed therethrough for permitting cooling fluid to enter the inner volume. At least one of the protuberances is shaped so as to cause cooling fluid passing through it to diverge in a circumferential direction as it enters into the inner volume. | 12-15-2011 |
