A gas turbine engine includes a buffer system that can communicate buffer supply air to a portion of the gas turbine engine The buffer system can include a first circuit and a second circuit The first circuit selects between a first bleed air supply having a first pressure and a second bleed air supply having a second pressure that is greater than the first pressure to render a first buffer supply air having an intermediate pressure The second circuit selects between a third bleed air supply and a fourth bleed air supply to communicate a second buffer supply air

Gas turbine engine buffer system

A gas turbine engine includes a heat exchanger, a mid-turbine frame, a passageway that extends through at least a portion of the mid-turbine frame and a first nozzle assembly. The heat exchanger exchanges heat with a bleed airflow to provide a conditioned airflow. The mid-turbine frame is in fluid communication with the heat exchanger. The conditioned airflow is communicated through the passageway and is received by the first nozzle assembly to condition gas turbine engine hardware.

Gas turbine engine buffer cooling system

The invention relates to a gas turbine assembly which substantially consists of at least one compressor, at least one first burner (8, 201), at least one second burner (9, 202) that is connected downstream of the first burner, and at least one turbine (5) that is connected downstream of the second burner. At least the first and second burner (8, 9) form a component of a tubular or quasi-tubular combustion chamber element (100, 200, 300) in the flow direction of the combustion path of said burners, said combustion chamber element being closed or quasi-closed and extending between the compressor (2) and the turbine (5). The combustion chamber elements are arranged around the rotor of the gas turbine assembly in the shape of a ring.

Gas turbine assembly and corresponding operating method

A vane carrier temperature control system for use in a gas turbine engine includes a first cooling air source, a second cooling air source, and an air temperature control system. The first cooling air source supplies a first portion of vane carrier cooling air extracted from a compressor section of the engine to a first section of a vane carrier that supports a plurality of rows of vanes within a turbine section of the engine. The second cooling air source supplies a second portion of vane carrier cooling air extracted from the compressor section to a second section of the vane carrier spaced from the first section in an axial direction defined by a direction of hot working gas flow through the turbine section. The air temperature control system controls a temperature of at least one of the first and second portions of vane carrier cooling air.

Vane carrier temperature control system in a gas turbine engine

A turbine cooling air generation system for a gas turbine engine includes a first fluid pathway connecting a compressor bleed outlet and a mixing valve, a second fluid pathway connecting the compressor bleed outlet and an input of a heat exchanger, and a third fluid pathway connecting an output of the heat exchanger and the mixing valve. The mixing valve is further connected to an input of a turbine stage active cooling system.

Turbine stage cooling

A turbine engine assembly for a generator including a turbine engine having a compressor section (12), a combustor section (30) and a turbine section (14). 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.

Cooling of turbine components using combustor shell air

A starting process for a gas turbine (28) that holds the turbine speed of rotation at an ignition speed setting during an ignition window (58), with the ignition speed setting being based on ambient air conditions to achieve a specified combustor air mass flow rate (52). A fuel flow rate may be set based on the fuel type and temperature to achieve a particular air/fuel ratio in a combustor. The fuel flow rate may be adjusted during the ignition window and thereafter based on a combustor inlet air temperature (46). Completion of ignition may be determined by a reduction (68) in a blade path temperature spread (66). After ignition, fuel flow is increased to accelerate the turbine to full speed. At any point, the fuel flow may be reduced, or its increase may be slowed, to avoid exceeding a temperature limit in the turbine.

Gas turbine starting process

Aspects of the invention relate to a system and method for actively managing blade tip clearances in a turbine engine, particularly under start up and steady state operating conditions. During start up, the rotor can be cooled while the blade ring can be heated. During steady state operation, the rotor may be heated while the blade ring may be cooled. The sources of the heating fluid and cooling fluid can be compressor delivery air, boiler steam and ambient air. The fluids can be compressed to maintain efficiency.

Tip clearance control

An ignition system for a combustor of a gas turbine engine is disclosed. The ignition system may include multiple igniters to reduce the likelihood of system failure. In at least one embodiment, the ignition system may be formed from a first igniter formed from an outer electrode containing at least one central electrode at a distal end of the outer electrode and may include a second igniter. The second igniter may be formed from a central electrode positioned within the outer electrode forming the first igniter. During operation, the igniters may spark to ignite the fuel mixture within the combustor

Multi-electrode igniter for a gas turbine engine

An interface (10) is provided between a combustor basket (12) and a transition (14) of a gas turbine engine (16). The interface (10) includes a tongue (18) formed in one of the combustor basket (12) and the transition (14), and a groove (20) formed in another of the combustor basket (12) and the transition (14). The tongue (18) is received within the groove (20) to define the interface (10). A sealing arrangement is also provided and includes an annular combustion outlet (12′″) and an annular transition inlet (14′″), with a labyrinth seal (72′″) positioned between an end (68′″) of the annular combustion outlet and an end (70′″) of the annular transition inlet to form a labyrinth path (34′″) there between. An inner diameter (24′″) of the annular combustion outlet (12′″) is axially aligned with an inner diameter (22′″) of the annular transition outlet (14′″).

Adam M. Foust

Papers

Cooling of turbine components using combustor shell air

Gas turbine starting process

Tip clearance control

Multi-electrode igniter for a gas turbine engine

Interface between a combustor basket and a transition of a gas turbine engine