1427results about "Turbine/propulsion engine ignition" patented technology

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 gas turbine generating apparatus according to the present invention comprises a gas turbine capable of a high-speed operation, a permanent-magnet-type generator driven at a high speed by the gas turbine, an inverter device (5) for converting alternating-current power generated by the generator into commercial alternating-current power, and a system-interconnection device (9) for interconnecting output of the inverter device to a commercial AC power supply system (10). The system-interconnection device controls the inverter device based on voltage of the commercial AC power supply system as a criterion such that output current of the inverter device is in phase with the voltage or out of phase with the voltage by a constant phase difference. A control device (18) is provided for operating the permanent-magnet-type generator (4) as a motor at a time of starting the gas turbine, and supplying alternating-current power from a direct-current power supply through the inverter device or another inverter device in addition to the inverter device to the generator for increasing a rotational speed smoothly from zero. The control device detects output current of the inverter device and controls the inverter device to keep the output current constant.

A combustor (22) for a gas turbine (10) includes a main burner oxidizer flow path (34) delivering a first portion (32) of an oxidizer flow (e.g., 16) to a main burner (28) of the combustor and a pilot oxidizer flow path (38) delivering a second portion (36) of the oxidizer flow to a pilot (30) of the combustor. The combustor also includes a flow controller (42) disposed in the pilot oxidizer flow path for controlling an amount of the second portion delivered to the pilot.

An engine power extraction control system controls the main propulsion engines and the electrical machines that are coupled thereto to supply an appropriate amount of aircraft thrust and electrical energy to the aircraft. The engines and electrical machines are also controlled so that the propulsion thrust that is generated is split between the various turbines in the main propulsion engines to maintain an adequate surge margin and to minimize residual thrust generation.

A premix chamber for a combustor of a gas turbine engine comprises a cylindrical chamber having a premix chamber wall, the cylindrical chamber having a chamber inlet end longitudinally separated from a chamber outlet end along a central axis, a chamber inlet plate in communication with the premix chamber wall at the chamber inlet end, the chamber inlet plate having a fuel nozzle inlet hole disposed through the chamber inlet plate, the chamber inlet plate further comprising a plurality of swirler passages disposed through the chamber inlet plate, and the chamber outlet end being open. A method of producing turbine gas is also disclosed.

A trapped vortex cavity afterburner includes one or more trapped vortex cavity stages for injecting a fuel / air mixture into a combustion zone. The trapped vortex cavity afterburner is operable to provide all thrust augmenting fuel used for engine thrust augmentation. Each stage has at least one annular trapped vortex cavity. The trapped vortex cavity afterburner may be a multi-stage afterburner having two or more trapped vortex cavity stages ganged for simultaneous ignition or operable for sequential ignition. One embodiment of the annular trapped vortex cavity is operable to raise a temperature of an exhaust gas flow through the afterburner about 100 to 200 degrees Fahrenheit. Each of the trapped vortex cavity stages may be operable to produce a single or a different amount of temperature rise of the exhaust gas flow through the afterburner. A chevron shaped trapped vortex cavity and having zig-zag shaped leading and trailing edges may be used.

A combustor assembly includes a combustor chamber having a primary and intermediate zone that provides for reduced flame temperatures. The combustor assembly includes first and second pluralities of injectors. The first plurality of injectors introduces fuel to a primary zone. A second plurality of injectors introduces fuel to an intermediate zone. During operation between initial start up and before the introduction of engine load, fuel is introduced into the primary zone only by the first plurality of injectors. Once engine load is applied to the engine, fuel is introduced into the intermediate zone by the second plurality of injectors. Introduction of additional volume of fuel allows the fuel-air ratio to remain constant regardless of engine operating conditions. The constant fuel-air ratio is maintained at a desired rate to lower flame temperatures and reduce nitrous oxide emissions.

A method of operating a gas turbine engine is provided. The method includes discharging pilot fuel into a combustion chamber from a nozzle through at least one pilot fuel outlet defined in a tip of the nozzle, discharging steam from the nozzle through a plurality of steam outlets that are spaced circumferentially about the plurality of pilot fuel outlets, and discharging primary fuel from the nozzle through a plurality of primary fuel outlets that are circumferentially aligned with the plurality of steam outlets. To facilitate mixing the primary fuel with the steam, the primary fuel is discharged from the nozzle tip at an oblique angle with respect to a centerline extending through the nozzle tip.

A gas turbine combustion system for burning air and fuel into exhaust gases. The gas turbine combustion system may include a combustor, a turbine nozzle integral with the combustor for providing the air to the combustor, and a fuel injector for providing the fuel to the combustor. The turbine nozzle and the fuel injector are positioned within the combustor such that a mixture of the air and the fuel flows in a first direction and the exhaust gases flow in a second direction.

Apparatus and method for detecting incipient lean blowoff conditions in a lean premixed combustion nozzle of a gas turbine. A sensor near the flame detects the concentration of hydrocarbon ions and / or electrons produced by combustion and the concentration monitored as a function of time are used to indicate incipient lean blowoff conditions.

A method and apparatus for distributing control of multiple engines amongst the engines in a power generation system that has a central controller, comprises the steps of: receiving aboard each engine from the central controller a control signal representative of a desired fuel flow; metering fuel at a fuel metering point aboard each engine; sensing at least one parameter aboard each engine proximate the fuel metering point that is representative of fuel flow; adjusting the fuel metering aboard each engine to cause the sensed parameter to correlate to the desired fuel flow; and transmitting a monitoring signal from each engine to the central controller that is representative of the adjusted fuel metering.