Abstract: A turbofan engine includes fan section including a plurality of fan blades, a gear train, a low spool including a low pressure turbine and a low pressure compressor, the low pressure turbine driving the plurality of fan blades through the gear train, and a high spool including a high pressure turbine driving a high pressure compressor. A fan nacelle at least partially surrounds a core nacelle to define a fan bypass flow path. A fan variable area nozzle is in communication with the fan bypass flow path and defines a fan nozzle exit area between the fan nacelle and the core nacelle. The fan variable area nozzle varies the fan nozzle exit area.

Abstract: A mixer having a housing, a duct within the housing, a first and a second injector arranged to inject a fluid at a centre zone of the duct, a third and a fourth injector arranged to inject the fluid at a wall zone of the duct. The first/third injectors are at a distance D1=v/2f1 or odd integer multiples of it from the second/fourth injectors in the absence of an acoustic node between them, or at a distance D1=?conv=v/f1 or full wave length integer multiples of it in the presence of an acoustic node between them. Advantageously f1 is greater than f2.

Abstract: There is disclosed an igniter for a gas turbine engine including: a base, a glow plug heater rod extending from the base along an axis and terminating in a rod end, and a plurality of fins disposed circumferentially around the glow plug heater rod, the fins extending axially along at least a portion of the glow plug heater rod. A gas turbine engine including such an igniter is further disclosed.

Abstract: A fuel system is provided for an aircraft having a fuel source. The fuel system includes a fuel oxygen reduction unit defining a liquid fuel supply path, a stripping gas supply path, a liquid fuel outlet path, and a stripping gas return path, wherein the stripping gas return path is in airflow communication with the fuel source.

Abstract: A panel for use with a shell as a combustor liner in a combustor section of a gas turbine engine includes a panel body having an outer surface defining a plurality of effusion holes for receiving the compressed gas to also be received in the combustion chamber of the combustor section. The panel further includes a flow guide extending from the outer surface of the panel body and configured to receive the compressed gas from an impingement hole of the shell and to direct the compressed gas over the outer surface of the panel body towards the plurality of effusion holes.

Abstract: A gas turbine engine component having a cooling passage includes a first wall defining an inlet of the cooling passage, a second wall generally opposite the first wall and defining an outlet of the cooling passage, a metering section extending downstream from the inlet, and a diffusing section extending from the metering section to the outlet. The metering section includes an upstream side and a downstream side generally opposite the upstream side. At least one of the upstream and downstream sides includes a first passage wall and a second passage wall where the first and second passage walls intersect to form a V-shape.

Abstract: There is disclosed an igniter for a gas turbine engine including: a base including a securing assembly configured to be matingly connected to a casing of the gas turbine engine, a glow plug heater rod, the glow plug heater rod extending from the base along an axis and terminating in a rod end, a flow impeding member extending along the axis and having a distal end spaced from the base and extending axially relative to the axis beyond the rod end, the flow impeding member sized and configured such that the distal end extends into a combustion chamber of the gas turbine engine upon the base connected to the casing.

Abstract: There is disclosed an igniter for a gas turbine engine including: a base; an inner conductor extending along an axis from the base to an end; a resistive heating element electrically connected to the end of the conductor; an outer conductor being electrically connected to the heating element, the outer conductor surrounding and extending axially along the inner conductor and forming a return electrical path, the outer conductor being radially spaced apart from the inner conductor by a gap, and an electrical insulator in the gap.

Abstract: There is disclosed an igniter for a gas turbine engine including: a base, a glow plug heater rod, the heater rod extending from the base along an axis and terminating in a rod end, a peripheral wall extending circumferentially around the rod end and having a distal end directed away from the base, a gap between the peripheral wall and the heater rod, the rod end being axially retracted relative the distal end of the peripheral wall.

Abstract: An apparatus and method for diffusing an airflow in a turbine engine can include a turbine center frame positioned between a high pressure turbine and a low pressure turbine of a turbine section of the engine. The turbine center frame can include two or more diffusion sections for diffusing the airflow. A stabilization section can be provided between the two or more diffusion sections to stabilize the airflow.

Abstract: The method allows to detect flameout in a combustor of a turbine system; it includes the steps of: A) measuring angular acceleration of a shaft of the or each turbine of the turbine system, B) calculating a flameout indicator as a function of the angular acceleration, and C) carrying out a comparison between the flameout indicator and at least one threshold.

Abstract: A propulsion device includes a propulsion body and a diversion assembly. The propulsion body includes a propulsion system and a housing accommodating the propulsion system. The housing has an air-intake opening and an air-discharge opening respectively on two opposite sides of the propulsion system. The diversion assembly includes first and second diversion annular sheets. The first diversion annular sheet is disposed outside the air-discharge opening of the housing and having a surrounding center. The first diversion annular sheet is swung relative to the air discharge opening by a first axis passing through the surrounding center. The second diversion annular sheet is disposed outside the air-discharge opening of the housing and concentrically disposed with the first diversion annular sheet. The second diversion annular sheet is swung relative to the air-discharge opening by a second axis passing through the surrounding center, and first axis intersects the second axis.

Abstract: There is disclosed a method of servicing a gas turbine engine having a spark plug igniter connected to a spark plug igniter socket of the engine, the method including: removing the spark plug from the spark plug igniter socket of the gas turbine engine; inserting a glow plug igniter having a base configured for matingly engaging the spark plug igniter socket; and securing the glow plug igniter to the gas turbine engine.

Abstract: There is provided an igniter for a gas turbine engine including: a base, a glow plug heater rod extending from the base along an axis and terminating in a rod end, the heater rod having a heating section extending axially between axially opposite ends of a heater contained within the heater rod, a heat spreader being in thermal conduction contact with the heating section of the heater rod and extending radially therefrom.

Abstract: A system for directing cooling air into a gas turbine combustor is provided. The system comprises a transition duct coupled to a flow sleeve, where air to be used for combustor cooling and in the combustion process enters a bellmouth of the transition duct, passes through a plurality of struts within the bellmouth, and is distributed to a passage located between the combustion liner and flow sleeve.

Abstract: An air turbine starter for starting an engine, comprising a housing defining an inlet, an outlet, and a flow path extending between the inlet and the outlet for communicating a flow of gas there through. A turbine member is journaled within the housing and disposed within the flow path for rotatably extracting mechanical power from the flow of gas. A gear train is drivingly coupled with the turbine member, a drive shaft is operably coupled with the gear train, and an output shaft is selectively operably coupled to rotate with the engine via a decoupler.

Abstract: The premixed pilot nozzle includes axially elongated tubes defined within a plenum between an outer shroud and a first shroud disposed radially inward of the outer shroud. The tubes extend between tube inlets defined through a forward face and tube outlets defined through an aft face. A second shroud is disposed radially inward of the first shroud, thereby defining a fuel plenum between the first shroud and the second shroud, and the fuel plenum is in communication with a gaseous fuel supply. A fuel injection port, which is positioned between the tube inlet and the tube outlet of each tube, is in fluid communication with the fuel plenum. An air supply configured to fluidly communicate with the tube inlet of each tube. The second shroud defines a second plenum therein, the second plenum being coupled to a source of a non-combustible fluid.

Abstract: A control device is a control device for a gas turbine including a plurality of combustors and is configured to select combustors to ignite in accordance with a target load on the basis of a predictor which defines a relationship between a load and the number and arrangement of combustors to ignite and a combustion temperature.

Abstract: A sleeve assembly for a combustion system includes a first duct including a first circumferential surface and a second duct including a second circumferential surface. The first circumferential surface radially overlaps the second circumferential surface such that the first duct and the second duct collectively define a continuous combustion chamber therein. The continuous combustion chamber is configured to receive high temperature gases flowing therethrough. The sleeve assembly further includes a wear insert removably coupled to the first circumferential surface and extending radially therefrom to engage the second circumferential surface.

Abstract: A turbine system includes a compressor section, an inlet cooling system coupled upstream of the compressor section and configured to cool ambient air entering the compressor section, and a turbine section coupled in flow communication with the compressor section and including at least one hot gas path component. The system further includes a controller configured to receive feedback parameters indicative of a temperature of the at least one hot gas path component, estimate a remaining life of the at least one hot gas path component based on the received feedback parameters, determine a desired power output of the turbine system based on the estimated remaining life of the at least one hot gas path component and a cooling capacity of the inlet cooling system, and control operation of the turbine system to cause the turbine system to generate the desired power output.