Abstract: The present disclosure is directed to a combustor assembly for a gas turbine engine. The combustor assembly includes a liner defining a combustion chamber therewithin and a pressure plenum surrounding the liner. The liner defines an opening and includes a walled chute disposed at least partially through the opening. A plurality of flow openings is defined through the walled chute.

Abstract: A gas turbine engine includes a main compressor section having a downstream most location, and a turbine section, with both the main compressor section and the turbine section housing rotatable components. A first tap taps air compressed by the main compressor section at an upstream location upstream of the downstream most location. The first tap passes through a heat exchanger, and to a cooling compressor. Air downstream of the cooling compressor is selectively connected to reach at least one of the rotatable components. The cooling compressor is connected to rotate at a speed proportional to a rotational speed in one of the main compressor section and the turbine section. A valve system includes a check valve for selectively blocking flow downstream of the cooling compressor from reaching the at least one rotatable component. A dump valve selectively dumps air downstream of the cooling compressor. A method is also disclosed.

Abstract: A ventilation system includes a cavity, a fluid motive force device, and a motive fluid supply system. The cavity includes different ventilation level requirements for a plurality of modes of operation. The fluid motive force device includes a suction port, an outlet port, and a motive fluid inlet port. The suction port is coupled in flow communication with the cavity to be vented. A flow supply to the motive fluid inlet port determines a ventilation flow through the suction port. The motive fluid supply system is coupled in flow communication with the motive fluid inlet port. An operation of the motive fluid supply system determines a flow of motive fluid from the motive fluid supply system to the motive fluid inlet port. The flow of motive fluid to the motive fluid inlet port generates a ventilation flow through the suction port approximately matching a current ventilation demand of said cavity.

Abstract: A system and method for detecting inlet temperature distortion of an engine are described. The method comprises obtaining an outside air temperature from at least one first sensor, obtaining an inlet temperature of the engine from at least one second sensor, determining an inlet temperature distortion based on a difference between the outside air temperature and the inlet temperature, comparing the inlet temperature distortion to a threshold, and issuing an alert when the inlet temperature distortion exceeds the threshold.

Abstract: A gas turbine engine comprising: a compressor; a first turbine; and a first compressor bleed valve in fluid communication with the compressor and configured to release bleed air from the compressor; wherein the first compressor bleed valve is configured to release bleed air to a downstream location in the engine, the downstream location being downstream of the first turbine; wherein the first compressor bleed valve is configured to open wherein the first compressor bleed valve is configured to open to at least two positions, to thereby release a variable amount of bleed air from the compressor.

Abstract: The invention relates generally to gas turbine engines used for electrical power generation. More specifically, embodiments of the present invention provide systems and ways for improving the life and reducing start-up time necessary for bringing gas turbine engines online and up to full power.

Abstract: A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A liner cooling structure of a duct assembly reduces pressure loss generated in the compressed air flow for cooling the liner. The duct assembly includes a liner, a transition piece, and a flow sleeve, and the transition piece and the flow sleeve form a transition piece channel through which a main stream of compressed air is introduced to the duct assembly. The liner cooling structure includes a first flow passage through which the main stream of compressed air passes in a first direction; and a second flow passage formed as a plurality of inlet holes in the flow sleeve to communicate with the first flow passage and configured to pass an auxiliary stream of compressed air in a second direction from outside the flow sleeve to inside the flow sleeve, the auxiliary stream joining the main stream such that the second direction forms an acute angle with the first direction.

Abstract: An air supply system for a gas turbine engine including a first duct which is connected or capable of being connected to a first compressor bleed air supply of a gas turbine engine; a second duct which downstream of the first compressor bleed air supply is connected or capable of being connected to a second compressor bleed air supply of the gas turbine engine; a nozzle by way of which air from the second duct is capable of being blown into the first duct; and an exhaust air duct having an opening which downstream of the nozzle is disposed in the first duct in such a manner that air blown by way of the nozzle into the first duct can flow out of the first duct through the opening into the exhaust air duct. A gas turbine engine and an aircraft are furthermore provided.

Abstract: A gas turbine engine includes an engine static structure. A rotating structure is configured to rotate relative to the engine static structure. The rotating structure has a target area with first and second directions. The first direction is greater than the second direction. A lubrication system includes a nozzle having a non-circular exit aimed at the target area. The exit provides a width and a height. The width is greater than the height. The width is oriented in the first direction.

Abstract: A fuel spray nozzle for a gas turbine engine, the fuel spray nozzle arranged to mix fuel and air and provide the mixture to a combustor of the engine and including: a first fuel supply conduit arranged to provide fuel to be mixed with air in a first ratio; a second fuel supply conduit arranged to provide fuel to be mixed with air in a second ratio, having a lower proportion of fuel than the first ratio; and a staging valve arranged to control the relative proportions of fuel provided through the first fuel supply conduit and the second fuel supply conduit, such that an increase or decrease in the flow in the first fuel supply conduit is accompanied by a corresponding decrease or increase in the flow in the second fuel supply conduit.

Abstract: Systems and methods for conditioning a fluid using bleed air from a bypass duct of a turbofan engine are disclosed. The system comprises a heat exchanger configured to facilitate heat transfer between a flow of bleed air from the bypass duct of the turbofan engine and the fluid, and a fluid propeller configured to drive the bleed air through the heat exchanger. The fluid propeller is disposed downstream of the heat exchanger.

Abstract: A cooling system for a gas turbine engine comprises a passage capable of receiving cooling air, a compartment radially adjacent thereto and axially aligned therewith, an opening therebetween, a valve within the opening, and a heat exchanger received in the compartment. The valve is moveable between a maximum open position and a minimum open position for increasing or decreasing airflow from the passage into the compartment. At the valve minimum open position, a leakage path is provided between the passage and the compartment, whereby cooling air is capable of passing from the passage to the compartment and toward the heat exchanger at all valve positions. A gas turbine engine is also disclosed.

Abstract: A liner assembly for use in a combustor of a gas turbine engine is disclosed. In various embodiments, the liner assembly includes a panel defining a left side and a right side and a hot side and a cold side, the panel having a dilution hole and a plurality of effusion holes extending between the hot side and the cold side. In various embodiments, the plurality of effusion holes includes a first subgrouping of effusion holes disposed downstream of the dilution hole and aligned in a generally left to right orientation toward a dividing line extending downstream of the dilution hole and a second subgrouping of effusion holes disposed downstream of the dilution hole and aligned in a generally right to left orientation toward the dividing line extending downstream of the dilution hole.

Abstract: A turbine engine includes a core engine including a first spool and a second spool rotatable about a main engine longitudinal axis, a boost spool powered by a secondary drive system, and an accessory gearbox coupled to the core engine and the boost spool. A differential gear system is coupled between the core engine, the boost spool and the accessory gearbox for distributing power between the boost spool, the core engine and the accessory gearbox.

Abstract: An injection assembly for a gas turbine combustor having a liner defining a combustion zone and a secondary combustion zone and a forward casing circumferentially surrounding at least a portion of the liner is provided. The injection assembly includes a thimble assembly and an injector unit. The thimble assembly, which is mounted to the liner, includes a thimble that extends through a thimble aperture in the liner. The injector unit, which is mounted to and extends through the forward casing, includes an injector blade that extends into the thimble. The injection assembly introduces a flow of fuel into a flow of air flowing through the thimble, such that fuel and air are injected into the secondary combustion zone in a direction transverse to a flow of combustion products from the primary combustion zone.

Abstract: An assembly for a gas turbine engine includes a vane assembly having an inner platform and an airfoil section extending from the inner platform. A first combustor panel extends at least partially along the vane from upstream of the vane, and a blade outer air seal extends at least partially along the vane from downstream of the vane. The first combustor panel and the blade outer air seal define a platform gap located between a leading edge of the airfoil section of the vane and a trailing edge of the airfoil section of the vane.

Abstract: A gas turbine includes a compressor; a combustor; a turbine configured to drive a rotational shaft of the compressor using combustion gas generated by the combustor; a cooling device configured to generate cooling air by bleeding compressed air from the compressor and cooling the compressed air, and to supply the cooling air to the turbine along the rotational shaft; a pressurizing device configured to increase pressure of the cooling air; a pressurizing device diffuser configured to provide a passage continuing in a turbine circumferential direction, on the outer side in the turbine radial direction to guide the cooling air having the increased pressure to the outer side of the pressurizing device; and a manifold disposed between the pressurizing device diffuser and a plurality of turbine vanes so that a ring-shaped passage communicates with the passage in the pressurizing device diffuser and a cooling passage provided inside each turbine vane.

Abstract: A bleed valve assembly of a gas turbine engine includes two or more valve segments extending circumferentially around a central longitudinal axis of the gas turbine engine, and a first splice bracket spanning a first joint between a first valve segment and a second valve segment of the two or more valve segments. The first splice is bracket secured to the first valve segment and the second valve segment. A second splice bracket spans a second joint between the first valve segment and the second valve segment. The second splice bracket is secured to the first valve segment and the second valve segment.

Abstract: An aspect includes a hybrid gas turbine engine system of a hybrid electric aircraft. The hybrid gas turbine engine system includes a gas turbine engine, an electric motor operable to perform an electric taxiing of the hybrid electric aircraft, and a controller. The controller is operable to prevent fuel flow to the gas turbine engine during at least a portion of the electric taxiing and monitor for a powered warm-up request during the electric taxiing. A powered warm-up state of the gas turbine engine is initiated based on detecting the powered warm-up request. The powered warm-up state adds heat to one or more components of the gas turbine engine prior to transitioning to a takeoff power state. The gas turbine engine transitions from the powered warm-up state to the takeoff power state after reaching a target temperature of the one or more components in the powered warm-up state.