Abstract: A gas turbine engine includes a plurality of rotatable components housed within a main compressor section and a turbine section. A cooling system is connected to tap air from said main compressor section. A first tap is connected to a first heat exchanger. The first heat exchanger is connected to a cooling compressor for raising a pressure of the tapped air downstream of the first heat exchanger. A second heat exchanger is downstream of the cooling compressor, and a connection is downstream of the second heat exchanger for delivering air to a bearing compartment. A connection intermediate the cooling compressor and the second heat exchanger delivers cooling air to at least one of the rotatable components.

Abstract: A propulsion system for an aircraft includes at least two gas turbine engines and at least one auxiliary propulsion fan. The at least one auxiliary propulsion fan is configured to selectively receive a motive force from either or both of the at least two gas turbine engines through at least one shaft operatively coupled to the at least one auxiliary propulsion fan.

Abstract: A gas turbine engine has a fan rotor, a turbine rotor driving the fan rotor, and an epicyclic gear reduction positioned between the fan rotor and the turbine rotor. The epicyclic gear reduction includes a ring gear, a sun gear, and no more than four intermediate gears that engage the sun gear and the ring gear. The fan drive turbine is configured to drive the sun gear to, in turn, drive the ring gears to, in turn, drive the fan rotor.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including at least a first compressor section and a second compressor section. A power ratio is provided by the combination of the first compressor section and the second compressor section. A method of design a gas turbine engine is also disclosed.

Abstract: A gas turbine engine includes a main compressor section and a turbine section. The turbine section has a first turbine blade and vane and a downstream turbine component. A tap is configured to tap air from the compressor section at a location upstream of a most downstream location. The tap is connected to a heat exchanger. The heat exchanger is connected to a cooling compressor. The cooling compressor is connected to the downstream turbine component. A second tap is configured to tap air from a location in the main compressor section. The second tap is connected through a check valve to a line leading to the downstream turbine component.

Abstract: A turbofan engine includes a fan section including a fan blade having a leading edge and hub to tip ratio of less than about 0.34 and greater than about 0.020 measured at the leading edge and a speed change mechanism with gear ratio greater than about 2.6 to 1. A first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.

Abstract: A gas turbine engine comprises at least two compressor rotors, including a first lower pressure compressor rotor and a second higher pressure compressor rotor. At least two corresponding air taps include a low tap for tapping low pressure compressor air from a location downstream of a first stage of the lower pressure compressor rotor, and upstream of a first stage of the higher pressure compressor rotor, and a high tap to tap air downstream of the first stage of the higher pressure compressor rotor. an air handling system selectively communicates both the low tap and the high tap to an air use destination. Air is selectively supplied from the low tap to the air handling system at a high power operation and from the high tap to the air handling system at a low power operation.

Abstract: A high pressure compressor has a downstream most end. A housing surrounds the compressor section and a turbine section. A low pressure turbine has a downstream most end. A first tap selectively taps high pressure cooling air from a location downstream of the downstream most end in the high pressure compressor and passes the high pressure cooling air through a heat exchanger. A second tap taps compressed air from a location upstream of the downstream most end in the high pressure compressor, and passes air over the heat exchanger, cooling the high pressure cooling air. A chamber is defined between the core engine housing and a nacelle airflow wall, and the second tap air flows through the chamber. The second tap air moves from the chamber into a core engine flow at a location downstream of the downstream most end of the low pressure turbine.

Abstract: A gas turbine engine having an engine axis and method of manufacturing the same is disclosed. The gas turbine engine may comprise a fan configured to drive air, a low pressure compressor section having a core flow path and configured to draw in and compress air flowing along the core flow path, a spool configured to drive the fan, and geared architecture configured to adjust the fan speed. The gas turbine engine may also include a housing defining a compartment that encloses the geared architecture. The housing is disposed between the core flow path and the axis, and includes a shielded mid-section that is in thermal communication with the core flow path of the low pressure compressor section. The shielded mid-section includes an outer layer and an insulator adjacent to the outer layer.

Abstract: A method of modulating cooling of gas turbine engine components includes the steps of identifying an input indicative of a usage rate for at least a first gas turbine engine component of a plurality of gas turbine engine components. A cooling system is operated for at least the first gas turbine engine component. The cooling system is moved between a higher cooling potential mode and a lower cooling potential mode based on the identified rate. A gas turbine engine is also disclosed.

Abstract: A gas turbine engine includes a plurality of rotatable components housed within a compressor section and a turbine section. A tap is connected to a location upstream of a downstream most location in the compressor section. The tap is connected to a heat exchanger. Downstream of the heat exchanger is a shut off valve and downstream of the shut off valve is a cooling compressor. The cooling compressor is connected to deliver cooling air through a chamber, and then to at least one of the plurality of rotatable components. The chamber is provided with at least one check valve configured to selectively allow flow directly from a more downstream location in the compressor section than the location upstream. The flow from the more downstream location has a higher pressure than a flow from the location upstream. There is a system for stopping operation of the cooling compressor. There is a control for closing the shut off valve.

Abstract: A gas turbine engine includes, among other things, a fan section including a fan rotor, a gear train defined about an engine axis of rotation, a first nacelle which at least partially surrounds a second nacelle and the fan rotor, the fan section configured to communicate airflow into the first nacelle and the second nacelle, a first turbine, and a second turbine followed by the first turbine. The first turbine is configured to drive the fan rotor through the gear train. A static structure includes a first engine mount location and a second engine mount location.

Abstract: An airfoil for a turbine engine includes an airfoil body with a cover mounting support, a first cover, and a second cover. The airfoil body includes a solid perimeter portion surrounding a recess formed into at least one of a suction side and a pressure side of the airfoil body, while the cover mounting support extends through the recess. The first cover can be engaged with a first edge of the recess and joined to a first portion of the cover mounting support by a first stir weld. The second cover can be engaged with a second edge of the recess, and joined to a second portion of the cover mounting support by a second stir weld.

Abstract: A bowed rotor prevention system for a gas turbine engine of an aircraft is provided. The bowed rotor prevention system includes a bowed rotor prevention motor operable to drive rotation of a starting spool of the gas turbine engine through an engine accessory gearbox. The bowed rotor prevention system also includes a controller operable to engage the bowed rotor prevention motor and drive rotation of the starting spool until a bowed rotor prevention threshold condition is met or a bowed rotor prevention shutdown request to halt rotation of the starting spool is received.

Abstract: An aspect includes a system including a high compressor of a gas turbine engine having a ratio of a cold-rotor build clearance to a span between 0.7% and 7%. The cold-rotor build clearance is defined for a plurality of rotor blades of the high compressor with respect to an engine casing assembly interior surface of the high compressor, and the span is defined as a gap between a rotor disk of the high compressor and the engine casing assembly interior surface of the high compressor for at least a last two stages of the high compressor closest to a combustor section of the gas turbine engine. The system also includes at least two bowed rotor management systems for the gas turbine engine to prevent damage to the rotor blades for a bowed rotor condition of the high compressor under a plurality of operating conditions.

Abstract: A gas turbine engine has a core engine with a compressor section and a turbine section. The compressor section includes a low pressure compressor and a high pressure compressor. A cooling air system taps compressed air and passes the compressed air through a heat exchanger. Cooling air passes over the heat exchanger to cool the compressed air, which is returned to the core engine to provide a cooling function. The heat exchanger is mounted through a flexible mount allowing movement between a static structure and the heat exchanger.

Abstract: A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (Dt) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is provided downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance from the exit point. A ratio of the distance to the diameter is greater than or equal to about 0.90.

Abstract: A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (Dt) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is provided downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance from the exit point. A ratio of the distance to the diameter is greater than or equal to about 0.90.

Abstract: A cryogenic cooling system for an aircraft includes a first air cycle machine, a second air cycle machine, and a means for collecting liquid air. The first air cycle machine is operable to output a cooling air stream based on a first air source. The second air cycle machine is operable to output a chilled air stream at a cryogenic temperature based on a second air source cooled by the cooling air stream of the first air cycle machine. An output of the second air cycle machine is provided to the means for collecting liquid air.

Abstract: A gas turbine engine includes a plurality of rotating components housed within a main compressor section and a turbine section. A first tap is connected to the main compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A cooling air valve is configured to selectively block flow of cooling air across the heat exchanger. A cooling compressor is connected downstream of the heat exchanger. A shut off valve stops flow between the heat exchanger and the cooling compressor. A second tap is configured to deliver air at a second pressure which is higher than the first pressure. A mixing chamber is connected downstream of the cooling compressor and the second tap. The mixing chamber is configured to deliver air to at least one of the plurality of rotating components. A system stops flow between the cooling compressor and the plurality of rotating components.