Abstract: A gas turbine engine includes an engine static structure housing that includes a compressor section and a turbine section. A combustor section is arranged axially between the compressor section and the turbine section. A core nacelle encloses the engine static structure to provide a core compartment. An oil tank is arranged in the core compartment and is axially aligned with the compressor section. A heat exchanger is secured to the oil tank and arranged in the core compartment.

Abstract: The present disclosure provides systems and methods related to thermal management systems for gas turbine engines. For example, a thermal management system comprises a thermally neutral heat transfer fluid circuit, a first heat exchanger disposed on the fluid circuit, and a second heat exchanger disposed on the fluid circuit.

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 includes a fan case radially outwardly of a core compartment. A compressor section is located within an engine core compartment and includes a front mount flange and an aft mount flange. An oil tank is mounted to at least one of the fan case or the front and aft mount flanges. The oil tank has a cooling structure integrated into an outer surface such that the oil tank is subjected to cooling air flow from a plurality of air sources.

Abstract: A gas turbine engine includes a plurality of rotating components housed within a compressor section and a turbine section. A first tap is connected to the compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap and configured to deliver air to an aircraft fuselage. A cooling compressor is connected downstream of the heat exchanger. A high pressure feed is configured to deliver air at a second pressure which is higher than the first pressure. The cooling compressor is configured to deliver air to at least one of the plurality of rotating components. A valve assembly that can select whether air from the first tap or air from the high pressure feed is delivered to the aircraft pneumatic system.

Abstract: An example turbomachine assembly includes, among other things, a nose cone of a turbomachine, and a pump at least partially within an interior of the nose cone. The pump is selectively rotated by a motor to communicate air to the interior.

Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. A core housing has an outer peripheral surface and a fan housing defining an inner peripheral surface. At least one bifurcation duct extends between the outer peripheral surface to the inner peripheral surface. The heat exchanger is received within the at least one bifurcation duct.

Abstract: A gas turbine engine comprises a compressor section having a downstream most end and a cooling air tap at a location spaced upstream from the downstream most end. The cooling air tap is passed through at least one boost compressor and at least one heat exchanger, and then passed to a turbine section to cool the turbine section, the boost compressor being controlled to provide a desired pressure to the turbine section.

Abstract: A gas turbine engine has a fan rotor delivering air into a bypass duct defined between an outer fan case and an outer interior housing. The fan rotor also delivers air into a compressor section, a combustor, a turbine section. A chamber is defined between the outer interior housing and an inner housing. The inner housing contains the compressor section, the combustor and the turbine section. A first conduit taps hot compressed air to be cooled and passes the air to at least one heat exchanger. The air is cooled in the heat exchanger and returned to a return conduit. The return conduit passes the cooled air to at least one of the turbine section and the compressor section. The heat exchanger has a core exhaust plane. The turbine section has at least a first and a downstream second rotor blade row, with the core exhaust plane located downstream of a center plane of the second blade row.

Abstract: The present disclosure provides systems and methods related to thermal management systems for gas turbine engines. For example, a thermal management system comprises a thermally neutral heat transfer fluid circuit, a first heat exchanger disposed on the fluid circuit, and a second heat exchanger disposed on the fluid circuit.

Abstract: An intercooled cooling system for a gas turbine engine is provided. The intercooled cooling system includes cooling stages in fluid communication with an air stream utilized for cooling. A first cooling stage is fluidly coupled to a bleed port of the gas turbine engine to receive and cool bleed air with the air stream to produce a cool bleed air. The intercooled cooling system includes a pump fluidly coupled to the first cooling stage to receive and increase a pressure of the cool bleed air to produce a pressurized cool bleed air. A second cooling stage is fluidly coupled to the pump to receive and cool the pressurized cool bleed air to produce an intercooled cooling air. The intercooled cooling system includes an air cycle machine in fluid communication to outputs of the cooling stages to selectively receive the cool bleed air or the intercooled cooling air.

Abstract: A first cooling stage is fluidly coupled to a bleed port of a compressor to receive and cool bleed air with the air stream to produce a cool bleed air. A cooling pump receives and increases a pressure of the cool bleed air to produce a pressurized cool bleed air. A second cooling stage is fluidly coupled to the pump to receive and cool the pressurized cool bleed air to produce an intercooled cooling air. A valve is downstream of the first cooling stage, the valve selectively delivering air into a mixing chamber where it is mixed with air from a tap that is compressed to a higher pressure than the air from the bleed port, and the valve also selectively supplying air from the first cooling stage to a use on an aircraft associated with the gas turbine engine. A method is also disclosed.

Abstract: An intercooled cooling system for a gas turbine engine is provided. The intercooled cooling system includes a plurality of cooling stages in fluid communication with an air stream utilized for cooling. A first cooling stage of the plurality of cooling stages is fluidly coupled to a bleed port of a compressor of the gas turbine engine to receive and cool bleed air with the air stream to produce a cool bleed air. The intercooled cooling system also includes a pump fluidly coupled to the first cooling stage to receive the cool bleed air and increase a pressure of the cool bleed air to produce a pressurized cool bleed air. A second cooling stage of the plurality of cooling stages is fluidly coupled to the pump to receive and cool the pressurized cool bleed air to produce an intercooled cooling air, which is provided to the gas turbine engine.

Abstract: A temperature detection device located downstream a cooling network coupled to a gas turbine engine and located upstream of at least one of a compressor rotor of a compressor section of the gas turbine engine and a turbine rotor of a turbine section of the gas turbine engine, is provided. The temperature detection device includes at least one thermocouple configured to detect an operational temperature of the cooling network. The temperature detection device communicates the operational temperature to a control system.

Abstract: A gas turbine engine includes a fan case radially outwardly of a core compartment. A compressor section is located within an engine core compartment and includes a front mount flange and an aft mount flange. An oil tank is mounted to at least one of the fan case or the front and aft mount flanges. The oil tank has a cooling structure integrated into an outer surface such that the oil tank is subjected to cooling air flow from a plurality of air sources.

Abstract: A gas turbine engine has a compressor section with a downstream most end and a cooling air tap at a location spaced upstream from the downstream most end. The cooling air tap is passed through at least one boost compressor and at least one heat exchanger, and then passed to a turbine section to cool the turbine section. The boost compressor is driven by a driveshaft which is driven by the turbine section. A boost turbine selectively drives the boost compressor.

Abstract: A gas turbine engine has a compressor section and a turbine section. A secondary cooling air includes a first fluid connection to tap cooling air and pass the cooling air through a plurality of tubes, and a second fluid connection for returning air from the tubes back to at least one of the compressor and turbine for cooling. A sensor senses a condition of the cooling air downstream of the tubes and a control compares the sensed condition of the cooling air to an expected condition, and to identify a potential concern in the cooling air system should the sensed condition differ from the expected condition by more than a predetermined amount.

Abstract: A gas turbine engine has a compressor section, a combustor, and a turbine section. An associated fluid is to be cooled and an associated fluid is to be heated. A transcritical vapor cycle heats the fluid to be heated, and cools the fluid to be cooled. The transcritical vapor cycle includes a gas cooler in which the fluid to be heated is heated by a refrigerant in the transcritical vapor cycle. An evaporator heat exchanger at which the fluid to be cooled is cooled by the refrigerant in the transcritical vapor cycle. A compressor upstream of the gas cooler compresses the refrigerant to a pressure above a critical point for the refrigerant. An expansion device expands the refrigerant downstream of the gas cooler, with the evaporator heat exchanger being downstream of the expansion device, and such that the refrigerant passing through the gas cooler to heat the fluid to be heated is generally above the critical point.

Abstract: A nose cone assembly for a gas turbine engine and method of circulating air in a gas turbine engine are disclosed. The nose cone assembly includes a nose cone having an aperture communicating air to an interior space of the nose cone and a discharge member communicating the air out of the nose cone.

Abstract: A gas turbine engine assembly comprising, a gearbox including a first housing that includes a first auxiliary gear drive on a first portion thereof, a second housing that includes a second auxiliary gear drive on a second portion thereof, and a third housing that includes a third auxiliary gear drive on a third portion thereof, the housings being interconnected so that the first portion of the first housing, the second portion of the second housing and the third portion of the third housing form a substantially triangular polyhedron shape, with the second portion of the second housing disposed between the first portion of the first housing and the third portion of the third housing. The first auxiliary gear drive, the second auxiliary gear drive and the third auxiliary gear drive project outwardly in mutually divergent directions.