Pratt-whitney Canada Corp

Timothy Redford

David Harold Menheere

Daniel Van Den Ende

Santo Chiappetta

Richard Alan Kostka

An impeller shroud assembly for a gas turbine engine includes an annular impeller shroud disposed about an axial centerline. The impeller shroud includes a shroud inducer portion and a shroud exducer portion disposed radially outward of the shroud inducer portion and extending to an outer radial end of the impeller shroud. The shroud inducer portion and the shroud exducer portion defining an impeller-facing surface of the impeller shroud. The impeller shroud has a pivot point defined between the shroud inducer portion and the shroud exducer portion. The impeller shroud assembly further includes a clearance control device connected to the shroud exducer portion of the impeller shroud proximate the outer radial end. The clearance control device is configured to pivot the shroud exducer portion of the impeller shroud about the pivot point between a first axial position and a second axial position.

Impeller shroud assembly and method for operating same

The gas turbine engine for an aircraft includes at least a low pressure spool with a low pressure turbine shaft operatively connected to at least one turbine, the low pressure turbine shaft rotatable about an engine axis, and a low pressure compressor operatively connected to a low pressure compressor shaft that is independently rotatable relative to the low pressure turbine shaft. A differential gearbox has an input operatively connected to the low pressure turbine shaft, a first output and a second output, the first output of the differential gearbox operatively connected to the low pressure compressor shaft and the second output of the differential gearbox operatively connected to an output shaft of the gas turbine engine. The differential gearbox permits the output shaft, the low pressure compressor shaft and the low pressure turbine shaft to rotate at different speeds.

Gas turbine engine with variable speed output

The aircraft can have a first engine secured to a first wing on a first side of a fuselage, and a second engine secured to a second wing on a second side of the fuselage, the second wing having a proximal end secured to the fuselage, and a distal end extending away from the fuselage. While operating the first engine, compressed gas can be conveyed from the first engine to a thrust generating device located at the distal end of the second wing.

Aircraft and method of operating same

An auxiliary power unit (APU) includes, in serial flow communication: an engine compressor, a combustor and a turbine, the turbine rotatable about an engine axis. A first shaft operatively connects the turbine to the engine compressor and extends non-parallel to the engine axis. A second shaft operatively connects the turbine to a load and extends non-parallel to the engine axis. A method of operating an APU is also described.

Gas turbine auxiliary power unit

The gas turbine engine includes a combustion section including an annular swirl combustor having a combustor inlet, and a compressor section including a centrifugal compressor with an impeller, the impeller compressing and swirling an airflow and discharging the compressed and swirled airflow from the impeller outlet into the combustor inlet. The turbine section includes a radial turbine having a turbine fuel inlet and a turbine fuel outlet, the radial turbine receiving a flow of fuel at the turbine fuel inlet and discharging the flow of fuel from the turbine fuel outlet of the radial turbine into the combustor inlet.

Gas turbine engine with fuel-cooled turbine

Systems and methods for conditioning a fluid using bleed air from a bypass duct of a turbofan engine are disclosed. In one embodiment, such 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-driven fluid propeller configured to drive the bleed air through the heat exchanger.

System and method for increasing bleed air flow to a heat exchanger with a fluid-driven fluid propeller

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.

System and method for conditioning a fluid using bleed air from a bypass duct of a turbofan engine

A system for controlling gas turbine engine rotor blades tip clearance is described. A rotor is mounted to an engine shaft, supported by a thrust bearing, for rotation within a gas path shroud circumscribing blades of the rotor, the gas path shroud having a non-cylindrical shape in the vicinity of the rotor blades. A rotary actuator is associated with the thrust bearing and configured for axial translation of the thrust bearing, to thereby axially translate the engine shaft and the rotor blades relative to the gas path shroud. This translation is configured to vary the blade tip clearance of the rotor.

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Campbellville, Canada

Timothy Redford