Rolls-Royce North American Technologies Inc.

Rolls-Royce Corporation

Paul O'Meallie

Kerry J Lighty

Alan W Smith

Douglas J Snyder

Sean Andrew Morgan

Matthew Jordan

Daniel G Edwards

Brian R Bennett

An example thermal management system may include a first heat exchanger including a bleed air inlet configured to receive input bleed air from a gas turbine engine and a bleed air outlet configured to output cooled bleed air. A turbine including a turbine inlet may be fluidically coupled to the bleed air outlet. The turbine may be configured to drive a shaft mechanically coupled to the turbine in response to expansion of the cooled bleed air through the turbine. A second heat exchanger may include an expanded bleed air inlet fluidically coupled to a turbine outlet of the turbine. The second heat exchanger may be configured to extract heat from at least one heat source using the expanded bleed air.

Bleed expander cooling with turbine

A cooling system includes a first cooling loop, a second cooling loop and a heat exchanger configured to transfer heat from the first cooling loop to the second cooling loop. The first cooling loop includes a flow restrictor, an inertial separator, and a pressure relief valve cooperating to effect diversion of vapor present in the first cooling loop due to a leak between the first cooling loop and the second cooling loop.

Vapor leak pressure relief and diversion system

A rotor assembly adapted for a gas turbine engine includes a shaft, a wheel, and a retaining nut. The shaft extends along an axis and includes a first tapered surface. The wheel is arranged circumferentially around the shaft and includes a second tapered surface. The retaining nut is fastened to the shaft and applies an axial force to the wheel to couple the wheel with the shaft.

Fan rotor with tapered drive joint

A shaft assembly for a gas turbine engine includes a shaft, a bearing, and an annular sump cover. The shaft is arranged along an axis and configured to rotate about the axis. The bearing is coupled to the shaft to facilitate rotation of the shaft about the axis. The annular sump cover is fixed relative to the axis and is configured to allow a lubrication fluid and a buffer air into an air chamber formed between the shaft, the bearing, and the annular sump cover.

Gas turbine engine cooled sump buffer air system

A cooling system includes a first cooling loop, a second cooling loop and a heat exchanger configured to transfer heat from the first cooling loop to the second cooling loop. The first cooling loop includes a vapor/liquid separation feature configured to separate vapor present in the first cooling loop due to a leak between the first cooling loop and the second cooling loop. The first cooling loop also includes a pressure sensor configured to detect an increase in pressure in the first cooling loop that may result from a leak of second coolant into the first cooling loop.

Vapor leak separation and detection system

A gas turbine engine includes a bypass duct, a compressor, and a pyrotechnic gas conduit. The bypass duct conducts bypass air between an outer bypass wall and an inner bypass wall. The compressor includes an impeller to compress engine air. The pyrotechnic gas conduit is configured to conduct a gas onto the impeller to start the gas turbine engine.

Gas turbine engine with pyrotechnic air starter

A method for cooling a hybrid electric aircraft propulsion system comprises transporting a coolant through a two-phase pumped loop (TPPL) in thermal contact with an electrical machine and a plurality of power modules to be cooled, where the TPPL includes: a parallel arrangement of cold plates; an evaporator; a condenser; a first control valve; a liquid receiver; and a pump. A sensor positioned upstream of the cold plates, and in some cases upstream of the liquid receiver, measures pressure and/or temperature of a return stream of the coolant and transmits measurement data to a first controller electrically connected to the first control valve. The first controller regulates flow of a first liquid stream through the first control valve based on the pressure and/or temperature measured by the sensor, thereby keeping the return stream at a temperature within a predetermined temperature range.

Thermal management system and method for cooling a hybrid electric aircraft propulsion system

An example method includes obtaining a representation of a change in rotational speed of an electric machine; obtaining a representation of an expected change in rotational speed of the electric machine; and determining, based on the obtained representation of the change in rotational speed of the electric machine and the representation of an expected change in rotational speed of the electric machine, whether a disconnect device has failed, wherein, when operating in an engaged state, the disconnect device is configured to couple rotational mechanical energy between the electric machine and a rotating device, and wherein, when operating in a disengaged state, the disconnect device is not configured to couple rotational mechanical energy between the electric machine and the rotating device.

Automatic testing of electric machine disconnect

Systems and methods of heat management of turbine engines including turbofans, turboprops and turboshafts and fan driven propulsion systems. The propulsion system may comprise a fan, nacelle, an electrical or mechanical heat source and a cooling system consisting of heat exchangers in the fan duct and on the nacelle and coolant pumps. The heat source can be a motor or a generator or turbine machinery or accessories rotationally coupled to rotating shafts. The heat management system transfers heat to the air in the fan flow path to provide additional fan thrust. The heat management system also transfers heat to structural members in the gas flow path that require anti-icing.

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Brownsburg, IN, United States of America

Paul O'Meallie