Pratt-Whitney Canada Corp

Kevin Ng

Eric Durocher

Daniel Alecu

Keith Morgan

Philippe Beauchesne-Martel

Johnny Vinski

Jean Dubreuil

Ghislain Plante

Adam Logan

Marc Jordan Adique

Hervé Turcotte

Martin Drolet

Marc Jordon Adique

A method of operating an engine of an aircraft, the engine having a bleed air system. The method includes in flight, directing pressurized air from a source of pressurized air external to the engine to the bleed air system of the engine. An aircraft, comprising a first engine having a bleed air system, a second engine having a bleed air system, and a source of pressurized air that is external to the first and/or the second engine, the source of pressurized air being selectively fluidly connectable to the bleed air system of the first and/or the second engine, is also disclosed.

Air system of multi-engine aircraft

A multi-engine aircraft includes a first engine drivingly engaged to a common rotatable load and a second engine drivingly engaged to the common rotatable load, the second engine having a bleed air system and a control system in communication with a compressed air switching system. The control system controls operation of the second engine and/or the compressed air switching system. The compressed air switching system includes a switching valve that is displaceable between at least a first position and a second position, the first position interconnecting a lower pressure inlet and a switch outlet, and the second position interconnecting a high pressure inlet and the switch outlet. The switch outlet is in communication with the bleed air system of the second engine. The control system actuates the switching valve to switch between the first and second positions.

Air system switching system to allow aero-engines to operate in standby mode

A gas turbine engine includes a compressor, an annular casing surrounding the compressor, and a bleed flow adapter mounted to an exterior side of the annular casing. The annular casing includes the exterior side and an interior side. The interior side surrounds a compressor bleed cavity located downstream of at least a portion of the compressor. The bleed flow adapter is in fluid communication with the bleed cavity. The bleed flow adapter includes an inlet end, an outlet end, and an inner diameter surface extending between the inlet end and the outlet end. The inner diameter surface defines a bleed passage. The bleed flow adapter further includes a fluid port formed through the inner diameter surface. The gas turbine engine further includes a bleed flow measurement system including a first pressure sensor in fluid communication with the bleed passage of the bleed flow adapter via the fluid port.

Method and apparatus for measuring compressor bleed flow

A method of operating a multi-engine aircraft having two or more gas turbine engines includes operating a first engine in a powered mode to provide motive power to the aircraft, and, in flight, operating a second engine in either a powered mode to provide motive power to the aircraft or in a standby mode to provide substantially no motive power to the aircraft. When operating the second engine in the powered mode, pressurized air is bled from a first bleed location of a compressor of the second engine. When operating the second engine in the standby mode, pressurized air is bled from a second bleed location of the compressor of the second engine and supplying the pressurized air to a bleed air system of the second engine. The second bleed location is downstream of the first bleed location within the compressor of the second engine.

A heat exchanger defines an annulus divided by a plurality of radial plates extending longitudinal along the annulus, into a plurality of channels. A first group of the channels form an oil passage and a second group of channels form a fuel passage. The channels in the respective first and second groups are circumferentially alternately positioned one to another, and heat transfer from the oil passage to the fuel passage takes place through the radial plates.

Heat exchanger for gas turbine engines

The liquid cooling system has a heat exchanger having a fluid inlet and an outlet; a fluid supply conduit leading to the inlet of the heat exchanger; a fluid return conduit extending from the outlet of the heat exchanger; a bypass conduit extending between the fluid supply conduit and the fluid return conduit; a thermal valve configured for selectively closing the bypass conduit, the valve having a temperature sensing element positioned downstream of both the heat exchanger and the bypass conduit, the temperature sensing element configured to selectively move the thermal valve in response to a temperature change of the liquid which the temperature sensing element is exposed to relative to a temperature threshold of the valve; and a deflector positioned between the temperature sensing element and at least one of the bypass conduit and the heat exchanger outlet.

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Hamilton, Canada

Kevin Ng