United Technologies Corporation

Rtx Corporation

Raytheon Technologies Corporation

Oliver V Atassi

Alan Retersdorf

Stephen H Taylor

Jonathan Gilson

Oleg Petrenko

Wesley K Lord

Nathan A Snape

Steven H Zysman

Andre M Hall

Robert M Miller

Ramons Andris Reba

William D Owen

Dennis Cicon

Fred W Schwark, Jr

Malcolm P MacDonald

Constantine Baltas

Anil Prasad

Michaela M Logue

Jr Fred W Schwark

A gas turbine engine assembly including a tap that is at a location up stream of the combustor section for drawing a bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow and communicate the heated bleed airflow into the turbine section where it is expanded to drive the turbine section.

Air recuperated engine with air reinjection

A gas turbine engine assembly includes a core engine that includes a core flow path where a core airflow is compressed in a compressor section, communicated to a combustor section, mixed with fuel and ignited to generate an exhaust gas flow that is expanded through a turbine section. The turbine section is coupled to drive the compressor section through an engine drive shaft. A tap is at a location up stream of the combustor section for drawing a bleed airflow. A bleed air heat exchanger places the bleed airflow in thermal communication with an auxiliary flow for heating the bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow.

Bottoming cycle for waste heat recovery and engine cooling

A turbine engine assembly includes a tap is at a location upstream of the combustor section where a bleed airflow is drawn. The bleed air is pressurized in an auxiliary compressor section, heated in an exhaust heat exchanger, and expanded through a power turbine that is coupled to drive the auxiliary compressor section.

Air bottoming cycle air cycle system source

An aircraft propulsion system includes a core engine that includes a core flow path where a core airflow is compressed in a compressor section, communicated to a combustor section, mixed with fuel and ignited to generate a gas flow that is expanded through a turbine section for powering a primary propulsor. The aircraft propulsion system further includes a tap that is at a location upstream of the combustor section where a bleed airflow is drawn, a heat exchanger where the bleed airflow is heated by the gas flow, a power turbine through which heated bleed airflow is expanded to generate a work output, and a secondary propulsor that is driven by the work output that is generated by the power turbine.

Air bottoming cycle driven propulsor

A turbine engine assembly includes a tap is at a location upstream of the combustor section where a bleed airflow is drawn. The bleed air is pressurized in an auxiliary compressor section, heated in an exhaust heat exchanger and expanded through a power turbine that is coupled to drive the auxiliary compressor section.

A gas turbine engine includes a core engine that has at least a compressor section, a combustor section and a turbine section disposed along a central axis. A fan is coupled to be driven by the turbine section. A fan nozzle is aft of the fan and defines an exit area. The fan nozzle has a body with an airfoil cross-section geometry. The body includes a wall that has a controlled mechanical property distribution that varies in material macro- or micro-structure by location on the wall in accordance with a desired flutter characteristic at the location.

Variable area fan nozzle with wall thickness distribution

A nacelle assembly according to an exemplary aspect of the present disclosure includes, among other things, a core nacelle defined at least partially about a core engine, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path, and a variable area fan nozzle in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section downstream of the first fan nacelle section. The first fan nacelle section and the second fan nacelle section are axially movable relative to one another to define an auxiliary port to vary a fan nozzle exit area and adjust fan bypass airflow. The auxiliary port is defined between the first fan nacelle section and the second fan nacelle section. The first fan nacelle section comprises a first acoustic system which provides an acoustic impedance configured to attenuate a noise characterized by a leading edge of the second fan nacelle section. The first acoustic system is defined at least in part within a trailing edge region of the first fan nacelle section. A method of reducing a total effective perceived noise level of a gas turbine engine with a variable area fan nozzle is also disclosed.

Variable area fan nozzle for gas turbine engine

A gas turbine engine is provided having a fan case and a translating sleeve positioned downstream from the fan case. A flow channel extends between the fan case and the translating sleeve. The flow channel includes an inner diameter and an outer diameter. A structural guide vane is positioned within the flow channel and extends from the inner diameter to the outer diameter. A liner is positioned between an aft end of the fan case and an aft end of the translating sleeve to reduce vibratory stress on the structural guide vane.

Liner for a gas turbine engine

A gas turbine engine includes a core engine that has at least a compressor section, a combustor section and a turbine section disposed along a central axis. A fan is coupled to be driven by the turbine section. A fan nozzle is aft of the fan and defines an exit area. The fan nozzle has a body that defines an airfoil cross-section geometry. The body includes a wall that has a controlled mechanical property distribution that varies by location on the wall in accordance with a desired flutter characteristic at the location.

A gas turbine engine includes a first airflow structure and a second airflow structure disposed aft of the first airflow structure. The second airflow structure includes a leading edge region. A thickness of the leading edge region is based on a thickness of a wake in the airflow produced by the first airflow structure when the airflow passes between the first airflow structure and the second airflow structure.

Noise reducing vane

A nacelle assembly for a high-bypass gas turbine engine includes a core nacelle defined about an engine centerline axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass flow path. A variable area fan nozzle is in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section. The second fan nacelle section is axially movable relative to the first fan nacelle section to define an auxiliary port at a non-closed position to vary a fan nozzle exit area and adjust fan bypass airflow. The second fan nacelle section includes an acoustic system that has an acoustic impedance located on a radially outer surface.

Gas turbine engine with noise attenuating variable area fan nozzle

A nacelle for a gas turbine engine system includes a housing that has a pressure side and a suction side. A passage extends between the pressure side and the suction side for permitting airflow from the pressure side to the suction side. The passage includes a first curved section that has an inlet at the pressure side. A linear section is connected with the first curved section. A second curved section is connected with the linear section and has an outlet at the suction side. The outlet has a cross sectional area smaller than a cross sectional area of the linear section.

Passive boundary layer bleed system for nacelle inlet airflow control

A method of controlling flutter of a fan nozzle includes, for a given design of a fan nozzle, determining a vibration mode that causes a flutter characteristic of the fan nozzle and, in response to the determined vibration mode, establishing a wall thickness distribution of at least one wall of the fan nozzle to include local thick portions and local thin portions that alter the flutter characteristic.

A gas turbine engine includes a core engine that has at least a compressor section, a combustor section and a turbine section disposed along a central axis. A fan is coupled to be driven by the turbine section. A fan nacelle is arranged around the fan, and a bypass passage extends between the fan nacelle and the core engine. A variable area fan nozzle (VAFN) extends at least partially around the central axis and defines an exit area of the bypass passage. The VAFN is selectively movable to vary the exit area. The VAFN includes a body that defines an airfoil cross-section shape. The VAFN includes a wall that has a mechanical property distribution in accordance with a computer-simulated vibration profile of a flutter characteristic of the VAFN.

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Longmeadow, MA, United States of America

Oliver V Atassi