United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

John H Mosley

William K Ackermann

Anthony P Cherolis

Jorge I Farah

James P Chasse

Bryan P Dube

Paul R Hanrahan

Steven D Porter

Ioannis Alvanos

Anthony R Bifulco

Gregory E Reinhardt

Joshua Daniel Winn

Philip S Stripinis

Christopher Treat

John J O'Connor

Michelle Diana Stearns

Robert Russell Mayer

Jonathan Lemoine

Paul K Sanchez

Kevin Zacchera

Andreas Sadil

Andrew D Burdick

Alberto A Mateo

Jonathan Ortiz

Seth A Max

Kalpendu J Parekh

Jeffrey Jay Lienau

David Frederick Glasspoole

Andrew D Milliken

Karl D Blume

Charles C Wu

James P Chrisikos

Roger Gates

Chris J Niggemeier

Hector M Pinero

Michael Luther Comeau

Paul E Coderre

Taryn Narrow

Robert Newman

Zachary Mott

Conan P Cook

Enzo Dibenedetto

Phililp S Stripinis

Douglas Paul Freiberg

Jon Pietrobon

Jonathan Oritz

David Glasspoole

A fluid supply system for a gas turbine engine includes a first component and a second component. A fluid fitting has a fluid junction. The fluid junction includes a first fluid port in fluid communication with and configured to supply a first cooling fluid to the first component. The fluid junction includes a second fluid port in fluid communication with and configured to supply a second cooling fluid to the second component. A primary pipe is fluidly connected to the fluid fitting. A flow meter is arranged upstream from the fluid junction and is configured to receive a cooling supply fluid from a fluid source. Thee flow meter is a wall having an orifice and a surface that is radiused and convex facing into the cooling supply fluid. A secondary pipe is secured to the fluid fitting and is configured to supply the second cooling fluid to the second component. A turbine section includes the first and second components. The primary pipe is configured to provide the cooling supply fluid from the fluid source to the fluid fitting.

Gas turbine engine cooling fluid metering system

A gas turbine engine includes a compressor section, a combustor, and a turbine section. The turbine section includes a high pressure turbine comprising a plurality of turbine blades. The gas turbine engine includes a tap for tapping air that is compressed by the compressor, to be passed through a heat exchanger to cool the air, the cooled air to be passed to the plurality of turbine blades. A sensor is located downstream of a leading edge of the combustor, and is configured to measure a characteristic of the cooled air. A controller is configured to compare the measured characteristic to a threshold and control an operating condition of the gas turbine engine based on the comparison.

Gas turbine engine control based on characteristic of cooled air

An inner diffuser case for use in a combustor of a gas turbine engine includes a multiple of struts across an annular flow path between an outer shroud and an inner shroud, at least one of the multiple of struts is an open strut.

Inner diffuser case struts for a combustor of a gas turbine engine

A cooling fluid system for a gas turbine engine includes a fluid source. A turbine section includes first and second components. A fluid supply system has a primary pipe that is configured to provide a cooling supply fluid from the fluid source to a fluid fitting having a fluid junction. The fluid junction is in fluid communication with and is configured to supply a first cooling fluid to the first component. The fluid junction is in fluid communication with and is configured to supply a second cooling fluid to the second component. A flow meter is upstream from the fluid junction and is configured to receive the cooling supply fluid.

A sealing arrangement includes a turbine static structure with a seal ring having groove with a contact surface. Also included is a bearing compartment with a second contact surface, and a piston seal positioned between the turbine static structure and the bearing compartment. The piston seal includes sides configured to contact the contact surfaces and is positioned in the groove.

Bearing compartment seal

A cooling system for a gas turbine engine turbine section includes a rotor supporting a blade having a cooling passage. A disc is secured relative to the rotor and it forms a cavity between the rotor and the disc. A bleed air source is in fluid communication with the cavity. An impeller is arranged in the cavity. The impeller is configured to increase a fluid pressure within the cavity to drive bleed air from the bleed air source and thereby provide a pressurized cooling fluid to the cooling passage.

Gas turbine engine turbine impeller pressurization

A mid-turbine frame for a gas turbine engine includes an outer frame case and an inner frame case. At least one spoke connects the outer frame case to the inner frame case. The spoke includes an inlet passage with at least one branch that extends in an axial direction.

Cooling passages for a mid-turbine frame

A mid-turbine frame for a gas turbine engine includes an inner frame case that includes a fluid passage. At least one spoke includes an inlet passage with at least one branch extending in an axial direction in fluid communication with the fluid passage in the inner frame case. A fitting fluidly connects the at least one branch to the fluid passage in the inner frame case. The fitting also includes a transfer tube connecting the cooling airflow passage to a cup boss. The transfer tube is fixed relative to the cylindrical portion and moveable relative to the cup boss. A swirler tube is in fluid communication with the fluid passage and is configured to direct cooling airflow in a radial inward and downstream direction.

A mid-turbine frame assembly includes an inner frame case which includes a first plurality of holes and a second plurality of holes. A plurality of tie rods is circumferentially spaced around the inner frame case and includes an inlet passage that is aligned with the first plurality of holes. A plurality of hollow airfoils is aligned with the second plurality of holes.

A gas turbine engine configured to rotate in a circumferential direction about an axis extending through a center of the gas turbine engine comprises a turbine stage. The turbine stage comprises a disk, a plurality of blades and a mini-disk. The disk comprises an outer diameter edge having slots, an inner diameter bore surrounding the axis, a forward face, and an aft face. The plurality of blades is coupled to the slots. The mini-disk is coupled to the aft face of the rotor to define a cooling plenum therebetween in order to direct cooling air to the slots. In one embodiment of the invention, the cooling plenum is connected to a radially inner compressor bleed air inlet through all rotating components so that cooling air passes against the inner diameter bore.

Turbine coolant supply system

The present invention relates to a system for delivering cooling air to a seal arrangement in a turbine stage of a gas turbine engine. The system comprises at least one vane having a passageway extending from an outer platform of the at least one vane to an inner platform of the at least one vane. A tube insert is positioned within the passageway. The tube insert has an inlet at one end for receiving cooling air from a source of cooling air and an outlet at a second end. A cover assembly is attached to the second end of the tube for receiving cooling air from the tube and delivering the cooling air to the seal arrangement. In a preferred embodiment, the cooling air is pre-swirled in the direction of rotation of a rotor stage of the turbine stage.

Integral swirl knife edge injection assembly

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Portland, CT, United States of America

John H Mosley