Honeywell International Inc.

Malak Fouad Malak

Balamurugan Srinivasan

Luis A Tapia

Jong Liu

Mark C Morris

Daniel Cregg Crites

David R Waldman

David Chou

Rajiv Rana

Vinayender Kuchana

Craig Mckeever

Dhinagaran Ramachandran

Gopal Samy Muthiah

Ardeshir Riahi

Kin Poon

Jyothishkumar Venkataramanan

Thomas E Strangman

Derek Raybould

Karthikeyan Paramanandam

Vighneswara Rao Kollati

Chandiran Jayamurugan

Paul Couey

Sridharan Reghupathy Narayanan

Sujatha Guntu

Don Mittendorf

Steve Halfmann

Paul Chipko

Milton Ortiz

Brent Ludwig

Stuart Andrew Harman

Jon Kettinger

Anish Kumar Gupta

Ramakumar Venkata Naga Bommisetty

Dhanvantri Joshi

Sridhar Murari

Deane Krmpotich

Luis Tapia

Anish Gupta

Sajatha Guntu

Craig Mcckeever

A turbine section for a gas turbine engine is annular about a longitudinal axis. The turbine section includes a first turbine with a first outlet, and a second turbine with a second inlet. The turbine section includes an inter-turbine duct extending from the first outlet to the second inlet and configured to direct a flow along a flow direction. The inter-turbine duct is defined by a hub and a shroud. The turbine section includes at least a first splitter blade positioned between the hub and the shroud. The first splitter blade includes a pressure side, a suction side, and at least one vortex generating structure having a leading end opposite a trailing end positioned on the suction side such that a first angle is defined between the vortex generating structure and the flow direction. The vortex generating structure extends in a radial direction from the suction side toward the hub.

Inter-turbine ducts with flow control mechanisms

An airfoil for a gas turbine engine includes a body with a first side wall and a second side wall joined at a leading edge and a trailing edge, the first side wall having a first interior surface and the second side wall having a second interior surface. The airfoil further includes an internal wall disposed within of the body and extending between the first interior surface and the second interior surface to define a supply passage and a leading edge passage. The internal wall defines a plurality of cooling holes to direct cooling air from the supply passage to the leading edge passage such that the cooling air impinges upon the leading edge. The airfoil further includes a first plurality of grooves formed in the first interior surface, each the first plurality of grooves extending in a chordwise direction within the leading edge passage.

Gas turbine engines with improved leading edge airfoil cooling

Cladding material is applied by laser to a net-shape. A method of cladding a host component includes installing the component in a fixture. A shroud component is located against the host component adjacent a select location for the cladding. Cladding is applied to the host component to the select location and adjacent to shroud component so that the shroud component defines an edge of the cladding as applied. The edge of the cladding as defined by the shroud component defines a desired cladding profile requiring no/approximately no post-cladding processing to remove over-cladded material.

Laser tip cladding to net-shape with shrouds

A turbine section is provided for a gas turbine engine. The turbine section is annular about a longitudinal axis. The turbine section includes a first turbine with a first inlet and a first outlet; a second turbine with a second inlet and a second outlet; an inter-turbine duct extending from the first outlet to the second inlet and configured to direct an air flow from the first turbine to the second turbine, the inter-turbine duct being defined by a hub and a shroud; and at least a first splitter blade disposed within the inter-turbine duct. The first splitter blade includes a pressure side facing the shroud, a suction side facing the hub, and at least one vortex generating structure positioned on the suction side.

An engine component includes a body having an internal surface and an external surface, the internal surface at least partially defining an internal cooling circuit. The engine component further includes a plurality of cooling holes formed in the body and extending between the internal cooling circuit and the external surface of the body. The plurality of cooling holes includes a first cooling hole with forward diffusion and lateral diffusion.

Gas turbine engine components with lateral and forward sweep film cooling holes

The present disclosure provides various embodiments of cooling circuits, turbine blades with cooling circuits, and methods of forming such turbine blades, having raised rib turbulator structures, which may be used in gas turbine engines. In one exemplary embodiment, a cooling circuit for directing a flow of fluid is disclosed, the cooling circuit includes a cooling circuit wall and a plurality of raised turbulator ribs, each turbulator rib of the plurality of raised turbulator ribs being spaced apart from the cooling circuit wall to allow the fluid to flow between the cooling circuit wall and the plurality of turbulator ribs.

Turbine blades and methods of forming turbine blades having lifted rib turbulator structures

A turbine rotor blade is provided for a turbine section of an engine. The turbine rotor blade includes a platform and an airfoil extending from the platform into a mainstream gas path of the turbine section. The airfoil includes a first side wall; a second side wall joined to the first side wall at a leading edge and a trailing edge; a tip cap extending between the first side wall and the second side wall; a first parapet wall extending from the first side wall; and a first cooling hole through the tip cap and the first parapet wall configured to deliver cooling air. The first cooling hole has a closed channel section and an open channel section. The open channel section forms a slot.

Turbine rotor blades with improved tip portion cooling holes

In accordance with an exemplary embodiment, a turbine rotor blade is provided for a turbine section of an engine. The turbine rotor blade includes a platform and an airfoil extending from the platform into a mainstream gas path of the turbine section. The airfoil includes a first side wall; a second side wall joined to the first side wall at a leading edge and a trailing edge; a tip cap extending between the first side wall and the second side wall; a first parapet wall extending from the first side wall; a first parapet wall cavity formed at least partially within the first parapet wall; and a first cooling hole extending between the first parapet wall cavity and a first surface of the first parapet wall such that cooling air flows through the first parapet wall cavity, through the first cooling hole, and out of the first parapet wall.

Turbine rotor blades with tip portion parapet wall cavities

An engine component includes a body; and a plurality of cooling holes formed in the body, at least one of the cooling holes having a multi-lobed shape with at least a first lobe, a second lobe, and a third lobe.

Gas turbine engine components with film cooling holes having cylindrical to multi-lobe configurations

Turbine blade airfoils, film cooling systems thereof, and methods for forming improved film cooled components are provided. The turbine blade airfoil has an external wall surface and comprises leading and trailing edges, pressure and suction sidewalls both extending between the leading and the trailing edges, an internal cavity, one or more isolation trenches in the external wall surface, a plurality of film cooling holes arranged in cooling rows, and a plurality of span-wise surface connectors interconnecting the outlets of the film cooling holes in the same cooling row to form a plurality of rows of interconnected film cooling holes. Each film cooling hole has an inlet connected to the internal cavity and an outlet opening onto the external wall surface. The span-wise surface connectors in at least one selected row of interconnected film cooling holes are disposed in the one or more isolation trenches.

Turbine blade airfoils including film cooling systems, and methods for forming an improved film cooled airfoil of a turbine blade

A turbine section of a gas turbine engine is annular about a longitudinal axis. The turbine section includes a first turbine with a first inlet and a first outlet; a second turbine with a second inlet and a second outlet; and an inter-turbine duct extending from the first outlet to the second inlet and configured to direct an air flow from the first turbine to the second turbine. The inter-turbine duct has a first station with a first meridional area, a second station with a second meridional area, and a third station with a third meridional area. The first station is upstream of the second station and the second station is upstream of the third station, and the second meridional area is less than or equal to the first meridional area.

Inter-turbine ducts with variable area ratios

In accordance with an exemplary embodiment, disclosed is an air-cooled turbine blade having an airfoil shape, including a convex suction side wall, a concave pressure side wall, the walls including an interior surface that defines an interior with the blade, a suction side flow circuit formed within the blade interior, a pressure side flow circuit formed within the blade interior; and a trailing edge pin bank positioned aft of the suction side and pressure side flow circuits. The turbine blade includes a wishbone-shaped architecture at a transition point between the suction side flow circuit and the pressure side flow circuit and the trailing edge pin bank.

Cooled turbine airfoil structures

Turbine blade airfoils, showerhead film cooling systems thereof, and methods for cooling the turbine blade airfoils using the same are provided. The airfoil has a leading edge and a trailing edge, a pressure sidewall and a suction sidewall both extending between the leading and the trailing edges, and an internal cavity for supplying cooling air. A showerhead of film cooling holes is connected to the internal cavity. Each film cooling hole has an inlet connected to the internal cavity and an outlet opening onto an external wall surface at the leading edge of the airfoil. A plurality of surface connectors is formed in the external wall surface. Each surface connector of the plurality of surface connectors interconnects the outlets of at least one selected pair of the film cooling holes.

Turbine blade airfoils including showerhead film cooling systems, and methods for forming an improved showerhead film cooled airfoil of a turbine blade

A turbine section of a gas turbine engine is provided. The turbine section is annular about a longitudinal axis and includes first turbine with a first inlet and a first outlet; a second turbine with a second inlet and a second outlet; an inter-turbine duct extending from the first outlet to the second inlet and configured to direct an air flow from the first turbine to the second turbine; and a first guide vane disposed within the inter-turbine duct.

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Tempe, AZ, United States of America

Malak Fouad Malak