Pratt-Whitney Canada Corp

United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Michael Leslie Clyde Papple

Sri Sreekanth

Dominic J Mongillo, Jr

Russell J Bergman

Shawn J Gregg

Mohammed Ennacer

Toufik Djeridane

Daniel Lecuyer

Si-Man Amy Lao

Alan Juneau

Dominique Michel Nadeau

Timothy J Jennings

Ricardo Trindade

Ghislain Plante

Panagiota Tsifourdaris

Domenico Di Florio

David A Niezelski

Rene Paquet

John R Battye

Francois Caron

Patrick Peloquin

Nicholas Grivas

XingYun Haggard

A component for a gas turbine engine includes, among other things, an airfoil that extends between a leading edge and a trailing edge and a cooling circuit disposed inside of the airfoil. The cooling circuit includes at least one core cavity that extends inside of the airfoil, a baffle received within the at least one core cavity, a plurality of pedestals positioned adjacent to the at least one core cavity and a first plurality of axial ribs positioned between the plurality of pedestals and the trailing edge of the airfoil.

Gas turbine engine airfoil cooling circuit arrangement

A turbomachine airfoil element comprises an airfoil having: an inboard end; an outboard end; a leading edge; a trailing edge; a pressure side; and a suction side. A span between the inboard end and the outboard end is 1.75-2.20 inches. A chord length at 50% span is 1.05-1.35 inches. At least two of: a first mode resonance frequency is 2400&#xb1;10% Hz; a second mode resonance frequency is 4950&#xb1;10% Hz; a third mode resonance frequency is 7800&#xb1;10% Hz; a fourth mode resonance frequency is 8700&#xb1;10% Hz; and a fifth mode resonance frequency is 12500&#xb1;10% Hz.

Turbomachine blade

A turbine component has a plurality of cooling passages each extending through a body of a structure between opposite hot and cold surfaces of the structure. According to one embodiment, at least one of the cooling passages includes a plurality of upstream paths defining respective inlet openings on the cold surface and merging into a number of downstream paths defining respective outlet openings on the hot surface.

Cooling passages in a turbine component

A turbine component includes a structure having hot and cold surfaces opposite each other; and a plurality of cooling holes extending between the cold and hot surfaces, each of the cooling holes including at least one projection or recess element to form part of a fluid path surface with enhanced in-hole heat convection.

Effusion cooling holes

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having a first path side and a second path side and a platform cooling circuit disposed on one of the first path side and the second path side of the platform. The platform cooling circuit includes a first core cavity, a cavity in fluid communication with the first core cavity, and a cover plate positioned to cover at least the cavity.

Platform cooling circuit for a gas turbine engine component

A component for a gas turbine engine according to an exemplary aspect of the present disclosure including, among other things, an airfoil that extends between a leading edge and a trailing edge and a cooling circuit disposed inside of the airfoil. The cooling circuit includes at least one core cavity that extends inside of the airfoil, a baffle received within the at least one core cavity, a plurality of pedestals positioned adjacent to the at least one core cavity and a first plurality of axial ribs positioned between the plurality of pedestals and the trailing edge of the airfoil.

A turbine blade for a gas turbine engine with an airfoil portion defined by a perimeter wall surrounding at least one enclosure, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the at least one enclosure and a gaspath of the gas turbine engine. The plurality of cooling holes includes at least one row of holes selected from the group consisting of a first row, a second row, a third row and a fourth row, with the first, second, third and fourth rows of holes respectively including the holes numbered PA-1 to PA-10, PB-1 to PB-3, HA-1 to HA-9, and SA-1 to SA-8 located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3.

High pressure turbine blade cooling hole distribution

A rotor assembly for a gas turbine engine, the rotor assembly comprises a plurality of cooling air deflectors made integral with the rotor disk to redirect air to a manifold at a bottom side of a corresponding blade retention slot on the periphery of the rotor disk.

Rotor assembly with cooling air deflectors and method

An internally-cooled airfoil for a gas turbine engine, wherein the airfoil comprises at least one substantially chordwise-extending and internally-projecting stiffener located on an internal surface of one of the sidewalls and adjacent to the insert.

Internally cooled airfoil for a gas turbine engine and method

An internally cooled airfoil for a gas turbine engine and a method of cooling in which at least two substantially parallel passages are in fluid communication with an exit plenum and adapted to reduce stagnation and improve strength, particularly in wide chord blades.

Internally cooled gas turbine airfoil and method

A partial rib for use in a turbine blade is disclosed which provides one or more of improved strength, air flow distribution and cooling. In one embodiment, the rib has a height of between 0.3 and 0.9 of the airfoil height.

Internally cooled turbine blade

A rotor assembly for a gas turbine engine, the rotor assembly comprises a plurality of cooling air deflectors mounted on the rotor assembly to redirect air to a manifold at a bottom side of a corresponding blade retention slot on the periphery of the rotor disk.

An internally cooled airfoil for a gas turbine engine, wherein a plurality of elongated cooling fins are provided inside the concave sidewall.

An internally cooled airfoil for a gas turbine engine, wherein a plurality of elongated cooling fins are provided inside the concave sidewall between two crossovers.

Growing community of inventors

Verdun, Canada

Michael Leslie Clyde Papple