United Technologies Corporation

Raytheon Technologies Corporation

Wieslaw A Chlus

Seth J Thomen

Stanley J Funk

Edward F Pietraszkiewicz

Anthony P Cherolis

David P Houston

Dominic J Mongillo, Jr

Richard H Page

Jesse R Christophel

Joseph F Englehart

Kenneth M Boucher

Clement A Senatro

Michael Gonsor

Arthur W Lucas, Jr

Erik R Granstrand

David M Konopka

Paul D Kudra

Luke A Hmiel

Stephen E Potz

Joan M Gudridge

A blade assembly includes a blade and a blade platform secured to the blade. The blade extends radially from the blade platform. The blade platform includes at least one platform airflow passage located therein. A gusset extends from the blade to the blade platform. The gusset includes a gusset airflow passage fluidly connected to the platform airflow passage to convey an airflow to the platform airflow passage. a gas turbine engine includes a combustor and a plurality of gas turbine engine components located in fluid communication with the combustor. The gas turbine engine component includes an airfoil portion and a platform secured to the airfoil portion. The platform includes at least one platform airflow passage positioned therein. A gusset extends from the airfoil portion to the platform. The gusset includes a gusset airflow passage fluidly connected to the platform airflow passage to convey an airflow to the platform airflow passage.

Blade platform gusset with internal cooling

An airfoil includes a cooling air passage for receiving a cooling air flow. A chevron including a first rib and a second rib extends from a common tip is disposed within the cooling passage for generating a turbulent flow to improve heat transfer. The chevron includes an angle between the first rib and the second rib that is greater than 90 degrees.

Obtuse angle chevron trip strip

An airfoil for a gas turbine engine includes pressure and suction walls spaced apart from one another and joined at leading and trailing edges to provide an airfoil that extends in a radial direction. The airfoil has a cooling passage arranged between the pressure and suction walls that extend toward a tip of the airfoil. The tip includes a pocket that separates the pressure and suction walls. Scarfed cooling holes fluidly connect the cooling passage to the pocket. The scarfed cooling holes include a portion that is recessed into a face of the suction wall and exposed to the pocket.

Suction side turbine blade tip cooling

An airfoil component for a gas turbine engine includes a platform joined to an airfoil. The platform includes a flow path surface that extends between spaced apart lateral surfaces. The airfoil extends from the flow path surface. A contoured surface adjoins the flow path surface and one of the lateral surfaces.

Gas turbine engine stator vane platform cooling

A gas turbine engine stage includes a rotor having a slot. A blade has a root received in the slot, and a shank that extends radially outward from the root to a platform that supports an airfoil. A seal is supported on the shank and in engagement circumferentially between and with the shank and the rotor within the slot. A gas turbine engine includes compressor and turbine sections. The turbine section has a rotor with a slot. A combustor is provided axially between the compressor and turbine sections. A turbine blade in the turbine section includes a root received in the slot and a shank that extends radially outward from the root to a platform that supports an airfoil.

Turbine under platform air seal strip

An airfoil assembly for a gas turbine engine is disclosed and includes a platform portion defining a portion of a gas flow path and a root portion for attachment of the turbine airfoil, the root portion including a bottom surface including a bottom area and a plurality of inlets that define a total inlet area as a ratio of the inlet area to the bottom area. An airfoil extends from the platform and including a plurality of cooling air passages in communication with the plurality of inlets.

Bell mouth inlet for turbine blade

A gas turbine engine has a turbine blade having a tip cap to close off internal cooling passages. The tip cap is formed with a plurality of purge holes, with there being at least two rows of purge holes. The increased number of purge holes spreads the cooling air outwardly across more of the surface area of the tip cap, and provides the tip cap with a better ability to withstand the extreme temperatures it faces in use.

Turbine blade with improved durability tip cap

A rotor assembly having a seal member in the root section of a rotor blade is disclosed. Various construction details are developed for blocking the flow of gases between adjacent rotor blades. In one detailed embodiment, a deformable seal member formed of a high temperature material is disposed between the root sections of adjacent rotor blades and engages the blades under operative conditions.

Rotor assembly for a rotary machine

A blade for a gas turbine engine includes a trailing edge cooling chamber, wherein the pressure and suction sides are connected by an array of spaced pedestals. The pedestals in a region most subject to thermal stresses are formed to be elliptical, such that the thermal stresses are spread over a larger area. Thus, stress concentrations are reduced. Also, various sized cylindrical pedestals may be used at other locations.

Airfoil with shaped trailing edge pedestals

A blade has an airfoil body having an internal cooling passageway network and a body tip pocket. At least one plate is secured within the body tip pocket and has inboard and outboard surfaces. A recess is in the outboard surface and an associated protrusion is on the inboard surface.

Turbine blade

A turbine blade tip plenum has means for preferentially directing or diverting cooling air from a cooling passageway network to cool areas subject to extreme heat.

A turbine blade has an undercut beneath its platform at a trailing edge. The undercut has a complex shape to move thermal stress concentration away from the platform. A preferred undercut shape has a pair of curved fillets at upper and lower extents of the undercut with an intermediate straight section. The intermediate straight section is preferably parallel to a principle stress field of the platform.

Turbine blade with trailing edge platform undercut

A turbine blade has a platform and an airfoil extending from a root at the platform to a tip. The airfoil has an internal cooling passageway network including at least one trailing edge cavity. A group of trailing edge holes extends from the trailing edge to the trailing edge cavity and a group of tip holes extends from the tip to the trailing edge cavity.

The present invention relates to an axial flow rotary machine and to a maintainable compressor bleed system therefore. The axial flow rotary machine comprises a primary flow path for a working medium, a secondary flow path for the working medium, a bleed system for allowing a bleed portion of the working medium flowing in the primary flow path to be delivered to the secondary flow path, a support structure for the bleed system located radially inboard of the bleed system, and the panels for gaining access to the bleed system. The panels are located radially outboard of the bleed system. The panels form the inner flow wall for the primary flow path.

Maintainable compressor stability bleed system

An integrated damper and windage cover 52 has a windage cover portion 56 cantilevered at the upstream end, free of contact with the blade platform 24, 32. The contact portion 54 is rigid and bears against the underside of the blade platform 24, spanning blade platform clearance 50.

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Wethersfield, CT, United States of America

Wieslaw A Chlus