United Technologies Corporation

Raytheon Technologies Corporation

Michael F Blair

Tracy A Propheter-Hinckley

Mark F Zelesky

Om Parkash Sharma

Dominic J Mongillo, Jr

Brandon W Spangler

Frederick M Schwarz

Paul M Lutjen

Thomas N Slavens

Joel H Wagner

Atul Kohli

William Abdel-Messeh

Susan M Tholen

Foster Philip Lamm

Justin D Piggush

Richard W Hoff

Richard H Page

David J Candelori

Samuel D Draper

Roger Gates

Raymond A Zatorski

Abbas Alahyari

A turbine engine system includes a heat source, a heat exchanger, a cooling medium inlet and a cooling medium outlet. The heat source includes a first passage. The heat exchanger includes a second passage and a third passage. The first and the second passages are configured in a closed loop circuit. The third passage is configured between the inlet and the outlet in an open loop circuit.

Turbine engine cooling system with an open loop circuit

Disclosed is a vane stage for a variable area turbine, comprising vane platforms comprising a first vane platform and a second vane platform, one of the vane platforms being a radial inner vane platform and another of the vane platforms being a radial outer vane platform, and a space therebetween defining an air flowpath; a primary vane body secured at opposing radial ends to the first vane platform and the second vane platform, a secondary vane body movably secured at the first vane platform, the secondary vane body being movable between a stowed position, wherein the secondary vane body is outside of the air flow path, and a deployed position, wherein the secondary vane body is extended between the plurality of vane platforms to form a vane trailing edge extension.

Vane for variable area turbine

A component according to an exemplary aspect of the present disclosure includes, among other things, a wall and a vascular engineered lattice structure formed inside of the wall. The vascular engineered lattice structure includes at least one of a hollow vascular structure and a solid vascular structure configured to communicate fluid through the vascular engineered lattice structure.

Gas turbine engine component having vascular engineered lattice structure

An exemplary gas turbine engine includes a turbine section operative to impart rotational energy to a compressor section. The turbine section includes at least a low-pressure turbine and a high-pressure turbine, and a number of stages in the low pressure turbine is from three to five.

Blades in a turbine section of a gas turbine engine

Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity are provided. In this regard, a representative turbine stage for a gas turbine engine includes a first set of rotatable blades operative to be positioned downstream of and adjacent to a first set of vanes, a number of blades of the first set of blades being less than a number of vanes of the first set of vanes.

Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity

An exemplary gas turbine engine includes a turbine section operative to impart rotational energy to a compressor section. The turbine section includes at least a low-pressure turbine and a high-pressure turbine, and a ratio of a number of stages in the low-pressure turbine to a number of stages in the high-pressure turbine is 2.

Cooled airfoils and gas turbine engine systems involving such airfoils are provided. In this regard, a representative cooled airfoil includes: an exterior surface defining a leading edge, a trailing edge, a suction side and a pressure side; an interior surface defining an interior cavity; trenches in the exterior surface oriented spanwise along the leading edge; and cooling holes communicating between the interior cavity and the trenches such that cooling air provided to the interior cavity flows from the interior cavity though the cooling holes into the trenches, the cooling holes having exterior apertures located in the trenches and interior apertures located at the interior surface.

Cooled airfoils and gas turbine engine systems involving such airfoils

A casting includes a wall thickness check feature for measuring thickness of a wall second aside an in-wall cooling passageway. The thickness is determined by observing the existence and/or size of an opening formed by the feature. The casting is cast from a pattern including portions forming the feature. To manufacture the pattern, a pattern-forming die is assembled with a ceramic feedcore and a refractory metal core (RMC). The assembling leaves an inlet portion of the RMC engaged to the ceramic feedcore and leaves an outlet portion of the RMC engaged to the die. A pattern-forming material is molded in the die at least partially over the ceramic feedcore and RMC. The die is disengaged from the pattern-forming material. The assembling engages a stepped projection of the RMC with a mating surface of the die.

Cooled wall thickness control

Turbine components, and in particular turbine blades, are provided with impingement cooling channels. Air is delivered along central channels, and the central channels deliver the air through crossover holes to core channels adjacent both a pressure wall and a suction wall. The air passing through the crossover holes impacts against a wall of the core channels.

Impingement skin core cooling for gas turbine engine blade

A turbine engine component has an airfoil portion having a pressure side and a suction side, a trailing edge discharge slot, and a suction side lip downstream of an exit of the trailing edge slot. The suction side lip is provided with negative features for increasing local heat transfer coefficient in the region of the suction side lip.

Local indented trailing edge heat transfer devices

A blade outer air seal (BOAS) casting core has first and second end portions and a plurality of legs. Of these legs, first legs each have: a proximal end joining the first end portion; a main body portion; and a free distal portion. Second legs each have: a proximal end joining the second end portion; a main body portion; and a free distal portion.

Blade outer air seal cores and manufacture methods

An embedded microcircuit for producing an improved cooling film over a surface of a part, comprising an inlet through which a coolant gas may enter, a circuit channel extending from the inlet through which the coolant gas may flow, and a slot film hole formed at a terminus of the circuit channel through which the coolant gas may exit a part.

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Manchester, CT, United States of America

Michael F Blair