United Technologies Corporation

Rtx Corporation

Raytheon Technologies Corporation

Ratheyon Technologies Corporation

Craig R McGarrah

Timothy M Davis

Carson A Roy Thill

Mark Broomer

Mark John Rogers

Russell E Keene

Tracy A Propheter-Hinckley

Dominic J Mongillo, Jr

Edward F Pietraszkiewicz

Shawn J Gregg

Jeffrey R Levine

Richard M Salzillo

Anthony B Swift

Lisa P O'Neill

Kirk David Hlavaty

Robert A Charbonneau

Kenneth A Lonczak

Heather Ann Terry

Kenneth P Botticello

Brandon W Spangler

Thomas E Clark

Jose R Paulino

William M Barker

David Wasserman

Anthony P Cherolis

Edwin Otero

Daniel J Whitney

Joseph F Englehart

Kevin W White

Edward R Bareiss

Robert L Hazzard

Kyle J Brevick

Egon Salimusaj

Timothy A Dunbar

Conway Choung

A method of forming a vane assembly for a gas turbine engine includes the steps of 1) determining a re-stagger angle for an airfoil on a ceramic matrix composite ('CMC') vane, 2) determine an adjusted position of a baffle to be inserted within a central chamber in the vane, inserting the baffle into a bathtub seal, and adjusting an upper face of the baffle to an orientation relative to a bathtub seal support face to account for the adjusted position of the baffle and 3) fixing the baffle upper face to the bathtub seal flange support face. A vane assembly and a gas turbine engine are also disclosed.

CMC vane with rotatable baffle design to accommodate re-stagger

A vane for a gas turbine engine, the vane including: an inner diameter platform and an outer diameter platform and an airfoil that extends between the inner diameter platform and the outer diameter platform, the airfoil having a leading edge, a trailing edge, a pressure side and a suction side, the inner diameter platform has a suction side mate face and a pressure side mate face, wherein the suction side mate face and the pressure side mate face of the inner diameter platform have an axial feather seal slot and a radial seal slot, the axial feather seal slot includes a leading edge portion that extends from the radial seal slot towards a leading edge of the inner diameter platform and a trailing edge portion that extends from the radial seal slot towards a trailing edge of the inner diameter platform, and a chamfer extends between the radial seal slot and the trailing edge portion of the axial feather seal slot.

Component with feather seal slots for a gas turbine engine

A blade outer air seal assembly includes a seal segment that has a base portion that extends between a first circumferential side and a second circumferential side and from a first axial side to a second axial side. A first wall is axially spaced from a second wall. The first and second walls extend from the base portion to an outer wall to define a circumferentially extending seal passage. A carrier has a platform. A portion of the platform is within the seal passage. A radially extending first hook protrudes from the platform and is configured to engage with a support structure.

Carrier for blade outer air seal

A seal for sealing a radially outer component of a gas turbine engine stator to a radially inner component thereof includes an axially resilient seal carrier adapted for mounting the seal carrier and including at a radially inner portion thereof, a pair of radially spaced, axially extending, radially resilient jaws adapted to clamp a sealing element such as a rope seal there between in sealing engagement with a radially inner component of the engine stator.

Gas turbine engine seal

A component for a gas turbine engine is disclosed. The component includes a platform and a rail having an outer radial surface and an inner radial surface, with the inner radial surface connected to the platform. The rail includes a plurality of apertures spaced circumferentially between the outer radial surface and the inner radial surface, the apertures disposed to form a truss-like structure.

Support rail truss for gas turbine engines

An assembly for a turbine engine includes a turbine engine first component, a turbine engine second component and a seal assembly. The first component includes a groove and a groove surface. The second component includes a tongue that extends into the groove to a tongue surface. The seal assembly at least partially seals a gap between the groove surface and the tongue surface. The seal assembly includes a rope seal and a clip that attaches the rope seal to the tongue. The rope seal is arranged within the groove between the groove surface and the tongue surface.

Assembly for sealing a gap between components of a turbine engine

A seal assembly for a gas turbine engine includes an engine static structure. First and second members fluidly separate cavities from one another. A seal assembly is captured by the engine structure. The seal assembly includes a carrier and a seal that engages the first member. The second member is captured by the carrier.

Gas turbine engine ring seal

A seal assembly is positioned within a cavity that extends circumferentially about an axial centerline of a gas turbine engine. The cavity includes a cavity wall. The seal assembly includes a seal and a seal protector. The seal extends circumferentially within the cavity. The seal protector extends circumferentially within the cavity. The seal protector is positioned between the seal and the cavity wall. The seal protector includes a locating feature that is operative to contact the seal to aid in axially positioning the seal protector relative to the seal.

Gas turbine engine seal assembly

An interface within a gas turbine engine includes a multiple of segmented components, each with a segment flange with a multiple of apertures, at least one of the multiple of apertures a first slot aperture. A full ring component with a ring flange that defines a multiple ring of apertures, at least one of the multiple of ring apertures a second slot aperture, the second slot aperture transverse to the first slot aperture.

Low leakage multi-directional interface for a gas turbine engine

An assembly for a turbine engine includes a turbine engine first component, a turbine engine second component and a flexible seal that is attached to the first component. The flexible seal at least partially seals a gap between the first component and the second component. The flexible seal includes a mount and a finger seal that sealingly engages the second component. The mount includes a boss that sealingly engages the first component.

Flexible finger seal for sealing a gap between turbine engine components

A molding technique is utilized to form the cooling channels in an airfoil for a gas turbine engine. The cooling channels are formed by lost core mold cores, which are leached away after metal has been molded around them. At least one of the mold cores is provided with a cone which will extend into the airfoil. The cone forms an opening in a wall of the airfoil, and the size of this opening is indicative of the thickness of the wall. By comparing the size to expected sizes, a determination can be made of whether the wall is of an acceptable thickness at locations where it is difficult to otherwise measure the wall.

Method for checking wall thickness of hollow core airfoil

The present invention provides a turbine blade having a revised under-platform structure including a unique coating combination that reduces mechanical stress factors within the turbine blade. The turbine blade includes a platform with an airfoil extending upwardly from the airfoil and a root portion extending downwardly from the platform. Two suction side tabs extend a first distance outward from a suction side of the root potion. Two pressure side tabs extend outward from a pressure side of the root portion. One of the two pressure side tabs extends outward a distance similar to the first distance, however, the other of the two pressure side tabs extends outward a distance much smaller than the first distance, which reduces stresses acting on the turbine blade. In addition, a plurality of coatings are systematically applied to the turbine blade to further reduce mechanical stress factors and improve cooling.

Turbine blade including revised platform

The present invention provides an improved cooling circuit for a trailing edge of a turbine blade. The cooling circuit includes an inlet passage that receives a airflow and distributes the airflow through a feed passage. The feed passage primarily includes trip strips, at least one barrier including cross-over holes, teardrop shaped protrusions, and pockets disposed along a trailing edge. The geometry and positioning of both the cross-over holes and teardrop shaped protrusions downstream of the cross-over holes have been optimized to maximize cooling efficiency and reduce airflow. An improved transition between the inlet passage and the feed passage is also provided, which is arcuate and allows the airflow to maintain attachment and flow unimpeded from the inlet passage to the feed passage. The geometry of the pockets disposed along the trailing edge is optimized to improve cooling.

Turbine blade including revised trailing edge cooling

Growing community of inventors

Southington, CT, United States of America

Craig R McGarrah