Abstract: The turbomachine includes a compressor, an inner annular casing, and an outer annular casing. The inner annular casing and the outer annular casing define at least one cavity therebetween. The clearance control system includes a manifold system including at least one conduit disposed within the cavities and configured to channel a flow of cooling fluid between the cavities. The clearance control system also includes an impingement system including a header and at least one plenum configured to channel the flow of cooling fluid to the inner annular casing. The conduits configured to channel the flow of cooling fluid to the impingement system. The clearance control system further includes a channel system including at least one channels configured to channel the flow of cooling fluid to the turbomachine. The channels are configured to control the flow of cooling fluid to the manifold system.

Abstract: An apparatus and method of cooling a hot portion of a gas turbine engine, such as a multi-stage compressor of a gas turbine engine, by reducing an operating air temperature in a space between a seal and a blade post of adjacent stages by routing cooling air through an inlet in the vane, passing the routed cooling air through the vane, and emitting the routed cooling air into the space.

Abstract: A method of making a shroud assembly by: forming a shroud frame consisting of a one-piece hook and rail assembly which includes a first rail and a second rail disposed at lateral edges and a first support and a second support connected to the first rail and the second rail, the first and second rails and the first and second supports being integrally joined to one another and defining a central aperture therebetween; joining a metallic backsheet to the shroud frame covering the central aperture; and attaching a honeycomb rubs strip to the metallic backsheet via an upper edge of the honeycomb seal structure.

Abstract: A turbine shroud apparatus for a gas turbine engine having a centerline axis includes: a shroud segment having: an arcuate body extending axially between forward and aft ends and laterally between opposed end faces, wherein each of the end faces includes seal slots formed therein; and an arcuate stationary seal member mounted to the body; a turbine vane disposed axially aft of the shroud segment; and a casing surrounding the shroud segment and the turbine vane; wherein the turbine vane is mounted to the case so as to bear against the stationary seal member, compressing it and forcing the shroud segment radially outward against the casing.

Abstract: A turbine shroud assembly comprises a honeycomb rub strip, a metallic backsheet disposed along an upper edge of said honeycomb seal structure, a shroud frame having a first rail and a second rail disposed at lateral edges of said metallic backsheet, a first support and a second support connected to the first rail and the second rail, a spline extending axially in the first rail and the second rail along laterally outer surfaces of the first rail and the second rail and, a spline seal having a first edge and a second opposite edge, the first edge disposed in each of the spline grooves and the second opposite edge being capable of positioning in an adjacent shroud assembly.

Abstract: A turbine shroud apparatus for a gas turbine engine having a centerline axis includes: a shroud segment having: an arcuate body extending axially between forward and aft ends and laterally between opposed end faces, wherein each of the end faces includes seal slots formed therein; and an arcuate stationary seal member mounted to the body; a turbine vane disposed axially aft of the shroud segment; and a casing surrounding the shroud segment and the turbine vane; wherein the turbine vane is mounted to the case so as to bear against the stationary seal member, compressing it and forcing the shroud segment radially outward against the casing.

Abstract: An airfoil for a gas turbine engine including an inflection that facilitates enhancing film cooling of the airfoil, without adversely affecting aerodynamic efficiency of airfoil is described. The airfoil includes a generally concave first sidewall and a generally convex second sidewall joined at a leading edge and at a trailing edge of the airfoil. A plurality of cooling openings extend between an internal cooling chamber and an external surface of the first sidewall. One cooling opening extends from the cooling chamber into the inflection at a relatively shallow injection angle with respect the airfoil external surface.

Abstract: A turbine cooling circuit includes a flow sleeve having an aft end for adjoining the rim of a turbine disk. A seal rotor is spaced from the aft end adjacent a row of inlet holes. An inducer has an outlet disposed radially outwardly from the sleeve inlet holes. And, a seal stator surrounds the seal rotor to define a rotary seal disposed radially inwardly from the inducer outlet.

Abstract: An airfoil for a gas turbine engine including an inflection that facilitates enhancing film cooling of the airfoil, without adversely affecting aerodynamic efficiency of airfoil is described. The airfoil includes a generally concave first sidewall and a generally convex second sidewall joined at a leading edge and at a trailing edge of the airfoil. A plurality of cooling openings extend between an internal cooling chamber and an external surface of the first sidewall.

Abstract: A turbine cooling circuit includes a flow sleeve having an aft end for adjoining the rim of a turbine disk. A seal rotor is spaced from the aft end adjacent a row of inlet holes. An inducer has an outlet disposed radially outwardly from the sleeve inlet holes. And, a seal stator surrounds the seal rotor to define a rotary seal disposed radially inwardly from the inducer outlet.

Abstract: A rotor blade for a gas turbine engine including a tip region that facilitates reducing operating temperatures of the rotor blade is described. The tip region includes a first tip wall and a second tip wall extending radially outward from a tip plate of an airfoil. The tip walls extend from adjacent a leading edge of the airfoil to connect at a trailing edge of the airfoil. A notch is defined between the first and second tip walls at the airfoil leading edge. At least a portion of the second tip wall is recessed to define a tip shelf.

Abstract: A rotor blade for a gas turbine engine including a tip region that facilitates reducing operating temperatures of the rotor blade is described. The tip region includes a first tip wall and a second tip wall extending radially outward from a tip plate of an airfoil. The tip walls extend from adjacent a leading edge of the airfoil to connect at a trailing edge of the airfoil. A notch is defined between the first and second tip walls at the airfoil leading edge. At least a portion of the second tip wall is recessed to define a tip shelf.

Abstract: A gas turbine engine rotor assembly includes a plurality of aerodynamic devices to direct airflow radially inward. The gas turbine engine rotor assembly includes a rotor shaft that includes a plurality of openings. The aerodynamic devices include a pair of vane segments and a pair of sidewalls. A contoured outer surface includes an opening and permits the aerodynamic device to be positioned against an inner surface of the rotor shaft, and a flange ring defines a pocket. The aerodynamic device fits within the pocket to concentrically align the openings.

Abstract: A segmented turbine cooling component, such as shroud segment for the high pressure and low pressure turbine sections of a gas turbine engine, useful in providing preferential cooling to the side rails or panels of the turbine component through the combination of cooling air fed into the intermediate pressure cavity between the discourager and primary seals and then through recesses in the bottom wall of the lower discourager seal slot that impinges air on the portion of the side panel below the discourager seal. The side panels of the turbine component have an upper primary seal slot and a lower discourager seal slot with a bottom wall having a plurality of alternating lands and recesses along the length thereof and a plurality of cooling air passages having outlets exiting into the lower slot above the lands.