Abstract: A method of nondestructive evaluation and related system. The method includes arranging a test piece (14) having an internal passage (18) and an external surface (15) and a thermal calibrator (12) within a field of view (42) of an infrared sensor (44); generating a flow (16) of fluid characterized by a fluid temperature; exposing the test piece internal passage (18) and the thermal calibrator (12) to fluid from the flow (16); capturing infrared emission information of the test piece external surface (15) and of the thermal calibrator (12) simultaneously using the infrared sensor (44), wherein the test piece infrared emission information includes emission intensity information, and wherein the thermal calibrator infrared emission information includes a reference emission intensity associated with the fluid temperature; and normalizing the test piece emission intensity information against the reference emission intensity.

Abstract: A flow control insert for a turbine vane is disclosed. The flow control insert is designed to be placed in a trailing edge cooling passage of the turbine vane to simultaneously reduce cooling air flow and increase heat transfer from the vane to the cooling air, thereby improving efficiency of the turbine via reduced cooling air flow requirement while maintaining vane temperature within a specified range. Two different flow control insert designs are disclosed, where either design fits inside the cooling passage opening and allows an existing vane casting design to be used.

Abstract: A turbine airfoil assembly for installation in a gas turbine engine. The airfoil assembly includes an endwall and an airfoil extending radially outwardly from the endwall. The airfoil includes pressure and suction sidewalls defining chordally spaced apart leading and trailing edges of the airfoil. An airfoil mean line is defined located centrally between the pressure and suction sidewalls. An angle between the mean line and a line parallel to the engine axis at the leading and trailing edges defines gas flow entry angles, ?, and exit angles, ?. Airfoil inlet and exit angles are substantially in accordance with pairs of inlet angle values, ?, and exit angle values, ?, set forth in one of Tables 1, 3, 5 and 7.

Abstract: A seal assembly between a disc cavity and a hot gas path in a gas turbine engine includes a rotating blade assembly having a plurality of blades that rotate with a turbine rotor during operation of the engine, and a stationary vane assembly having a plurality of vanes and an inner shroud. The inner shroud includes a radially outwardly facing first surface, a radially inwardly facing second surface, and a plurality of grooves extending into the second surface. The grooves are arranged such that a space having a component in a circumferential direction is defined between adjacent grooves. During operation of the engine, the grooves guide purge air out of the disc cavity toward the hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path.

Abstract: A cooling channel (36, 36B) cools an exterior surface (40 or 42) or two opposed exterior surfaces (40 and 42). The channel has a near-wall inner surface (48, 50) with a width (W1). Interior side surfaces (52, 54) may converge to a reduced channel width (W2). The near-wall inner surface (48, 50) may have fins (44) aligned with a coolant flow (22). The fins may highest at mid-width of the near-wall inner surface. A two-sided cooling channel (36) may have two near-wall inner surfaces (48, 50) parallel to two respective exterior surfaces (40, 42), and may have an hourglass shaped transverse sectional profile. The tapered channel width (W1, W2) and the fin height profile (56A, 56B) increases cooling flow (22) into the corners (C) of the channel for more uniform and efficient cooling.

Abstract: An airfoil in a gas turbine engine includes an outer wall, a cooling fluid cavity, and a plurality of cooling fluid passages. The outer wall has a leading edge, a trailing edge, a pressure side, a suction side, and radially inner and outer ends. The cooling fluid cavity is defined in the outer wall, extends generally radially between the inner and outer ends of the outer wall, and receives cooling fluid for cooling the outer wall. The cooling fluid passages are in fluid communication with the cooling fluid cavity and include zigzagged passages that include alternating angled sections, each section having both a radial component and a chordal component. The cooling fluid passages extend from the cooling fluid cavity toward the trailing edge of the outer wall and receive cooling fluid from the cooling fluid cavity for cooling the outer wall near the trailing edge.

Abstract: A thermal management arrangement (110) in a gas turbine engine (60), including: a conduit-arrangement (62) providing fluid communication between a compressor section (156) and: a relatively thermally responsive portion (52) of a turbine vane carrier (10); and a relatively thermally unresponsive portion (48) of a first turbine vane carrier. The conduit-arrangement includes: a general cooling flow outlet (122) disposed proximate the relatively thermally responsive portion of the turbine vane carrier and configured to discharge a general cooling flow (124); and an impingement flow outlet (118) disposed proximate the relatively thermally unresponsive portion and configured to discharge an impingement flow (120). The thermal management arrangement is configured such that a flow rate of the impingement flow is effective to accelerate a thermal response of the relatively thermally unresponsive portion toward a thermal response of the relatively thermally responsive portion.

Abstract: A system for cooling a wall (24) of a component having an outer surface with raised ribs (12) defining a structural pocket (10), including: an inner wall (26) within the structural pocket and separating the wall outer surface within the pocket into a first region (28) outside of the inner wall and a second region (40) enclosed by the inner wall; a plate (14) disposed atop the raised ribs and enclosing the structural pocket, the plate having a plate impingement hole (16) to direct cooling air onto an impingement cooled area (38) of the first region; a cap having a skirt (50) in contact with the inner wall, the cap having a cap impingement hole (20) configured to direct the cooling air onto an impingement cooled area (44) of the second region, and; a film cooling hole (22) formed through the wall in the second region.

Abstract: An integrated thermal barrier coating and cooling flow metering plate for a turbine vane are disclosed. On an existing vane design, the thickness of the thermal barrier coating is increased in order to provide more thermal protection around the vane material itself. The increased insulation around the vane allows the volume of cooling air flow to be reduced, while still maintaining the vane temperature within specification. The reduced cooling air flow is obtained by adding a flow metering plate at the inlet of a vane trailing edge cooling circuit, thereby increasing turbine efficiency via reduced cooling air flow requirements, while allowing an existing vane casting design to be used.

Abstract: During full load operation of gas turbine engine operation, a valve system is maintained in a closed position to substantially prevent air from passing through a piping system of a shell air recirculation system. Upon initiation of a turn down operation, which is implemented to transition the engine to a turning gear state or a shut down state, the valve system is opened to allow air to pass through the piping system. A blower is operated to extract air through at least one outlet port of the shell air recirculation system from an interior volume of an engine casing portion associated with the combustion section, to convey the extracted air through the piping system, and to inject the air into the interior volume of the engine casing portion through at least one inlet port of the shell air recirculation system to circulate air within the engine casing portion.

Abstract: A cooling fluid air injection system for use in a gas turbine engine includes at an external cooling fluid source, at least one rotor cooling pipe, which is used to inject cooling fluid from the source into a rotor chamber, a piping system that provides fluid communication between the source and the rotor cooling pipe(s), a blower system for conveying the cooling fluid through the piping system and the rotor cooling pipe(s) into the rotor chamber, and a valve system. The valve system is closed during full load engine operation to prevent cooling fluid from the source from passing through the piping system, and open during less than full load engine operation to allow cooling fluid from the source to pass through the piping system.

Abstract: A triple hook ring segment including forward, midsection and aft mounting hooks for engagement with respective hangers formed on a ring segment carrier for supporting a ring segment panel, and defining a forward high pressure chamber and an aft low pressure chamber on opposing sides of the midsection mounting hook. An isolation plate is provided on the aft side of the midsection mounting hook to form an isolation chamber between the aft low pressure chamber and the ring segment panel. High pressure air is supplied to the forward chamber and flows to the isolation chamber through crossover passages in the midsection hook. The isolation chamber provides convection cooling air to an aft portion of the ring segment panel and enables a reduction of air pressure in the aft low pressure chamber to reduce leakage flow of cooling air from the ring segment.

Abstract: A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages effect a pumping of cooling fluid from the disc cavity toward the hot gas path during operation of the engine.

Abstract: A squealer tip formed from a pressure side tip wall and a suction side tip wall extending radially outward from a tip of the turbine blade is disclosed. The pressure and suction side tip walls may be positioned along the pressure sidewall and the suction sidewall of the turbine blade, respectively. The pressure side tip wall may include a chamfered leading edge with film cooling holes having exhaust outlets positioned therein. An axially extending tip wall may be formed from at least two outer linear surfaces joined together at an intersection forming a concave axially extending tip wall. The axially extending tip wall may include a convex inner surface forming a radially outer end to an inner cavity forming a cooling system. The cooling system may include one or more film cooling holes in the axially extending tip wall proximate to the suction sidewall, which promotes increased cooling at the pressure and suction sidewalls.

Abstract: A gas turbine engine blade (20), including: an airfoil (24) including a pressure side exterior surface (34), a suction side exterior surface (36), and a first rib (130) spanning between the pressure side exterior surface and the suction side exterior surface. The airfoil (24) is twisted from a base end (30) of the airfoil to a tip end (32) of the airfoil. The first rib is twisted from a base end of the first rib to a tip end of the first rib. The pressure side exterior surface, the suction side exterior surface, and the first rib are cast as a monolith.

Abstract: A casting core (200) for a twisted gas turbine engine blade, including: an airfoil portion (202) having: an airfoil base end (208), an airfoil tip end (210), a concave side exterior surface (212), a convex side exterior surface (214), a leading edge (204), and a trailing edge (206). The airfoil portion is twisted in a radial direction from the airfoil base end to the airfoil tip end. The airfoil portion includes a first void (220) between the concave side exterior surface and the convex side exterior surface and extending radially to define the shape of a rib of an airfoil to be cast around the core. A first leading edge surface and a first trailing edge surface of the void are twisted from the airfoil base end to the airfoil tip end.

Abstract: A mesh (35) of cooling channels (35A, 35B) with an array of cooling channel intersections (42) in a wall (21, 22) of a turbine component. A mixing chamber (42A-C) at each intersection is wider (W1, W2)) than a width (W) of each of the cooling channels connected to the mixing chamber. The mixing chamber promotes swirl, and slows the coolant for more efficient and uniform cooling. A series of cooling meshes (M1, M2) may be separated by mixing manifolds (44), which may have film cooling holes (46) and/or coolant refresher holes (48).

Abstract: A seal assembly between a disc cavity and a hot gas path in a gas turbine engine includes a stationary vane assembly and a rotating blade assembly axially upstream from the vane assembly. A platform of the blade assembly has a radially outwardly facing first surface, an axially downstream facing second surface defining an aft plane, and a plurality of grooves extending into the second surface such that the grooves are recessed from the aft plane The grooves are arranged such that a circumferential space is defined between adjacent grooves During operation of the engine, the grooves impart a circumferential velocity component to purge air flowing out of a disc cavity through the grooves to guide the purge air toward a hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path.

Abstract: A seal assembly between a disc cavity and a turbine section hot gas path includes a stationary vane assembly and a rotating blade assembly downstream from the vane assembly and including a plurality of blades that are supported on a platform and rotate with a turbine rotor and the platform during operation of the engine. The platform includes a radially outwardly facing first surface, a radially inwardly facing second surface, a third surface, and a plurality of grooves extending into the third surface. The grooves are arranged such that a space is defined between adjacent grooves. During operation of the engine, the grooves guide purge air out of the disc cavity toward the hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path.

Abstract: A cooling channel (36, 36B) cools an exterior surface (40 or 42) or two opposed exterior surfaces (40 and 42). The channel has a near-wall inner surface (48, 50) with a width (W1). Interior side surfaces (52, 54) may converge to a reduced channel width (W2). The near-wall inner surface (48, 50) may have fins (44) aligned with a coolant flow (22). The fins may highest at mid-width of the near-wall inner surface. A two-sided cooling channel (36) may have two near-wall inner surfaces (48, 50) parallel to two respective exterior surfaces (40, 42), and may have an hourglass shaped transverse sectional profile. The tapered channel width (W1, W2) and the fin height profile (56A, 56B) increases cooling flow (22) into the corners (C) of the channel for more uniform and efficient cooling.