Abstract: One embodiment of the present invention is a unique turbine blade for a gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and turbine blades for gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: A cooling apparatus bifurcates and regulates cooling airflow provided to a mid-turbine frame. The cooling apparatus includes a flow metering tube and a metering plate. The flow metering tube includes a top portion and a tube portion, wherein the top portion includes a first central aperture that directs a first cooling airflow into the tube portion and a first plurality of apertures located circumferentially around the central aperture that directs a second cooling airflow to a portion outside the tube portion. The metering plate is located on the top portion of the flow metering tube, wherein the metering plate includes a second central aperture aligned with the central aperture of the flow metering tube and a second plurality of apertures located circumferentially around the central aperture, wherein a size of the second central aperture meters the first cooling airflow and a size and number of the second plurality of apertures meters the second cooling airflow.

Abstract: A gas turbine engine component includes a cooling hole. The component includes a first wall having an inlet, a second wall having an outlet and a metering section extending downstream from the inlet and having a substantially convex first surface and a substantially concave second surface. The component also includes a diffusing section extending from the metering section to the outlet. A gas turbine engine wall includes first and second surfaces and a cooling hole extending between an inlet at the first surface and an outlet at the second surface. The cooling hole includes a metering section commencing at the inlet and a diffusing section in communication with the metering section and terminating at the outlet. The metering section includes a top portion having a first arcuate surface and a bottom portion having a second arcuate surface. The first and second arcuate surfaces have arcs extending in substantially similar directions.

Abstract: A gas turbine engine comprises a compressor, a combustion chamber, an outer casing, an inner casing and a cooling arrangement. The outer casing surrounds the compressor and the combustion chamber and the combustion chamber has turbine nozzle guide vanes. The compressor has load carrying outlet guide vanes connected to the outer casing and the inner casing. The turbine nozzle guide vanes connect the outer casing and the inner casing. The cooling arrangement comprises a cooling air duct located between the compressor and the combustion chamber. The compressor outlet guide vanes carry at least one aerodynamic fairing. A support structure supports the cooling air duct from the inner casing at two spaced positions and the support structure forms a chamber with the inner casing. The support structure comprises at least one hollow duct and each hollow duct locates behind a respective one of the aerodynamic fairings.

Abstract: The present invention is an article containing internal cooling channels located near at least one surface. In an embodiment, the cooled article includes a base material, a first layer, and a second layer. Here, the first layer is bonded to the base material and the second layer is bonded to the first layer, wherein at least one closed cooling channel is disposed within a portion of the first layer and a portion of the second layer.

Abstract: A gas turbine engine combustion chamber heat shield and seal assembly comprises a heat shield and a seal. The heat shield has an aperture and the seal is located in the aperture in the heat shield. The seal comprises an annular member having an upstream end, a middle and a downstream end. The upstream end of the seal has a diameter greater than the diameter of the aperture in the heat shield, the middle has a diameter less than the diameter of the aperture in the heat shield and the downstream end of the seal has a diameter greater than the diameter of the aperture in the heat shield.

Abstract: A cooling air flow ejector has a primary nozzle position to entrain air from a secondary nozzle. Mixed air is provided into a downstream flow conduit to be directed to an exhaust liner for a gas turbine engine. A variable area device controls a volume of air passing through a primary nozzle. A control for the variable area device to control the size of an orifice within the variable area device controls the volume of air reaching the exhaust liner to be cooled.

Abstract: An axial-flow turbine according to an embodiment includes a plurality of nozzle structures and a plurality of blade structures. At least one nozzle structure includes an outer ring diaphragm and an inner ring diaphragm. The outer ring diaphragm and the inner ring diaphragm form an annular opening portion which extends in a circumferential direction therebetween. A nozzle is provided in a portion of a region of the annular opening portion in the circumferential direction, and a closing part is provided in another portion of the region of the annular opening portion in the circumferential direction. The closing part closes this other portion of the region to prevent a working fluid from flowing into this other portion of the region. A closing part medium passage is provided in the closing part and is configured to flow a cooling medium which cools the closing part.

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: 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: Fuel (12) is supplied to a rotatable portion (118) of a gas turbine engine (10) comprising a rotor (24) and at least one blade (26, 26.1) operatively coupled thereto, so as to provide for cooling at least one of the rotor (24) or the at least one blade (26, 26.1) by transforming the fuel (12) to a vapor or gaseous state. The fuel (12) is discharged in a vapor or gaseous state from the rotatable portion (118) directly into a combustion chamber (16) of the gas turbine engine (10).

Abstract: An axial-flow turbine has turbine nozzles, a heat shield plate, a first communication hole formed in the turbine rotor and connected to the space, to flow a cooling medium, a first opening formed in at least any one of the two adjacent rotor disks, to be connected to the space, a second communication hole connected to the space through the first opening, to communicate with an implant unit of the turbine rotor blade in the rotor disk, a third communication hole connected to the second communication hole, to communicate along an effective length of the turbine rotor blade, a second opening formed in a side face of the turbine rotor blade, to be connected to the third communication hole, and a third opening formed in an outer circumferential end face of the turbine rotor blade, to be connected to the third communication hole.

Abstract: A gas turbine engine includes a fan, a compressor section fluidly connected to the fan, a combustor fluidly connected to the compressor section, a turbine section fluidly connected to the combustor, and a buffer system. The buffer system includes a heat exchanger having a first inlet, a first outlet, a second inlet, and a second outlet. The first outlet is configured to provide a cooled pressurized fluid. The buffer system includes first and second air sources that are selectively fluidly coupled to the first inlet, and a third air source that are fluidly coupled to the second inlet. Multiple fluid-supplied areas are located remotely from one another and are fluidly coupled to the first outlet. The multiple fluid-supplied areas include a bearing compartment. A method and a buffer system are also disclosed.

Abstract: A heat exchange module is provided for a turbine engine. The heat exchange module includes a duct and a plurality of heat exchangers. The duct includes a flowpath defined radially between a plurality of concentric duct walls. The flowpath extends along an axial centerline through the duct between a first duct end and a second duct end. The heat exchangers are located within the flowpath, and arranged circumferentially around the centerline.

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a body portion having an exterior surface and an internal surface. A cavity is disposed inside of the body portion. A cooling hole extends between the exterior surface and the internal surface and includes a metering section having an outlet and an inlet. The inlet is shaped dissimilar to the outlet.

Abstract: Bi-directional nacelle ventilation and cooling systems for use with aircraft and related methods are disclosed. An example apparatus includes a passageway to fluidly couple an opening formed in a nacelle of an aircraft engine and an engine compartment of the nacelle. The opening provides an inlet into the engine compartment when passive airflow is available to vent or cool the engine compartment and the opening to provide an outlet from the engine compartment when forced air is needed to vent or cool the engine compartment. A fan is positioned in the passageway to provide the forced air when the passive air is unavailable.

Abstract: A gas turbine may include turbine blades configured to improve stream adhesion by selectively attracting or reducing repulsion of charged particles carried by a combustion gas stream.

Abstract: A gas turbine engine with a closed loop liquid metal cooling fluid system for cooling stator vanes within the turbine, in which the stator vanes include a liquid metal cooling passage lined with Tantalum or Molybdenum and a liquid metal cooling fluid of Bismuth or Lead or Zinc or Tin or alloy mixtures of these metals.

Abstract: An aircraft compound cooling system includes a power thermal management system for cooling one or more aircraft components, an air cycle system, a vapor cycle system, and a turbine cooling circuit for cooling bleed air and cooling turbine components in a high pressure turbine in the engine. An air to air FLADE duct heat exchanger is disposed in a FLADE duct of the engine and a valving apparatus is operable for selectively switching the FLADE duct heat exchanger between the turbine cooling circuit and the air cycle system. A vapor cycle system includes a vapor cycle system condenser that may be in heat transfer cooling relationship with the air cycle system. An air cycle system heat exchanger and an engine burn fuel to air heat exchanger in the vapor cycle system condenser may be used for cooling a working fluid in a refrigeration loop of the vapor cycle system.

Abstract: A turbine exhaust casing having an outer casing, an inner casing, an annular exhaust gas path defined between outer and inner flow path walls, and a turbine exhaust casing cavity located radially outward and radially inward from the gas path. A plurality of structural struts support the inner casing to the outer casing, and a fairing surrounds each of the struts in an area extending between the outer and inner flow path walls. A first purge air path extends through at least one of the struts for conducting purge cooling air radially inward to the inner casing, and a second purge air path extends through the strut for further conducting the purge cooling air radially outward to provide a flow of purge air to a location of the exhaust casing cavity radially outward from the outer flow path wall.

Abstract: A reheat combustor for a gas turbine engine includes a fuel/gas mixer for mixing fuel, air and combustion gases produced by a primary combustor and expanded through a high pressure turbine. Fuel injectors inject fuel into the mixer together with spent cooling air previously used for convectively cooling the reheat combustor. The fuel mixture is burnt in an annular reheat combustion chamber prior to expansion through low pressure turbine inlet guide vanes. The fuel/gas mixer and optionally the combustion chamber define cooling paths through which cooling air flows to convectively cool their walls. The fuel injectors are also convectively cooled by the cooling air after it has passed through the fuel/gas mixer cooling paths. The low pressure turbine inlet guide vanes may also define convective cooling paths in series with the combustion chamber cooling paths.

Abstract: In a gas turbine, a plurality of burners, which are arranged concentrically to the rotational axis in a regular arrangement, each directing hot gas through an associated combustion chamber outlet into a turbine, at the inlet of which a second plurality of stator blades are arranged in a ring, uniformly spaced apart around the rotational axis. Cooling openings are provided, which are distributed over the circumference, through which cooling air is injected into the hot gas flow at the combustion chamber outlet. An improvement of the flow conditions in the hot gas is achieved by the cooling openings being divided into first cooling opening groups and second cooling opening groups. The arrangement of the first cooling opening groups corresponds to the arrangement of the stator blades, and in that the arrangement of the second cooling opening groups corresponds to the regular arrangement of the burners.

Abstract: A transition piece of a combustor which is sustainable even under conditions of more severe thermal environments is provided in the present invention. A transition piece of a combustor is provided with a cylindrical trunk main body, a cylindrical exit trunk part which is joined to the downstream end of the trunk main body and which cooperates with the trunk main body to constitute a trunk part, and an inner flange which extends from the downstream end part of the exit trunk part toward the outer periphery side of the exit trunk part. The exit trunk part and the inner flange are of an one-piece product. On the exit trunk part, there is formed a groove, and there is formed a cooling fluid passage which opens at the groove.

Abstract: A system includes a turbine combustor that includes a head end portion having a head end chamber, a combustion portion having a combustion chamber disposed downstream from the head end chamber, a cap disposed between the head end chamber and the combustion chamber, and a flow separator configured to separate a first exhaust flow from an oxidant flow. The flow separator is configured to direct the first exhaust flow into the head end chamber. The turbine combustor also includes a mixing region configured to mix the first exhaust flow with the oxidant flow to provide an oxidant-exhaust mixture.

Abstract: A turbine exhaust cylinder for an industrial gas turbine engine with an external blower that delivers cooling air to a heat shield jacket surrounding an outer diameter cylinder to provide impingement cooling for the cylinder. Spent impingement cooling air is collected as passed through a space formed between a fairing and a strut to provide cooling to these parts. The cooling air then provides cooling for the inner diameter cylinder before being discharged into the turbine exhaust gas or out from the turbine exhaust cylinder.

Abstract: One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique cooling system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for cooling one or more objects of cooling. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: 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 having an exterior surface that extends in a radial direction to a tip. A tip trench is provided in the tip and wrapping at least a portion of the airfoil from the pressure side wall around the leading edge to the suction side wall. The tip trench is provided by a recess.

Abstract: An oscillating heat pipe of a gas turbine engine includes a plurality of channels that define a continuous loop through which a fluid flows, and an evaporator of a gas turbine engine. The fluid flows through the evaporator to accept heat from a first fluid. The first fluid is located near or in an engine core. The oscillating heat pipe also includes condenser of the gas turbine engine. The fluid flows through the condenser to reject heat to a second fluid, and the second fluid is located outwardly of the engine core.

Abstract: An embodiment of the present invention is a gas turbine engine including a compressor, a turbine, an annular combustor, an exhaust duct, a first engine shaft bearing, and a second engine shaft bearing. The turbine has an axial flow direction toward the compressor. The combustor has an axial flow direction away from the compressor. The exhaust duct is disposed between the compressor and the combustor. The first engine shaft bearing is disposed on an axial side of the compressor opposite the turbine. The second engine shaft bearing is disposed on an axial side of the turbine opposite the compressor.

Abstract: An integrated turbomachine oxygen plant includes a turbomachine and an air separation unit. One or more compressor pathways flow compressed air from a compressor through one or more of a combustor and a turbine expander to cool the combustor and/or the turbine expander. An air separation unit is operably connected to the one or more compressor pathways and is configured to separate the compressed air into oxygen and oxygen-depleted air. A method of air separation in an integrated turbomachine oxygen plant includes compressing a flow of air in a compressor of a turbomachine. The compressed flow of air is flowed through one or more of a combustor and a turbine expander of the turbomachine to cool the combustor and/or the turbine expander. The compressed flow of air is directed to an air separation unit and is separated into oxygen and oxygen-depleted air.

Abstract: A system for removing heat from a turbine includes a component in the turbine having a supply plenum and a return plenum therein. A substrate that defines a shape of the component has an inner surface and an outer surface. A coating applied to the outer surface of the substrate has an interior surface facing the outer surface of the substrate and an exterior surface opposed to the interior surface. A first fluid channel is between the outer surface of the substrate and the exterior surface of the coating. A first fluid path is from the supply plenum, through the substrate, and into the first fluid channel, and a second fluid path is from the first fluid channel, through the substrate, and into the return plenum.

Abstract: The present invention is a coated article containing at least two adjacent cooling holes that are substantially uncoated. In an embodiment, the coated article includes a base material, an outer surface and a coating on a portion of the outer surface. Here, the adjacent cooling holes are arranged along an axis on the outer surface and extend within the base material at an angle to the outer surface. The coating may include at least one unfinished edge oriented substantially parallel to the axis and set off at an oblique angle and a distance from the adjacent cooling holes.

Abstract: In a high pressure turbine used in combination with a contra-rotating low pressure turbine, in which the geometry of the inlet guide vane of the low pressure turbine induces a reflection of the incident shock wave from the high pressure turbine airfoil that propagates back upstream to impact the blade thereby denigrating both the performance of the high pressure turbine and its durability. The shock wave reflection is eliminated by fitting the downstream vanes with holes or indentations that allow jets of compressor bleed air to penetrate the boundary layer and induce a net velocity component across the main-flow direction in the near-wall region. This induced flow component acts to cancel the net flow component created by the moving shock from the upstream blade row thereby obviating the reflection that occurs naturally to cancel the flow.

Abstract: The invention concerns a turbojet for aircraft including engine located in nacelle, and thermal exchanger intended to cool a fluid participating in the engine propulsive system, characterized in that said thermal exchanger is located on engine external wall, an interstitial space within which air can circulate being arranged between the engine external wall and a lower wall of said thermal exchanger. The invention also concerns an aircraft provided with at least one such turbojet.

Abstract: A turbine has a first and second inner wall, an inner casing and a shield element. The first and second inner walls are mounted to the inner casing such that an inner volume for working fluid is separated from an outer volume for cooling fluid. The first and second inner walls and inner casing are arranged to form a cavity in outer volume. The shield element is arranged inside cavity to separate cavity in an inner and outer region in a radially outer position to the inner region. A gap is formed between the first and second inner walls for working fluid between the inner volume and inner region. The shield element is arranged in cavity to form a fluid inlet for injecting cooling fluid from outer region to inner region for generating a predefined recirculation of working fluid and cooling fluid inside inner region.

Abstract: A gas turbine engine component includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally from the metering section and a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section.

Abstract: An operation method for a gas turbine is provided which includes compressing a working fluid by a compressor, feeding this compressed working fluid into at least one combustion chamber where it is overheated, expanding this overheated working fluid in at least one expansion turbine to produce energy, carrying out a first tapping of the compressed work fluid from the compressor to feed it into a first cavity of the turbine for cooling, carrying out a second tapping of the working fluid downstream from the first tapping to feed it into a second cavity of the turbine upstream from the first cavity, for cooling, and fluidly connecting the first tapping to the second tapping to selectively feed the first tapping by a part of the second tapping during partial load operation conditions to keep the temperature of the first cavity within the acceptable limits for the resistance of the materials.

Abstract: A turbine nozzle having a non-linear cooling conduit is disclosed. In one embodiment, a turbine nozzle includes: an airfoil, at least one endwall adjacent the airfoil, and a fillet region connecting the airfoil and the at least one endwall, the fillet region including an outer surface. The turbine nozzle also includes a non-linear cooling conduit located within the fillet region and adjacent the outer surface of the fillet region, the non-linear cooling conduit allows fluid flow through the fillet region. The non-linear cooling conduit spans substantially along an axial length of the airfoil between a leading edge of the airfoil and a trailing edge of the airfoil. Additionally, the non-linear cooling conduit includes an arc profile substantially similar to an arc profile of the airfoil.

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: An end-wall component of the mainstream gas annulus of a gas turbine engine has a cooling arrangement including one or more circumferentially extending rows of ballistic cooling holes through which, in use, dilution cooling air is jetted into the mainstream gas to reduce the mainstream gas temperature adjacent the end-wall. A portion of the cooling holes are first cooling holes angled such that the direction of the dilution cooling air jetted therethrough has, on entry into the mainstream gas annulus, a component in one tangential direction. A portion of the cooling holes are second cooling holes angled such that the direction of the dilution cooling air jetted therethrough has, on entry into the mainstream gas annulus, a component in the opposite tangential direction. The first and second cooling holes are arranged such that the cooling air from jets having entry components in opposing tangential directions collide and coalesce.

Abstract: An air injection system for use in a gas turbine engine includes at least one outlet port through which air is extracted from the engine only during less than full load operation, at least one rotor cooling pipe, which is used to inject the air extracted from the outlet port(s) into a rotor chamber, a piping system that provides fluid communication between the one outlet port(s) and the rotor cooling pipe(s), a blower system for extracting air from the engine through the outlet port(s) and for conveying the extracted air 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 air from passing through the piping system, and open during less than full load engine operation to allow air to pass through the piping system.

Abstract: A gas turbomachine system includes a compressor portion including an inlet portion, a turbine portion fluidically connected to, and mechanically linked with, the compressor portion, and a combustor assembly including at least one combustor fluidically connected to the turbine portion. An inlet system is fluidically connected to the inlet portion of the compressor portion. The inlet system includes an inlet chiller. An inlet chiller condensate recovery system is fluidically connected to the inlet system. The inlet chiller condensate recovery system includes an inlet fluidically connected to the inlet chiller and an outlet fluidically connected to one of the compressor portion and the combustor assembly.

Abstract: A gas turbine includes a liner, a casing surrounding the liner, a hula seal flexibly connected to an aft end of the liner and a liner aft support mechanism. The liner is configured to receive compressed gas and fuel at an upstream end, the mixture of the compressed gas and the fuel being burned in a combustion core area of the liner to yield hot exhaust gasses. The liner aft end support mechanism is located downstream from an area where a highest temperature on an outer surface of the liner is attained, and upstream to a portion where the hula seal is connected to the liner, and is configured to movably support the liner inside the casing. The liner aft end support mechanism includes at least three individual support elements configured to allow a part of the individual support elements to move in the flow direction relative to at least one of the liner or the casing.

Abstract: A combustor includes a first shroud extending circumferentially inside the combustor and at least partially defining an inlet passage. A second shroud extends circumferentially inside the combustor. The second shroud defines an outlet passage. A first plate extends radially inside the second shroud downstream from the inlet passage of the first shroud and upstream from the outlet passage of the second shroud. The first plate generally defines an inlet port and an outlet port. A second plate extends radially around the first plate downstream from the inlet port and upstream from the outlet port of the first plate. A first fluid flow path extends from the inlet passage to the inlet port. A second fluid flow path extends from the outlet port to the outlet passage. A baffle extends from the first shroud to the first plate. The baffle separates the first and second fluid flow paths.

Abstract: A combustor of a combustion turbine engine is described. The combustor may include an inner radial wall, which defines a combustion chamber downstream of a primary fuel nozzle, and an outer radial wall, which surrounds the inner radial wall so to form a flow annulus therebetween, and the combustor may include a socket extending from the outer radial wall into the flow annulus. The socket may include: a mouth formed through the outer radial wall; a floor offset a predetermined distance from an outboard surface of the inner radial wall; impingement ports formed through the floor; and an axial nozzle.

Abstract: An accessory gear box configured to be mounted on a fan casing of a turbojet engine, including: an input shaft; a plurality of accessory drive shafts; and a compartment forming an interface with an accessory arranged on the housing of the accessory gear box at a site of the accessory, the interface compartment including an inlet and an outlet configured to allow air to circulate inside the compartment, to insulate the accessory gear box from heat generated by the accessories.

Abstract: A combustor generally includes a shroud that that defines at least one inlet passage extends circumferentially inside the combustor. A first plate extends radially inside the shroud downstream from the inlet passage. The first plate defines at least one inlet port, at least one outlet port and at least partially defines at least one fuel nozzle passage. The shroud at least partially surrounds a sleeve that extends around the fuel nozzle passage. A tube at least partially surrounded by the sleeve may extend through the fuel nozzle passage. The tube, the sleeve, and the first plate may at least partially define an outlet passage. A first fluid flow path generally extends from the at inlet passage to the inlet port, and a second fluid flow path extends generally from the outlet port to the outlet passage.

Abstract: A turbine vane for a gas turbine engine includes inner and outer platforms joined by a radially extending airfoil. The airfoil includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface. The inner and outer platforms respectively include inner and outer sets of film cooling holes. One of the inner and outer sets of film cooling holes are formed in substantial conformance with platform cooling hole locations described by one of the sets of Cartesian coordinates set forth in Tables 1 and 2. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate. The cooling holes with Cartesian coordinates in Tables 1 and 2 have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.08 mm).

Abstract: Embodiments of methods and apparatus for providing compressor extraction cooling are provided. According to one example embodiment, a method is disclosed for controlling compressor extraction cooling. The method can include providing a cooling medium. The method can include extracting air from a compressor associated with a gas turbine. The method can also include introducing the cooling medium to the compressor extraction air, wherein the compressor extraction air is cooled by the cooling medium prior to or during introduction to the turbine section. Furthermore, method can include selectively controlling at least one of the compressor extraction air or the cooling medium based at least in part on a characteristic associated with the gas turbine.

Abstract: A gas turbine engine component includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface and a diffusing section in communication with the inlet and extending to the outlet. The diffusing section includes a plurality of crenellation features that encourage lateral spreading of cooling air flowing through the cooling hole.