Abstract: A process is provided for forming shaped air holes, such as for use in turbine blades. Aspects of the disclosure relate to forming shaped portions of air holes using a short pulse laser, forming a metered hole corresponding to each shaped portion, and separately finishing the shaped portion using a short-pulse laser. In other embodiments, the order of these operations may be varied, such as to form the shaped portions and to finish the shaped portions using the short-pulse laser prior to forming the corresponding metered holes.

Abstract: A rotor blade comprising an airfoil portion and a root portion, and an internal cooling circuit having flow passages in the root portion and the airfoil portion, wherein the internal cooling circuit includes: a first flow passage; and a non-integral plug. The plug may include a plug channel configured to correspond to a desired level of coolant flow through the first cooling passage. The plug may be connected to the rotor blade in a fixed blocking position relative to the first flow passage.

Abstract: A platform cooling arrangement in a turbine rotor blade having a platform at an interface between an airfoil and a root, wherein the rotor blade includes an interior cooling passage that extends from a connection at the root to the approximate radial height of the platform, wherein, the interior cooling passage comprises a high-pressure coolant region and a low-pressure coolant region, and wherein a pressure side of the platform comprises a topside extending circumferentially from the airfoil to a pressure side slashface. The platform cooling arrangement may include: a platform cavity formed within the pressure side of the platform, a high-pressure connector that connects the platform cavity to the high-pressure coolant region of the interior cooling passage; a low-pressure connector that connects the platform cavity to the low-pressure coolant region of the interior cooling passage; and a pin bank formed within the platform cavity that includes radial pins.

Abstract: In a turbine rotor blade having a platform at an interface between an airfoil and a root, wherein the airfoil and the root include an interior cooling passage formed therein, wherein, in operation, the interior cooling passage comprises at least a high-pressure coolant region and a low-pressure coolant region, platform cooling arrangement that includes: a platform slot extending circumferentially from a mouth formed through the pressure side slashface; a high-pressure connector that connects the platform slot to the high-pressure coolant region of the interior cooling passage; a low-pressure connector that connects the platform slot to the low-pressure coolant region of the interior cooling passage; and a platform cooling cartridge removably engaged within the platform slot, the platform cooling cartridge comprising one or more cartridge cooling channels.

Abstract: A platform cooling arrangement for a turbine rotor blade having a platform and an interior cooling passage and, in operation, a high-pressure coolant region and a low-pressure coolant region, wherein the platform includes a topside, which extends from the airfoil to a pressure side slashface, and an underside. The platform cooling arrangement may include: an airfoil manifold that resides near the junction of the pressure face of the airfoil and the platform; a slashface manifold that resides near the pressure side slashface; a high-pressure connector that connects the airfoil manifold to a high-pressure coolant region of the interior cooling passage; a low-pressure connector that connects the slashface manifold to a low-pressure coolant region of the interior cooling passage; cooling apertures that extend from a starting point along the pressure side slashface to a connection with the airfoil manifold, bisecting the slashface manifold therebetween; and a plurality of non-integral plugs.

Abstract: A platform cooling arrangement in a turbine rotor blade having a platform at an interface between an airfoil and a root, wherein the rotor blade includes an interior cooling passage that in operation, includes at least a high-pressure coolant region and a low-pressure coolant region, and wherein the platform includes a platform underside. The platform cooling arrangement may include: a plate that comprises a plate topside; a channel formed on the plate topside, the channel comprising an upstream end and a downstream end, and being open through the plate topside such that, upon attaching the plate to the platform, the platform underside comprises a channel ceiling; a high-pressure connector that connects the upstream end of the channel to the high-pressure coolant region of the interior cooling passage; and a low-pressure connector that connects the downstream end of the channel to the low-pressure coolant region of the interior cooling passage.

Abstract: Turbine blade airfoils, showerhead film cooling systems thereof, and methods for cooling the turbine blade airfoils using the same are provided. The airfoil has a leading edge and a trailing edge, a pressure sidewall and a suction sidewall both extending between the leading and the trailing edges, and an internal cavity for supplying cooling air. A showerhead of film cooling holes is connected to the internal cavity. Each film cooling hole has an inlet connected to the internal cavity and an outlet opening onto an external wall surface at the leading edge of the airfoil. A plurality of surface connectors is formed in the external wall surface. Each surface connector of the plurality of surface connectors interconnects the outlets of at least one selected pair of the film cooling holes.

Abstract: The present disclosure is directed to the use and manufacture of cooling features within a component used in a hot gas path, such as within a turbine. In one embodiment, channels are formed within an external surface of the component and filled with a removable material. The external surface and channels may then be coated with one or more layers, such as a structural layer and/or top coat. The removable material may then be removed to leave the channels free of the removable material.

Abstract: A soluble and substantially ring-shaped casting core may include a cooling channel in a piston having at least one inflow and at least one outflow. The casting core may form a closed ring and the inflow may be arranged directly adjacent to the outflow. The casting core may have a cross-sectional taper for forming a choke in the circumferential direction between the inflow and the outflow.

Abstract: Methods for testing turbine blades. One method for testing turbine blades includes measuring dimensions of each of a first set of turbine blades. The method also includes testing airflow through first openings in each of the first set of turbine blades to determine airflow properties of each of the first set of turbine blades. The method includes determining a relationship between the dimensions and the airflow properties of each of the first set of turbine blades. The method includes measuring dimensions of each of a second set of turbine blades. The method also includes determining airflow properties for each of the second set of turbine blades based at least partially on the dimensions of the second set of turbine blades and the relationship.

Abstract: There are a multiplicity of methods of making through-holes. In particular in the production of a multiplicity of film-cooling holes, as in gas turbine blades or combustion chamber elements, small time advantages are also important when making a hole. The method according to the invention, to make the hole close to the final contour in each case in sections in a top and a bottom region in order to then produce the final contour with other laser parameters, achieves time advantages.

Abstract: A method of manufacturing a cooled turbine blade for use in a gas turbine engine. The method includes forming an inner blade pattern, the inner blade pattern including an inner spar and a plurality of inner spar cooling fins. The method also includes forming an inner blade core, removing the inner blade pattern from the inner blade core, forming an outer blade pattern, forming a casting shell, removing the outer blade pattern from the casting shell, and casting the cooled turbine blade in the casting shell. The method also includes removing the casting shell from the cast cooled turbine blade, and removing the inner blade core from the cast cooled turbine blade.

Abstract: A platform cooling configuration in a turbine rotor blade that includes platform slot formed through at least one of the pressure side slashface and the suction side slashface; a removably-engaged impingement insert that separates the platform into two radially stacked plenums, a first plenum that resides inboard of a second plenum; a high-pressure connector that connects the first plenum to the high-pressure coolant region of the interior cooling passage; a low-pressure connector that connects the second plenum to the low-pressure coolant region of the interior cooling passage.

Abstract: A system for manufacturing an airfoil includes an outer surface of the airfoil, a cavity inside the airfoil, and a collimator outside of the airfoil. The system further includes a fluid column flowing from the collimator toward the outer surface of the airfoil, and a laser beam inside the fluid column creates a confined laser beam directed at the outer surface of the airfoil. A method for manufacturing an airfoil includes confining a laser beam inside a fluid column to create a confined laser beam, directing the confined laser beam at an outer surface of the airfoil, and creating a passage through the outer surface of the airfoil with the confined laser beam.

Abstract: A component is provided and comprises at least one wall comprising a first and a second surface. At least one film cooling hole extends through the wall between the first and second surfaces and has an exit region at the second surface. The second surface has a non-planar curvature in the vicinity of the exit region. The film cooling hole is tapered at the exit region, such that the curvature of the film cooling hole in the exit region conforms to the non-planar curvature of the second surface, thereby forming a curved exit region. A method is also provided for forming at least one film cooling hole in the component. The method includes forming a straight section in the component wall, such that the straight section extends through the first surface, and tapering the film cooling hole, such that the curvature of the film cooling hole in the exit region conforms to the non-planar curvature of the second surface, thereby forming a curved exit region for the film cooling hole.

Abstract: An airfoil for a gas turbine engine includes pressure and suction surfaces provided by pressure and suction walls that extend in a radial direction and are joined at a leading edge and a trailing edge. A cooling passage is arranged between the pressure and suction walls and extends to the trailing edge. Elongated pedestals are arranged in the cooling passage and interconnect the pressure and suction walls. The elongated pedestals are spaced apart from one another in the radial direction and extend from a plane to the trailing edge. A metering pedestal includes at least a portion that is arranged between the plane and the trailing edge. The portion is provided between adjacent elongated pedestals in the radial direction.

Abstract: A configuration of cooling channels through the interior of a turbine rotor blade, the turbine rotor blade including a platform at an interface between an airfoil and a root. In one embodiment, the configuration of cooling channels includes: an interior cooling passage that is configured to extend from a connection with a coolant source in the root to the interior of the airfoil; a platform cooling channel that traverses at least a portion of the platform; a turndown extension that includes a first section, which comprises a connection with the platform cooling channel, and a second section, which comprises a radially oriented cooling channel; and a connector that extends from a connector opening formed through an outer face of the root to a connection with the interior cooling passage and, therebetween, bisects the second section of the turndown extension.

Abstract: The invention introduces a reinforcement of a box girder of a wind turbine blade. The reinforcement prevents the transverse shear distortion of the blade structure, when the blade is loaded during operation. The reinforcement connects the corners diagonally opposite inside the girder, and fixes them in relation to each other. The reinforcement increases the blade's resistance to overall collapse. The reinforcement comprises one or more individual element, such as rods or plates.

Abstract: A configuration of cooling channels through the interior of a turbine rotor blade having a platform, wherein the rotor blade includes an airfoil cooling channel that includes a cooling channel formed within the airfoil and an outboard airfoil supply channel. The configuration of cooling channels may include: a platform cooling channel that comprises a cooling channel that traverses at least a portion of the platform, the platform cooling channel having an upstream end and a downstream end; an outboard platform supply channel, which comprises a cooling channel that stretches from a second coolant inlet formed in the root to the upstream end of the platform cooling channel; and an inboard platform return channel, which comprises a cooling channel that stretches from the downstream end of the platform cooling channel to a termination point formed in the root.

Abstract: A method of manufacturing a metal blading sector for low-pressure guide vanes of a turbomachine of which at least one blade includes an internal cavity configured to accept or communicate with a gas detection probe and at least one hole formed in the wall constituting a passage for gas from a low-pressure zone of the turbomachine toward the cavity and the probe through the fitting, into a casting mold, of a core corresponding to the cavity and the casting of a molten metal in the cavity of the casting mold. The core includes, for each hole for communication with the cavity, a protrusion penetrating the internal surface of the mold and constituting the only element holding the core in position in the casting mold.

Abstract: Methods of fabricating coated components using multiple types of fillers are provided. One method comprises forming one or more grooves in an outer surface of a substrate. Each groove has a base and extends at least partially along the outer surface of the substrate. The method further includes disposing a sacrificial filler within the groove(s), disposing a permanent filler over the sacrificial filler, disposing a coating over at least a portion of the substrate and over the permanent filler, and removing the first sacrificial filler from the groove(s), to define one or more channels for cooling the component. A component with a permanent filler is also provided.

Abstract: An airfoil includes a body that includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface. A cooling passage is arranged interiorly of the exterior airfoil surface and provides an interior surface. The interior cooling surface includes micro-bumps that protrude from the interior cooling surface into the cooling passage. The micro-bumps are discrete from and noncontiguous relative to one another in multiple directions along the interior cooling surface. The micro-bumps may be provided while forming the airfoil or using correspondingly shaped micro-depressions on an airfoil core.

Abstract: A turbine rotor blade formed from three pieces with a pressure wall side piece and a suction wall side piece bonded to an intermediate piece so that a pressure side cooling circuit can be formed as a separate cooling circuit from a suction side cooling circuit. The intermediate piece has film cooling holes and blade tip cooling holes and trailing edge exit slots formed in it that are enclosed when the outer two pieces are boded to it.

Abstract: A method of manufacturing a component is provided. The method includes forming one or more grooves in an outer surface of a substrate. Each groove extends at least partially along the surface of the substrate and has a base, a top and at least one discharge point. The method further includes forming a run-out region adjacent to the discharge point for each groove and disposing a coating over at least a portion of the surface of the substrate. The groove(s) and the coating define one or more channels for cooling the component. Components with cooling channels are also provided.

Abstract: Methods of fabricating coated components using multiple types of fillers are provided. One method comprises forming one or more grooves in an outer surface of a substrate. Each groove has a base and extends at least partially along the outer surface. A sacrificial filler is deposited within the groove, a second filler is deposited over the sacrificial filler, and a coating is disposed over at least a portion of the outer surface and over the second filler. The method further includes removing the sacrificial filler and at least partially removing the second filler from the groove(s), to define one or more channels for cooling the component.

Abstract: A thin wall turbine blade used in a gas turbine engine, in which the blade is cast in conventional grain from a super alloy using the lost wax process as a single piece, and then the blade walls are machined to remove enough material to leave a thin wall. The blade is cast with a wall thickness greater than the designed for thin wall in order that any core shifting during the casting process will be accounted for in the machining process. prior to machining, a scanning process is used to measure the actual wall thickness on all portions of the blade wall in order to determine how much material must be removed to leave the blade wall with the proper thinness.

Abstract: A method of manufacturing an aerofoil structure capable of being diffusion bonded and superplastically formed to create a substantially hollow cavity within the aerofoil structure, the method comprising: providing a metallic plate for forming the aerofoil structure; joining mounting elements to opposing end surfaces of said metallic plate; dividing said plate along a plane extending substantially in a span-wise direction so as to produce two metallic panels each with one of said mounting elements joined thereto; assembling the two metallic panels so that the surfaces of the panels opposite to the surfaces which have been divided are facing each other; and joining the two metallic panels to one another to form the aerofoil structure; wherein the mounting elements are joined to one another to form the root of the aerofoil.

Abstract: Fabricating a turbine component (50) by casting a core structure (30), forming an array of pits (24) in an outer surface (32) of the core structure, depositing a transient liquid phase (TLP) material (40) on the outer surface of the core structure, the TLP containing a melting-point depressant, depositing a skin (42) on the outer surface of the core structure over the TLP material, and heating the assembly, thus forming both a diffusion bond and a mechanical interlock between the skin and the core structure. The heating diffuses the melting-point depressant away from the interface. Subsurface cooling channels (35) may be formed by forming grooves (34) in the outer surface of the core structure, filling the grooves with a fugitive filler (36), depositing and bonding the skin (42), then removing the fugitive material.

Abstract: An airfoil for a gas turbine engine is provided that includes a first sidewall and a second sidewall coupled together at a leading edge and a trailing edge, such that a cavity is defined therebetween. A central plenum and an impingement chamber are defined within the cavity. The central plenum channels cooling fluid to the impingement chamber where cooling fluid impinges on the sidewalls. Cooling fluid is discharged from the impingement chamber via film cooling holes.

Abstract: An airfoil includes a leading edge, a trailing edge, and a pair of sides extending from the leading edge to the trailing edge. The airfoil also includes an internal cooling flow passage defined between the sides, wherein the passage has a passage axis along which cooling air is to flow. The airfoil further includes a plurality of flow paths extending through at least one of the sides such that the flow paths are configured to discharge cooling air from the passage, wherein each of the flow paths has a broken flow path axis oriented to intersect the passage axis at an acute angle.

Abstract: A cooling system for use in regulating a temperature of a component is described herein. The cooling system includes a plurality of flow control assemblies that are defined across a sidewall of the component for channeling a cooling fluid across a surface of the sidewall. The plurality of flow control assemblies are configured to adjust a plurality of fluid flow characteristics of the cooling fluid. The plurality of flow control assemblies includes a first flow control assembly configured to adjust a first fluid flow characteristic, and at least a second flow control assembly configured to adjust a second fluid flow characteristic that is different than the first fluid flow characteristic.

Abstract: An airfoil for a gas turbine engine includes a first sidewall and a second sidewall coupled together at a leading edge and a trailing edge, such that a cavity is defined therebetween. A plurality of cooling circuits are defined within the cavity. Each cooling circuit channels cooling fluid through at least one cooling chamber to facilitate cooling the airfoil. More specifically, a cascade impingement circuit, a down pass circuit, a flag tip circuit, and a trailing edge circuit are provided. The cascade impingement circuit includes a central chamber and a plurality of impingement chambers.

Abstract: The present application thus provides an airfoil for use in a turbine. The airfoil may include a wall, an internal cooling plenum, and a cooling hole extending through the wall to the cooling plenum. The cooling hole may include an offset counterbore therein.

Abstract: A turbine blade is provided and includes a tip end carrying a shroud and at least one fin, which extends radially away from the shroud. The fin includes a first sidewall and a second sidewall, which are spaced apart, arranged parallel to each other, and are connected to the shroud, and a cutting edge, which is connected to the first and second sidewalls. The cutting edge thereby creates a hollow space between the sidewalls, the shroud, and the cutting edge, and further extends radially away from the first and second sidewalls. Also provided is a method of manufacturing the blade by casting the blade as single piece with the hollow fin or by forging the blade; and machining the fin to create the first and second sidewalls and cutting edge thereby opening the hollow space between said sidewalls and the cutting edge.

Abstract: A method of fabricating a component is provided. The method includes forming one or more grooves in a surface of a substrate, where the substrate has at least one hollow interior space. Each of the one or more grooves extends at least partially along the substrate surface and has a base and a top. The base is wider than the top, such that each of the one or more grooves comprises a re-entrant shaped groove. The method further includes forming one or more access holes through the base of a respective groove, to connect the groove in fluid communication with respective ones of the hollow interior space(s), and disposing a coating over at least a portion of the substrate surface. The one or more grooves and coating define one or more re-entrant shaped channels for cooling the component. A component with one or more re-entrant shaped channels and a method of coating a component are also provided.

Abstract: A turbine rotor blade made from the spar and shell construction in which the shell is a thin wall shell made from a high temperature resistant material that is formed by a wire EDM process, and where the shell is secured to the spar using a retainer that is poured into retainer occupying spaces formed in the shell and the spar, and then hardened to form a rigid retainer to secure the shell to the spar. The spar and the shell both have grooves facing each other to form a retainer groove. A retaining material, such as a liquid or a powdered metal, is poured into the grooves and hardened to form a retainer to secure the shell to the spar. The retaining material also forms a seal on the top of the spar and between the spar and shell.

Abstract: A blade (10) for a gas turbine has a blade airfoil (11), the blade wall (18) of which encloses an interior space (17). For cooling the blade wall (18), the blade wall (18) includes a cooling arrangement (19) which has a radial passage (20) extending in the longitudinal direction of the blade and from which a multiplicity of cooling passages (21, 22), extending in the blade wall (18), branch in the transverse direction, and from which a multiplicity of film-cooling holes (23) are led to the outside in the transverse direction. Particularly efficient cooling is made possible by the distribution of the film-cooling holes (23) along the radial passage (20) being selected independently of the distribution of the cooling passages (21, 22) along the radial passage (20).

Abstract: A method for creating a platform cooling passage in a turbine rotor blade, wherein the turbine rotor blade comprises a platform at an interface between an airfoil and a root, wherein the platform includes a platform topside along an outboard surface. The method may include the steps of forming a recessed area along the platform topside; forming a coverplate; and affixing the coverplate to the platform topside. The coverplate may be configured to correspond to the shape of the recessed area such that, when affixed to the platform topside in a desired manner, the coverplate substantially encloses the recessed area to form the platform cooling passage therein.

Abstract: A turbine component may generally comprise an airfoil having a base and a tip disposed opposite the base. The airfoil may further include a pressure side surface and a suction side surface extending between a leading edge and a trailing edge. An airfoil circuit may be at least partially disposed within the airfoil and may be configured to supply a medium through the airfoil. The turbine component may also include a curved passage defined in the airfoil so as to be in flow communication with the airfoil circuit. Additionally, an outlet may be defined through the pressure side surface or the suction side surface of the airfoil. The outlet may be in flow communication with the curved passage and may have a cross-sectional area that is greater than a cross-sectional area of the curved passage.

Abstract: A thin wall turbine blade used in a gas turbine engine, in which the blade is cast in conventional grain from a super alloy using the lost wax process as a single piece, and then the blade walls are machined to remove enough material to leave a thin wall. The blade is cast with a wall thickness greater than the designed for thin wall in order that any core shifting during the casting process will be accounted for in the machining process. prior to machining, a scanning process is used to measure the actual wall thickness on all portions of the blade wall in order to determine how much material must be removed to leave the blade wall with the proper thinness.

Abstract: A cooling system comprises serpentine cooling fluid passages cast into a work piece with carefully controlled turning radii and profiles. Individual interdigitated baffles are contoured in the plane of coolant flow to have walls that thicken and then round off at their distal ends. The outside radii at these turns is similarly rounded and controlled such that the coolant flow will not be swirled into eddies.

Abstract: A turbine blade is provided. The turbine blade includes a support structure and a shell which surrounds the support structure and which is connected to and at a distance from the support structure by at least one spacing element. For example the spacing element can be a solder globule in order to form a space through which a cooling medium can flow between the support structure and the shell. A method for the production of a turbine blade having a support structure and a shell which surrounds the support structure and which is connected to and at a distance from the support structure is also provided. The shell is soldered to the support structure in at least at one place of the support structure in order to connect the shell to and at a distance from the support structure.

Abstract: A configuration of cooling channels through the interior of a turbine rotor blade having a platform, wherein the rotor blade includes an airfoil cooling channel that includes a cooling channel formed within the airfoil and an outboard airfoil supply channel. The configuration of cooling channels may include: a platform cooling channel that comprises a cooling channel that traverses at least a portion of the platform, the platform cooling channel having an upstream end and a downstream end; an outboard platform supply channel, which comprises a cooling channel that stretches from a second coolant inlet formed in the root to the upstream end of the platform cooling channel; and an inboard platform return channel, which comprises a cooling channel that stretches from the downstream end of the platform cooling channel to a termination point formed in the root.

Abstract: A turbine blade of an axial turbine includes internal cooling fluid passages with radially outwardly extending passages connected to holes in the blade root. The holes are generally core printouts providing stability to the core during the casting process, but are not needed and need to be closed to guarantee the functioning of the cooling system. This is achieved by at least one covering plate. The plate is held by at least two slots located at the root of the turbine blade. Thus, the supply holes for cooling fluid located at the root section are closed by a simple mechanical device, e.g., a plate that does not require any subsequent brazing/welding operations. In addition, the plate is removable to facilitate inspection/cleaning, or further processing of the blade at service intervals.

Abstract: A turbine engine component has an airfoil portion, which airfoil portion is bounded by a platform at one end. The platform has an as-cast open cavity bordered by at least one as-cast landing. A plate is welded to the at least one as-cast landing to cover and close the as-cast open cavity. A process for forming the turbine engine component is described.

Abstract: A configuration of cooling channels through the interior of a turbine rotor blade, the turbine rotor blade including a platform at an interface between an airfoil and a root. In one embodiment, the configuration of cooling channels includes: an interior cooling passage that is configured to extend from a connection with a coolant source in the root to the interior of the airfoil; a platform cooling channel that traverses at least a portion of the platform; a turndown extension that includes a first section, which comprises a connection with the platform cooling channel, and a second section, which comprises a radially oriented cooling channel; and a connector that extends from a connector opening formed through an outer face of the root to a connection with the interior cooling passage and, therebetween, bisects the second section of the turndown extension.

Abstract: A method is provided for milling a thermal barrier coated metal part. This method includes selectively removing a portion of a ceramic coating using a mechanical cutting tool, thereby forming a counterbore, and machining the metal part through the counterbore. A drilling head for drilling thermal barrier coated metal parts is also provided. The drilling head comprises a mechanical cutting tool, which is operable to mill through ceramic, and an electrode for electrical discharge machining. The electrode may be used to mill the metal part, and may be interchangeable with the mechanical cutting tool.

Abstract: A platform cooling arrangement in a turbine rotor blade having a platform at an interface between an airfoil and a root, wherein the rotor blade includes an interior cooling passage that extends to the approximate radial height of the platform, and wherein, a pressure side of the platform comprises a planar topside that extends circumferentially from the airfoil to a pressure side slashface, and a suction side of the platform comprises a substantially planar topside that extends circumferentially from the airfoil to a suction side slashface.

Abstract: A platform cooling arrangement in a turbine rotor blade having a platform at an interface between an airfoil and a root, wherein the rotor blade includes an interior cooling passage that extends from a connection at the root to the approximate radial height of the platform, wherein, the interior cooling passage comprises a high-pressure coolant region and a low-pressure coolant region, and wherein a pressure side of the platform comprises a topside extending circumferentially from the airfoil to a pressure side slashface. The platform cooling arrangement may include: a platform cavity formed within the pressure side of the platform, a high-pressure connector that connects the platform cavity to the high-pressure coolant region of the interior cooling passage; a low-pressure connector that connects the platform cavity to the low-pressure coolant region of the interior cooling passage; and a pin bank formed within the platform cavity that includes radial pins.

Abstract: An airfoil assembly for a turbine engine, that may comprise: an airfoil that includes at least one panel opening formed therein; and an aperture panel that includes at least one cooling aperture; wherein the panel opening and the aperture panel are configured such that, upon installation of the aperture panel within the panel opening, a substantially smooth outer surface of the airfoil is formed and at least one of the cooling apertures of the aperture panel forms an exit passageway from at least one hollow internal cavity within the airfoil.