Abstract: A gas turbine rotating blade comprises a serpentine passage provided in the blade width direction in plural rows in a blade profile portion except the portion near the trailing edge having a small blade thickness, a steam cooling passage provided around the outer periphery of a platform, an air passage provided in the blade width direction in the vicinity of the trailing edge of blade profile portion, an impingement plate provided under the platform on the inside of the steam cooling passage, slot holes formed so as to branch off from the air passage and be directed to the trailing edge, and slits formed in the platform above the impingement plate. Therefore, the blade profile portion is cooled by the steam passing through the serpentine passage and the air passing through the air passage, so that the cooling construction is not complicated, and cooling is performed effectively. For the platform, the outer periphery thereof is cooled by steam and the inside thereof by air effectively.

Abstract: In a steam cooled type gas turbine, taking the effect of efficiency and power output into consideration, a rear stage is sufficiently cooled by air in place of cooling all of the stages by steam. In a rotor of the gas turbine having a steam cooled moving blade, a moving blade disposed in a front stage is cooled by steam, and a moving blade disposed at the rear stage is cooled by cooling air through a stationary blade disposed in front of the moving blade at the rear stage.

Abstract: A cooling system is disclosed for cooling the platforms of turbine blades fixed to the periphery of a turbine rotor, each turbine blade having an airfoil portion, a platform with opposite longitudinal edges, a root portion affixed to the turbine rotor, and a shank portion connecting the root portion to the platform. Adjacent turbine blades are located such that a longitudinal edge of one platform is adjacent to a longitudinal edge of an adjacent platform. A cavity is formed bounded by a portion of the periphery of the turbine rotor, the shank portions of the adjacent turbine blades, and the platforms of the adjacent turbine blades. A cooling air passage supplies cooling air to the cavity. One or more cooling channels are formed in a side of the platform facing the turbine rotor, each cooling channel having a cooling air collector at an end located adjacent to the air supply passage and communicating with the cavity.

Abstract: A hollow airfoil is provided having a pressure side wall, a suction side wall, a cavity formed between the pressure and suction side walls, a plurality of cooling ports disposed within the pressure side wall, and a plurality of passages, each extending between the cavity and one of the cooling ports. Each passage has a first wall adjacent the suction side wall, a pair of passage side walls extending substantially toward the pressure side wall, and a second wall adjacent the pressure side wall. In a first embodiment, each passage further includes a pair of fillets extending between the passage side walls and the second wall. In a second embodiment, each passage includes a jog adjacent each cooling port.

Abstract: In the invention, propagating broad band and tonal acoustic components of noise characteristic of interaction of a turbomachine blade wake, produced by a turbomachine blade as the blade rotates, with a turbomachine component downstream of the rotating blade, are reduced. This is accomplished by injection of fluid into the blade wake through a port in the rotor blade. The mass flow rate of the fluid injected into the blade wake is selected to reduce the momentum deficit of the wake to correspondingly increase the time-mean velocity of the wake and decrease the turbulent velocity fluctuations of the wake. With this fluid injection, reduction of both propagating broad band and tonal acoustic components of noise produced by interaction of the blade wake with a turbomachine component downstream of the rotating blade is achieved.

Abstract: A coolable double walled structure, for use and exposure in a hot gas flow environment, includes a jet issuing wall, a target wall spaced from the jet issuing wall defining at least one cavity therebetween. The jet issuing wall includes at least one impingement rail having a greater thickness than the jet issuing wall such that a bottom portion of the impingement rail extends within the cavity formed between the jet issuing wall and the target wall. A number of impingement ports are disposed within each impingement rail to provide flow communication through the jet issuing wall to the target wall. Due to the relatively thick dimensions of the impingement rail, the impingement ports disposed within respective impingement rails can be directionalized so as to provide for local cooling needs. In one embodiment, the entry areas of the impingement ports are bell shaped so as to lower inlet pressure loss.

Abstract: The invention relates to a cooled shroud in a gas turbine stationary blade which is able to flow a cooling air in the entire area of an inner shroud for cooling thereof. Three stationary blades are fixed to the inner shroud 2, a cover 13, 14 is provided to form a space 21 and space 22a, 22b and 22c, respectively. The cooling air is introduced through an independent air passage 3A of a leading edge of each stationary blade into the spaces 22a, 22b and 22c and is flown therefrom through a tunnel 18 and air reservoirs 19-2, 19-3 and 19-4 to be blown out of a trailing edge while cooling surfaces of the shrouds and the trailing edges. Also, a portion of the cooling air from the space 22b is flown into the space 21 through a tunnel 11, a leading edge side passage 12 and an endmost tunnel 11 and is then blown out of the trailing edge through a tunnel 18 and an air reservoir 19-1, so that the leading edge portion, the endmost portion and the endmost trailing edge portion are cooled.

Abstract: A gas turbine engine turbine blade squealer tip includes an airfoil tip cap having an airfoil shape and pressure and suction sides joined together at chordally spaced apart leading and trailing edges of the tip cap. The squealer tip further includes a plurality of squealer ribs extend outward from the tip cap between the leading and trailing edges and substantially parallel channels between the squealer ribs. The squealer ribs and channels may have rectangular cross-sections and may also include a trailing edge rib along a portion of the suction side including the trailing edge and/or a leading edge rib curved around an axially extending portion of the leading edge. The squealer ribs are preferably angled in an axially aft direction from the leading edge on the suction side towards the trailing edge on the pressure side. Channel ribs may be disposed between the squealer ribs and across the channels and may be arranged such that only one of the channel ribs is disposed across each of the channels.

Abstract: A turbine vane-system cooling system uses three internal cooling cavities 1, 12, 13) separated by two radial walls (9, 10). The upstream cavity (11) uses a helical ramp (30) and is fed through an intake (22) at the vane root (3). The middle cavity (12) also is fed at the vane root (3) and includes a compartmented, multi-perforated lining (40). The air is exhausted from each compartment through impact orifices and enters the succeeding compartment through slots (42) and then is finally exhausted through a vane-head orifice (21). The vane side walls opposite the downstream cavity (13) have double skins with bridging elements. The air passes through these double skins but circulates centrifugally in the upstream portion (15) of the downstream cavity (13) and enters this cavity's downstream portion (16) to be exhausted through slots (19) in the trailing edge (6). A third wall (14) divides the downstream cavity (13) into two parts (15, 16).

Abstract: A cooling medium path structure for cooling a gas turbine blade. The structure includes a disk-side cooling medium path, a blade-side cooling medium path formed at the root portion of the blade and a delivery block disposed between the two cooling medium paths so as to establish communication therebetween. The delivery block is provided with an elastic engaging section which comes into elastic and line-contact with the disk-side cooling medium path and the blade-side cooling medium path. Thus, the sealing property of the contact portions of the structure is secured so as to allow a cooling medium to be supplied without leaking from the cooling medium paths. The heat of the cooling medium, generated as a result of cooling the high-temperature portion of the gas turbine, can be recovered so as to make the best use of the heated medium.

Abstract: A gas turbine stationary blade having a cooling medium passage and outer and inner shrouds in a blade section, comprising a hollow portion in the blade section, the hollow portion including a plurality of chambers divided in the chord direction, the cooling medium passage being formed by connecting the divided chambers in series, one of the chambers located in the center with respect to the chord direction being formed as a sealing air passage in order to allow a cooling medium to flow from the outside of the outer shroud to the bore of the inner shroud.

Abstract: A first-stage turbine blade includes an airfoil having a profile according to Table I. The airfoil has a plurality of cooling air passages extending linearly from the root portion to the tip portion of the airfoil. The blade includes a shank having a pair of cavities in communication through the blade dovetail with a plenum in the wheel space for supplying cooling air to the passages in the airfoil. The cooling passages in the airfoil terminate in a recess at the tip portion which has an opening adjacent the trailing edge of the airfoil and along the suction side to enable egress of cooling air into the hot gas stream on the low pressure side of the airfoil. The majority of the cooling passages are turbulated. Certain of those passages are arranged in rows lying adjacent to the pressure and suction sides of the airfoil.

Abstract: A convectively cooled turbine blade has two distinct cooling air passage systems. The first system cools the blade leading edge and emits cooling air through outlet passageways in the leading edge arranged in showerhead array. The second system includes a five-pass series flow passage comprising five cooling passage sections that extend in series through the remainder of the blade. One of the passage sections includes a plurality of recesses near the trailing edge of the turbine blade to retain cooling air flow to the trailing edge adjacent the root portion of the blade.

Abstract: A turbine blade including an airfoil section having contoured surface geometries between the consecutive ribs or turbulators on the leading edge passage walls is described. The contoured surface geometries increase the overall heat transfer surface area between adjacent ribs as compared to the heat transfer surface area of a smooth inter-rib wall. The contoured surface geometries do not, however, change the overall serpentine passage geometry or the large scale flow characteristics of the turbulated cooling passageways. The contoured surface geometries may have many different geometries (e.g., triangular, conical, semi-cylindrical, cylindrical columns or indentations in the side wall such as semi-circular dimples).

Abstract: A turbine airfoil includes a plurality of internal ribs joined to laterally opposite first and second sidewalls to define a pair of serpentine cooling circuits therein. A span rib extends longitudinally between a root and tip of the airfoil. And, slant ribs extend outwardly from opposite sides of the span rib to define respective portions of the serpentine circuits.

Abstract: A product and a method for manufacturing the product include a metallic basic body with at least one longitudinal duct disposed therein and with a number of transverse ducts branching off from the longitudinal duct. A covering layer lies on the outside of the basic body. The covering layer serves as a protective or adhesive layer. An enrichment layer covers walls of the longitudinal duct and of the transverse ducts and parts of the covering layer. A ceramic heat-insulating layer can also be provided on the outside.

Abstract: A turbine airfoil includes a plurality of internal ribs joined to laterally opposite first and second sidewalls to define a serpentine cooling circuit therein. A span rib extends longitudinally between a root and tip of the airfoil. And, a slant rib extends obliquely to the span rib to define respective portions of the serpentine circuit longitudinally above and below the slant rib.

Abstract: A rotor disk for a gas turbine engine having an axis of rotation and an axial centerline is provided which includes a rim, forward and aft hubs, and forward and aft webs. The forward web extends between the rim and the forward hub, and the aft web extends between the rim and the aft hub. The forward and aft hubs are separated. Each web is skewed from the axial centerline. A circumferentially extending cavity is formed between the webs and the hubs.

Abstract: In connection with a turbine rotating disk with disk grooves, constituted by disk fingers, for receiving turbine blades, as well as with measures for guiding a cooling air flow from a chamber located in front of the disk to one behind the disk, a cooling air transfer channel is provided in at least some of the disk fingers, which respectively starts at the front disk face, and makes a transition into a cooling air exhaust channel, also essentially extending in the radial direction in the disk finger, whose outlet opening on the rear disk face side lies closer to the disk axis than the disk ring section, which has the disk grooves and is widened in the disk axial directions. By means of this it is possible to convey a larger cooling air flow which, in an advantageous manner, indirectly aids the cooling of the disk in the area of the disk grooves.

Abstract: A gas turbine stationary blade, having a simple structure in which sufficient cooling is achieved and the drop in pressure of cooling vapor is decreased so that the turbine efficiency is prevented from lowering. The shape of a vapor passage is simplified to prevent the drop in pressure because an outer shroud (3) of the stationary blade and a blade unit (2) are cooled with vapor, while an inner shroud (4) is cooled with the air supplied from another system.

Abstract: The internally air cooled turbine cast blade is molded with a ceramic core for defining the interior cavity with coolant passages where the ceramic core is configured to fair radially inwardly at the attachment portion of the blade until they form a single wall, parallel to and centered in the blade attachment. The blade is cast in the mold by pouring molten metal into the mold, curing and the core is leached once the metal is cured and the mold removed. Another embodiment includes a metering plate attached to the root of the blade for metering coolant to provide a coolant to gas flow ratio at the film cooling holes.

Abstract: A turbine cooling apparatus comprising a turbine disk having a plurality of moving blades, a torque tube coupled coaxially to one surface side of the turbine disk and having a thick stepped inner wall portion in the central portion thereof, an air separator fitted on the torque tube with the inner surface thereof in contact with the outer surface of the torque tube so that a passage through which cooling air is supplied to the moving blades via the turbine disk is defined between the air separator and the torque tube, and a torque tube cooling hollow portion provided along and in the vicinity of the outer surface of the thick stepped wall portion of the torque tube.

Abstract: When cooling air from a gas turbine engine compressor is supplied in greater quantities than is needed, the flow thereof is reduced firstly and mainly by ensuring that the air leaves a blade stage cooling thereby at a radius less than that of its entry. A pressure head results which slows the cooling air flow rate prior to its entering a plenum chamber defined by the disc of the stage and fixed structure. There is a reduced tendency of the air to enter the gas stream.

Abstract: A coolable blade (10) essentially comprises a blade root (11) and a blade body (1), which is composed of a pressure-side wall (6) and a suction-side wall (5). They are connected to one another essentially via a trailing-edge region (4) and a leading-edge region (3) in such a way that at least one hollow space (2) used as a cooling-fluid passage is formed. An essentially radially running, diverging cooling passage (7) is arranged in the trailing-edge region (4).

Abstract: A blade or vane for a gas turbine engine includes a primary cooling system (42) with a series of medial passages (44, 46a, 46b, 46c, 48) and an auxiliary cooling system (92) with a series of cooling conduits (94). The conduits of the auxiliary cooling system are parallel to and radially coextensive with the medial passages and are disposed in the peripheral wall (16) of the airfoil between the medial passages and the airfoil external surface (28). The conduits are chordwisely situated in a zone of high heat load (104, 106) so that their effectiveness is optimized. The conduits may also be chordwisely coextensive with some of the medial passages so that coolant in the medial passages is protected from excessive temperature rise. The chordwise dimension C of the conduits is limited so that potentially damaging temperature gradients do not develop in the airfoil wall (16).

Abstract: A gas turbine engine rotor blade includes an airfoil having a concave side wall and a convex side wall joined together at leading and trailing edges. Concave and convex tip walls extend from adjacent the leading edge along the respective concave and convex side walls to adjacent the trailing edge and are spaced apart to define a tip cavity therebetween. A hole or channel is disposed in a trailing edge tip region connecting the tip cavity to the trailing edge for channeling cooling fluid through the trailing edge tip region.

Abstract: A gas turbine moving blade in which the moving blade operating in a high temperature operating gas is cooled from its interior by steam and the steam after being used for the cooling is recovered, by use of a simple structure. A supply side passage (10), through which steam (6) is supplied directly into a cooling passage (51) of a blade leading edge side (11), is provided within a blade root portion (2) and a pocket (19). One end of the pocket is connected to the supply side passage and the other end of which is connected to a cooling passage (52) of a blade trailing edge side (12). The pocket is covered with a plate, which extends in a blade chord direction on an outer peripheral side face of the blade root portion between a blade platform and a rotor plate. Thus, efficiency of the gas turbine can be effectively enhanced.

Abstract: Internal hot corrosion which occurs in hollow uncooled gas turbine engine components is reduced by increasing the oxygen potential of the gas which fills the hollow portion of the component. In particular, blades and vanes which operate in conditions where there is a pressure differential across the component have deliberate venting holes introduced in a sealing structure at one end of the component to permit gas flow through the hollow portion of the component.

Abstract: A rotor blade includes a dovetail and an airfoil joined thereto. The airfoil includes first and second spaced apart sides joined together laterally at opposite leading and trailing edges, and spanwise at a root and opposite tip. A serpentine cooling circuit extends inside the airfoil for channeling air therethrough for cooling the blade. The serpentine circuit includes first and second passes and a first bend therebetween for firstly receiving the cooling air in turn from the dovetail. A tip circuit is disposed between the tip and the serpentine circuit at the first bend for separating the tip from the first bend and providing cooling thereof near the trailing edge.

Abstract: A gas turbine blade having a plurality of shower head cooling holes (1), bored in a blade leading edge portion, are arranged along a single line which is biased so as to correspond to the direction of swirling flow in a combustor. The swirling flow is clockwise or counter-clockwise. In particular, if the flow is clockwise, the cooling holes (1) are biased to the left hand side, as seen from upstream of the blade, toward a blade hub side (HS) from a blade tip side (TS). Also, if the flow is counter-clockwise, the cooling holes (1) are arrayed on one line which is biased to the reverse side relative to the latter case. The influence of the swirling flow is thus avoided, and the pressure at each position of the cooling holes (1) is constant, so that uniform cooling can be attained.

Abstract: A steam cooled blade operating in a high temperature gas with a high cooling effect on a trailing edge portion has an impingement plate, disposed in the blade lengthwise direction within a trailing edge side cooling passage formed on a trailing edge side of a moving blade. The impingement plate partitions the trailing edge side cooling passage into a convection cooling steam passage, into which steam is introduced, and an impingement cooling steam passage, formed along a trailing edge of the moving blade. The steam cooled blade further has a by-pass passage 14 for joining the steam, after it is used for cooling the moving blade, flowing from a blade tip side of the impingement cooling steam passage with the steam flowing to a leading edge direction of the moving blade from the convection cooling steam passage.

Abstract: A coolable airfoil (25) for a gas turbine engine includes a leading edge portion (52) cooled by a combination of radial and impingement flows through internal passages (80) and (90) separated by an internal wall (95) which is imperforate along a radially inner portion (100) thereof and apertured along a radially outer portion (105) thereof

Abstract: An effective cooling structure that is intended to cool a platform of a gas turbine moving blade. Cooling air passages (B1, C1, D1; B2, C2, D2; and B3, C3 and D3) are provided through the interior and along the peripheral edge of the platform (2) to thereby cause a cooling air to pass therethrough.

Abstract: An airfoil having a trailing edge cavity formed by a leading wall and a trailing edge connected by a pair of side walls which converge at the trailing edge to define a cooling passage of substantially triangular cross section; a plurality of guide vanes arranged within the passage, spaced from the leading wall and trailing edge, and configured so that cooling gas flow introduced a generally radial direction is forced to flow in a direction toward the trailing edge.

Abstract: In a thermally loaded blade for a turbomachine, the front edge of such a blade is made from a ceramic material for providing better protection. The front edge has a shell-shaped form which covers at least one part of the head stage against which flow occurs. The connection between front edge and head stage is produced on the one side by a coupling-forming indentation and on the other side by an elastic element. The elastic element produces a retaining action at the location where it is placed and a frictional connection, having an effect on the indentation, on the other side of the blade. Cooling air bores in the head stage provide for a cooling effect on the ceramic front edge.

Abstract: A coolable airfoil having a passage for discharging particulates entrained in cooling air is disclosed. Various construction details are developed to provide clean cooling air to the leading edge of the airfoil. In one embodiment, a chordwise duct extends past a spanwisely extending supply passage to the inlet of a spanwisely extending rear passage such that momentum effects direct the flow of a portion of cooling air and particulates from the duct, past the entrance of the supply passage and into the rear passage and a larger portion of the flow of cooling air into the supply passage for cooling the leading edge of the airfoil.

Abstract: A double wall hot gas path part for turbines, such as an airfoil, combustor, duct or shroud, comprises an outer skin and an inner support wall that are metallurgically bonded to one another. The double wall contains integral channels for passage of cooling air adjacent to the skin. The skin may be a metal alloy skin or a microlaminate structure, including microlaminate composite structures. Microlaminate composites typically have a lower density than that of the material used for the support wall, and a simplified internal geometry, which promote weight reductions in the parts and increases in engine operating efficiency.

Abstract: Internal heat transfer of air cooled turbine blades of a gas turbine engine is enhanced by orienting the impingement holes that interconnect the air-cooled feed passage to the hybrid passageway internally of the blade in a more radial direction to minimize pressure drop losses. By improved mixing of the cooling air in the hybrid passageway the additional mechanical turbulence promoters can be utilized without sacrificing pressure losses.

Abstract: An air-cooled gas turbine blade providing improved cooling characteristics is disclosed. The turbine blade includes several internal passages for conveying cooling air therethrough during turbine operation to provide the desired cooling effect. Two distinct passages are provided to cool the airfoil leading and trailing edges, respectively. Two serpentine cooling passages are disposed so as to efficiently cool each side of the airfoil. Disposed in the middle of the airfoil is an additional, distinct passage. The platform is cooled by three serpentine cooling passages. Two of these passages are in outlet fluid communication with the inlet to the middle airfoil passage. As cooling air traverses these two passages, heat is transferred, from the base simultaneously cooling it and warming the air. This warmed air is next directed through the middle airfoil passage, providing a slight warming effect to the center portion of the airfoil.

Abstract: An apparatus for cooling the trailing edge portion of a gas turbine vane. Two radially extending passages connected to the outer shroud direct cooling fluid to a plenum formed about mid-span adjacent the trailing edge. Two arrays of cooling fluid passages extend from the plenum. One array extends radially outward toward the outer shroud while the other array extends radially inward toward the inner shroud. The plenum distributes the cooling fluid to the two arrays of passages so that it flows radially inward and outward to manifolds formed in the inner and outer shrouds. The manifolds direct the spent cooling fluid to a discharge passage.

Abstract: A turbine blade including an airfoil section having a double-wall construction for side-wall impingement cooling on the pressure side and a multi-pass serpentine along the suction side of the blade, is described. More particularly, and in one embodiment, the airfoil section includes a pressure side wall and a suction side wall which are joined together at a leading edge and a trailing edge. The blade also includes a leading edge, or tip, and a trailing edge, or tail. The airfoil section also includes a leading edge cavity having a plurality of radial film air holes, and an inner cavity which is a three pass serpentine. As cooling air flows along the passageways, it convectively cools the portions of the turbine blade adjacent these passageways. The airfoil section further includes a trailing edge cavity to cool the trailing edge flow region of the airfoil section.

Abstract: A turbulator configuration is formed on an inner surface of one or more walls defining a cooling passage in a rotor blade of a gas turbine engine, where the cooling passage has a centerline extending therethrough. The turbulator configuration includes a plurality of spaced turbulator pairs oriented with respect to the centerline so that the direction of a boundary layer flow of the coolant adjacent the cooling passage inner surface is consistent with a pair of counter-rotating flow circulations created by Coriolis forces on the coolant flowing into and through the cooling passage when the rotor blade is rotated, whereby heat transfer between the coolant and the wall is enhanced. The turbulator pairs are angled either toward or away from the centerline for leading and trailing walls of the cooling passage depending upon whether the coolant is flowing either radially outward or radially inward.

Abstract: A gas turbine rotor blade includes a shroud which is cooled effectively with cooling gas used for cooling the blade profile so that the temperature of the shroud is reduced and the life of the gas turbine rotor blade is extended. A two-step groove is engraved in the shroud along the tip of the blade profile and the upper portion of the two-step groove is plugged. Second cooling holes are bored along the direction of the plane of the shroud so as to be connected to and communicate with the first cooling holes which are bored in the blade profile in a direction along the longitudinal axis of the blade for passing cooling gas through the blade. Consequently, most of the cooling gas, after cooling the blade profile, can be used for cooling the shroud, so that the shroud is cooled effectively and the temperature thereof is prevented from rising. Preventing the temperature from rising can also prevent reduction of the creep resistance of the shroud and prevent the blade root on the shroud from being turned up.

Abstract: A gas turbine rotor, more particularly a gas turbine rotor blade, which is manufactured so as to be longer, larger, and thin walled and is provided on a rear side of a gas turbine blade array. The gas turbine rotor blade is cooled with cooling air flowing interiorly thereof. A cavity is provided to facilitate the flow of cooling air. The cavity is formed inside a rotor root and inside a hub unit disposed adjacent to a rotor profile unit. Projections are provided inside of the cavity so that the projections, which protrude from an inner wall of the cavity, project into the cooling air flow. Consequently, cooling efficiency of the rotor blade can be improved, and the strength of the portions forming the cavity can be significantly increased. The device also allows the core for cooling to be manufactured and easily set.

Abstract: A cooling structure for the leading edge area of an airfoil having a plurality of passages wherein each passage has a radial component and a downstream component relative to the leading edge axis, and the outlet of each passage has a diffuser area formed by conical machining, wherein the diffuser area is recessed in the wall portion downstream of the passage.

Abstract: A cooled turbine guide vane comprises a hollow aerodynamic portion between inner and outer platforms. The interior of the aerodynamic portion is partitioned into a leading edge duct and a main cavity in which a perforated tubular member is disposed, being spaced from the intrados and extrados side walls of the vane by longitudinal ribs. The tubular member is divided by a partition into two cavities on the intrados and extrados sides of the partition. A first cooling circuit is constituted by the leading edge duct and the intrados cavity of the tubular member, and a second cooling circuit is constituted by the extrados cavity and a cooling system for the inner platform, both circuits being supplied with air by the same source from the outer platform. Cooling air from each circuit passes through the perforations of the tubular member to impinge on the inside face of the respective side wall of the vane and is then guided towards the trailing edge where it escapes through slits in the trailing edge.

Abstract: An airfoil for use in a turbomachine such as a stationary vane in a gas turbine. The airfoil has a plurality of longitudinally extending ribs in its trailing edge region that form first cooling fluid passages extending from the airfoil cavity to the trailing edge of the airfoil. The first cooling fluid passages are tapered so that their height and width decrease as they extend toward the trailing edge. Turbulating fins are spaced along the length of each passage to increase the heat transfer. The ribs have a plurality of radially extending passages spaced along their length so as to form an array of interconnected longitudinal and radial passages. The airfoil is formed by a casting process using a core that has longitudinal and radial fingers that correspond to the longitudinal and radial passages of the airfoil.

Abstract: A method for forming a slot in a metal component having first and second opposite surfaces utilizes a first laser beam having a pulse rate and power to vaporize the metal. The laser beam is traversed across the first surface at a feed rate so that each beam pulse vaporizes the component metal at a spot. The laser beam pulse rate and feed rate are effective so that successive spots do not substantially overlap each other. The laser beam is traversed in repeated steps in a series of passes so that the spots collectively form a continuous slot to a depth below the first surface. In a subsequent step, a second laser beam is positioned in the slot to drill a hole through a base of the slot, with the second laser beam being repositioned to drill a plurality of the through holes spaced apart from each other along the length of the slot.

Abstract: A method for cooling a stator vane is provided, comprising the steps of: (a) providing a hollow stator vane having a high pressure and a standard pressure chamber disposed within the hollow stator vane, adjacent the leading edge of the airfoil, and a supply chamber, disposed within the hollow stator vane, aft of the high and standard pressure chambers, and forward of the trailing edge; the stator vane further includes first and second inlet apertures, and first and second exit apertures; the first inlet apertures extend between the high pressure chamber and the supply chamber, and the second inlet apertures extend between the standard pressure chamber and the supply chamber; the first exit apertures extend between the high pressure chamber and the exterior of the stator vane, and the second exit apertures extend between the standard pressure chamber and the exterior of the stator vane; (b) determining the magnitudes of the gas flow pressure gradient facing the stator vane, and the position of the gradient rel

Abstract: A turbulator configuration formed on an inner surface of at least one wall of a cooling passage for a rotor blade in a gas turbine engine is disclosed, where the cooling passage has an upstream end, a downstream end, and a longitudinal axis therethrough. The turbulator configuration includes a first set of longitudinally spaced turbulators having a specified height and length for tripping a thermal boundary layer along the cooling passage wall to enhance heat transfer between the coolant and the wall, as well as a second set of longitudinally spaced turbulators having a specified height and length for generating turbulence in a core flow region of the cooling passage to reduce the thickness of the thermal boundary layer downstream therefrom.