Abstract: A turbine blade with a squealer tip, the squealer tip being formed by a tip rail extending along the suction side wall of the blade from a point just passed the leading edge of the tip to the trailing edge of the blade, and the tip floor being slanted from the tip rail downward to the pressure side wall of the blade. A first cooling hole opens onto the tip floor at a location adjacent to the pressure side wall of the tip rail, the wall being concave in shape to redirect the flow along the curvature toward the pressure side. A second cooling hole located on the pressure side wall and slanted upward injects cooling air to push the hot gas flow up and over the tip floor. In a second embodiment, the tip floor includes a deflector upstream from and adjacent to the first cooling hole to push the flow upward from the first cooling hole.

Abstract: A blade of a fluid-flow machine has at least one cavity 2 arranged in the blade 1, with the cavity 2 connecting to a fluid supply, and at least one outlet opening 3 connecting the cavity 2 to a flow path, in which the blade 1 is arranged. The outlet opening 3 is of a nozzle-type design and essentially inclined in a direction of a main flow extending longitudinally to a corresponding surface of the blade 1. The outlet opening 3 extends over at least a part of the blade 1 height and the outlet opening 3 forms a step in a contour of the blade 1 profile, with a fluid jet issuing from the outlet opening 3 essentially tangentially attaching onto the blade surface.

Abstract: An impeller for preventing air-lock in a pump, the impeller in its basic embodiment includes a top plate, the top plate having a center axis and a peripheral edge, a plurality of vanes, each vane attached to the top plate, at least one small protrusion projecting from the periphery of the top plate, and at least one ventilation channel having an outlet located near the protrusion, and an inlet located near the center axis of the top plate. The present invention also contemplates a centrifuge pump and an aerator device having the impeller of the present invention.

Abstract: The rotor blade structure for the pneumatic devices includes a blade mounted into the preset chamber of a drive rotor of a pneumatic device. The blade with predefined edge and thickness has an assembly side, an external side and two windshield surfaces. At least one windshield surface is provided with a recessed portion. There may be a single recessed portion or a plurality of recessed portions. Hence, the recessed portion could increase the compression area of blade's windshield surface, thus improving the driving force of the blade and torsion of drive rotor with higher practicability.

Abstract: A turbine engine blade has an attachment root, a platform outboard of the attachment root, and an airfoil extending from the platform. The airfoil has pressure and suction sides extending between leading and trailing edges. An internal cooling passageway network includes at least one inlet in the root and a plurality of outlets along the airfoil. The passageway network includes a leading spanwise cavity fed by a first trunk. A streamwise cavity is inboard of a tip of the airfoil. A spanwise feed cavity feeds the streamwise cavity absent down-pass. A second trunk feeds the spanwise feed cavity.

Abstract: A blade or a vane preferably used in a gas turbine engine, the blade or vane being formed of a metallic core or insert with a ceramic airfoil surface bonded to the metallic insert. The insert includes a series of concave and convex portions arranged along the outer surface, in which a plurality of hoop fibers are secured within a resin matrix to firmly secure the metallic insert to the ceramic airfoil body. Axial loads between the insert and the ceramic body are transferred to the hoop fibers. The hoop fibers are strong enough to resist extending in length that would permit the hoop fibers from sliding over the widest portion of the insert.

Abstract: An article includes a blade casting core combination. The combination includes a ceramic feedcore and a metallic core. The ceramic feedcore has: a root end; a tip end; a leading end; a trailing end; a first side; a second side; and a plurality of legs extending between the root and tip ends and arrayed between the leading and trailing ends. The metallic core has: a first face; a second face; a first portion extending from the feedcore trailing end; and a second portion extending from the tip end. The article may be a pattern where the core is embedded in a wax or may be a shell formed from such a pattern. The article may be used in a method for forming the resultant blade.

Abstract: A blade for a rotor of a wind energy turbine includes opposite upper and lower faces, opposite leading and trailing edges connecting the upper and lower and at least one air outlet opening arranged on at least one of the upper face, the lower face and the trailing edge for discharging pressurized air into the air around at least one of the upper and lower faces and the leading and trailing edges.

Abstract: A hollow blade has an internal cooling passage, an open cavity situated at the free end of the blade and defined by an end wall extending over the entire end of the blade and by a rim. The blade includes cooling channels connecting the internal cooling passage and the outside face of the pressure side wall. The pressure side wall presents a projecting end portion whose outside face is inclined relative to the outside face of the pressure side wall. The cooling channels are disposed in the end portion, being parallel to the outside face of the end portion so that they open out into the tip of the end portion towards the free end of the blade.

Abstract: Covers on bucket tips are provided with an access slot along a circumferentially facing edge of the cover at a location where moisture removal grooves of the bucket will intersect with the cover. Machine tooling is then applied to form the grooves along the suction side and adjacent the leading of the bucket. The machining runs out through the access slot forming grooves along the side edge of the bucket cover within the slot. The axial extent of the covers overlying leading and trailing edges of the buckets is maintained while enabling moisture removal by centrifugal action.

Abstract: A turbine airfoil usable in a turbine engine and having at least one cooling system. The cooling system may include a pressure side serpentine cooling channel and a suction side serpentine cooling channel. The cooling channels may be nested within each other to optimize heat exchange between the cooling fluids and the materials forming the airfoil, to reduce the amount of cooling fluids required, to reduce the required pressure of the cooling fluids, and to provide other benefits. The pressure side serpentine cooling channel may pass cooling fluids chordwise towards the trailing edge, and the suction side serpentine cooling channel may pass cooling fluids chordwise towards the leading edge.

Abstract: An airfoil for a gas turbine engine includes a root, a tip, a leading edge, a trailing edge, and opposed pressure and suction sidewalls extending generally along a radial axis. The airfoil includes a tip cap extending between the pressure and suction sidewalls; and spaced-apart suction-side and pressure-side tip walls extending radially outward from the tip cap to define a tip cavity therebetween. The pressure-side tip wall includes a continuously concave curved arcuate portion, at least a section of which extends circumferentially outward from a radial axis of the airfoil.

Abstract: An airfoil for a gas turbine engine includes a root, a tip, a leading edge, a trailing edge, and opposed pressure and suction sidewalls extending generally along a radial axis. The airfoil includes a tip cap extending between the pressure and suction sidewalls; and spaced-apart suction-side and pressure-side tip walls extending radially outward from the tip cap to define a tip cavity therebetween. The pressure-side tip wall includes a continuously concave curved arcuate portion, at least a section of which extends circumferentially outward from a radial axis of the airfoil. At least a portion of the pressure-side tip wall is recessed from the pressure sidewall to define an outwardly facing tip shelf, such that the pressure-side tip wall and the tip shelf define a trough therebetween.

Abstract: A turbine blade with a tip squealer and method of rebuilding a turbine blade for a gas turbine engine. The blade is of the type including an airfoil having first and second spaced-apart sidewalls defining an interior void and joined at a leading edge and a trailing edge. The first and second sidewalls extending from a root disposed adjacent the dovetail to a tip cap for channeling combustion gases, and a squealer tip including at least one tip rib extending outwardly from the tip cap. The method includes the steps of removing the squealer tip, including the at least one rib tip, from the tip cap and adding new material to the tip cap to serve as a new squealer tip. A plurality of spaced-apart notches is formed in the new material between the leading edge and the trailing edge of the airfoil. At least one hole is formed in each notch communicating with the interior void of the airfoil for channeling cooling air from the interior void of the airfoil to thereby form a squealer tip.

Abstract: A turbine bucket includes a cooling circuit through a dovetail section, a shank section and an airfoil section. The cooling circuit is configured to maximize cooling ability and maximize a useful life at base load operation at firing temperatures of up to 2084° F. while minimizing negative effects on performance. The cooling circuit includes a plurality of cooling holes having predetermined positions and sizes, resulting in increased cooling flow near a trailing edge of the airfoil section and effecting turbulation in the airfoil section to increase bulk and local creep margins throughout the airfoil section.

Abstract: A blade has an airfoil body having an internal cooling passageway network and a body tip pocket. At least one plate is secured within the body tip pocket and has inboard and outboard surfaces. A recess is in the outboard surface and an associated protrusion is on the inboard surface.

Abstract: A blade for a rotor of a gas turbine engine is constructed with a spar and shell configuration. The spar is constructed in an integral unit or multi-portions and includes a first wall adjacent to the pressure side and a second wall adjacent to the suction side, a tip portion extending in the spanwise direction and extending beyond the first wall and the second wall and a root portion extending longitudinally, an attachment portion having a central opening for receiving the root portion and a platform portion. The root portion fits into the central opening and is secured therein by a pin extending transversely through the attachment and the root portion. The shell fits over the spar and is supported thereto by a plurality of complementary hooks extending from the spar and shell. The ends of the shell fit into grooves formed on the tip portion and the platform.

Abstract: The cooling system includes a plurality of generally radially extending passages within the airfoil for convectively cooling the airfoil. A predetermined number of the passages exit through the airfoil tip. One or more of the remaining passages exit into a plenum adjacent the trailing edge and airfoil tip region of the airfoil and flow radially inwardly along one or more passageways adjacent the trailing edge region. The passageways exit in openings along the pressure side of the airfoil. In this manner, the trailing edge region is convectively cooled as well as film cooled by the film of air exiting the holes.

Abstract: A turbine blade tip plenum has means for preferentially directing or diverting cooling air from a cooling passageway network to cool areas subject to extreme heat.

Abstract: A turbine blade for a turbine engine having at least one secondary flow deflector proximate to a blade tip for reducing the effective flow path between the blade tip and an adjacent outer seal. The turbine blade may be a superblade having a central opening forming a hollow turbine blade. The turbine blade may include a secondary flow deflector on upstream sides of the pressure side wall and the suction side wall. The downstream sides of the pressure and suction side walls may include chamfered corners. The secondary flow deflector reduces the effective flow path between the blade tip and an outer seal in numerous ways.

Abstract: A tip cap for a turbine blade. The tip cap may include a HS-188 sheet material with a thickness of less than about 0.079 inches (about 2 millimeters) and a number of holes positioned in the sheet material.

Abstract: An airfoil. The airfoil may include a first number of cooling holes and a second number of cooling holes positioned within the airfoil. The first number of cooling holes and the second number of cooling holes each may include a turbulated section and a non-turbulated section.

Abstract: A turbine blade squealer tip has a continuous squealer tip wall extending radially outwardly from and continuously around a tip cap. A recessed tip wall portion of the tip wall is recessed inboard from a pressure side of an airfoil outer wall of an airfoil of the blade forming a tip shelf therebetween. A plurality of film cooling shelf holes are disposed through the tip shelf to an internal cooling circuit of the blade and are spaced away from a junction between the recessed tip wall portion and the tip shelf. The exemplary embodiment of the airfoil includes shelf hole centerlines of the holes passing through pierce points in the shelf. At least a majority of the shelf hole centerlines are angled in outboard directions away from and outboard of the squealer tip wall. A majority of centerlines are angled away from vertical lines passing through the pierce points at first component angles in a range between 2 degrees and 16 degrees.

Abstract: A cooled fluid flow component for a combustion engine which directs cooling fluid through complementary guided-flow regions to ensure effective cooling of the component tip end, without producing overcooled regions. The component includes multiple channels fluidly linked by a first turning zone. A contoured boundary member divides the turning zone into two guided-flow regions which cooperatively ensure that the tip is cooled appropriately. According to one aspect of the invention, the first guided-flow region forms a vortex that cools a region adjacent a channel-dividing partition, while the second guided flow region ensures the region adjacent the component tip is cooled appropriately. A method of cooling a internally-cooled fluid guide component is also provided.

Abstract: A turbine blade including an airfoil having a profile in accordance with Table 1 is disclosed. The turbine blade has a plurality of cooling passages extending radially outward through the airfoil. The aerodynamic profile of the airfoil has been reconfigured to reduce airfoil stress for a given operating temperature, thereby resulting in increased creep rupture life of the airfoil.

Abstract: A method facilitates fabricating a rotor assembly for a gas turbine engine. The method comprises providing a plurality of rotor blades that each include an airfoil, a dovetail, a shank, and a platform, wherein the platform extends between the shank and the airfoil, and wherein the dovetail extends outwardly from the shank, and forming a cooling circuit within a portion of the shank to supply cooling air to the rotor blade for supplying cooling air to the rotor blade for impingement cooling a portion of the rotor blade and for supplying cooling air to the rotor blade for purging a cavity defined downstream from the rotor blade.

Abstract: A turbine stage one bucket has an airfoil having a plurality of cooling holes passing through the airfoil from 0% span to 100% span whereby cooling air exits the airfoil tip into the hot gas path. X and Y coordinate values are given in Table I, locating the holes relative to the airfoil profile at airfoil profile sections of 5%, 50% and 90% span, Table I also giving the hole diameters. In this manner, cooling hole optimization for this turbine bucket airfoil is achieved. The cooling holes are also located in relation to the profile of the bucket airfoil given by the X, Y and Z coordinate values of Table II, the two coordinate systems having the same origin.

Abstract: A turbine blade has a platform and an airfoil extending from a root at the platform to a tip. The airfoil has an internal cooling passageway network including at least one trailing edge cavity. A group of trailing edge holes extends from the trailing edge to the trailing edge cavity and a group of tip holes extends from the tip to the trailing edge cavity.

Abstract: A turbine blade including an airfoil having a profile in accordance with Table 1 is disclosed. The turbine blade has a plurality of cooling passages extending radially outward through the airfoil. The aerodynamic profile of the airfoil has been reconfigured to further reduce overall heat load to the airfoil while paying particular attention to the leading edge region. Specifically, the airfoil leading edge, which is the life-limiting location of the turbine blade has been reconfigured to lower heat load and allow for increased cooling, thereby increasing turbine blade life.

Abstract: An open cooling circuit for a gas turbine airfoil and associated tip shroud includes a first group of cooling holes internal to the airfoil and extending in a radially outward direction generally along a leading edge of the airfoil; a second group of cooling holes internal to the airfoil and extending in a radially outward direction generally along a trailing edge of the airfoil. A common plenum is formed in the tip shroud in direct communication with the first and second group of cooling holes, but a second plenum may be provided for the second group of radial holes. A plurality of exhaust holes extends from the plenum(s), through the tip shroud and opening along a peripheral edge of the tip shroud.

Abstract: Holes 38 and 39 have upstream opening portions 38b and 39b and downstream opening portions 38a and 39a which have a larger cross-sectional area than upstream opening portions 38b and 39b, and are formed at top portion TP of each moving blade. Holes 38 and 39 have tapered shapes T1 and T2 or step portions, and preferably, downstream opening portions 38a and 39a are eccentrically formed toward the moving direction. When tip squealer 37 is formed, hole 38 is formed so that its opening portion is provided at the side surface of tip squealer 37. Without covering the holes for cooling which are formed at the top portion of the turbine blade due to rubbing or the like, the turbine blade is accurately cooled and stably driven.

Abstract: An improved turbine blade having a tip configured to provide a reduced blade tip clearance between the blade tip and the inside of a turbine casing and to provide an improved flow path for the discharge of cooling air from the turbine blade tip.

Abstract: A method for fabricating a rotor blade for a gas turbine engine facilitates reducing operating temperatures of a tip portion of the rotor blade. The method comprises forming an airfoil including a first sidewall and a second sidewall connected at a leading edge and a trailing edge to define a cavity therein, wherein the first and second sidewalls extend radially between a rotor blade root and a rotor blade tip plate, and forming a first tip wall extending from the rotor blade tip plate along the first sidewall, such that at least a portion of the first tip wall is at least partially recessed with respect to the rotor blade first sidewall to define a tip shelf that extends from the airfoil trailing edge towards the airfoil leading edge. A second tip wall is formed to extend from the rotor blade tip plate along the second sidewall.

Abstract: Imperfect design of shroud or identical parts of the blades of steam/gas turbines, including the adjoining seals, leads to the decrease in reliability and efficiency. These drawbacks are eliminated by way of drilling of radial holes in the shroud of the blades. The transfer of steam through the shroud holes results in the relief of the pressures gradient on the surface of the shroud and prevents the formation of metal oxide, salt and other deposits on the inner surfaces of the shroud. The quantity and diameter of the holes, as well as their corresponding disposition and the values of radial clearances in shroud crowning seals regulate their efficiency.

Abstract: An air turbine starter with an improved seal assembly is provided. A fluid conduit or channel allows higher pressure fluid into one side of a bearing seal. For example, the seal assembly may include a face seal, and a second seal forming a chamber therebetween and a flow passage opens into the chamber for exposing the chamber to the higher pressure.

Abstract: A curved turbulator configuration is in a radial cooling passage of an airfoil, where the radial cooling passage is defined by at least a leading wall and a trailing wall, and the airfoil includes a tip and a root. The curved turbulator configuration includes a number of spaced curved turbulator pairs positioned along a center-line on an inner surface of the leading wall and a number of spaced complementary curved turbulator pairs positioned along a center-line on an inner surface of the trailing wall. An electrode for forming the curved turbulator configuration in the radial cooling passage includes a leading face having a curved turbulator pattern, a trailing face having a complementary curved turbulator pattern, a conductive core, and an insulating coating disposed thereon that is partly removed.

Abstract: A gas turbine engine blade includes pressure and suction sides extending between leading and trailing edges and root to tip. The pressure side includes a tip rib recessed therein to define a tip shelf terminating in an inclined ramp. The ramp may be aligned with streamlines of combustion gas flow for preventing interruption thereof and the increase of heat transfer therefrom.

Abstract: A gas turbine blade (1) with a tip squealer (7) comprises cooling passages (12, 13) that provide cooling fluid for the cooling of the tip squealer (7). A first portion of the cooling passages (12) extends from an internal hollow space (11) through an end cap to a tip pocket. A second portion (13) of the cooling passage extends from the end cap in part through the rails (8) of the tip squealer (7) to the tip crown (9). The second portion (13) is partially bounded by the rail (8) and partially open to the tip pocket (14). The cooling passages according to the invention provide an improved cooling of the tip squealer and have the advantage that cooling fluid can cool the tip squealer (7) even in the event of clogging of the second portions (13) of the passages.

Abstract: An open cooling circuit for a gas turbine bucket wherein the bucket has an airfoil portion, and a tip shroud, the cooling circuit including a plurality of radial cooling holes extending through the airfoil portion and communicating with an enlarged internal area within the tip shroud before exiting the tip shroud such that a cooling medium used to cool the airfoil portion is subsequently used to cool the tip shroud.

Abstract: A blade (1) for a gas turbine comprises a tip cap (4) and a tip squealer (6) and passages (15, 25) for cooling fluid extending from a hollow space (5) to the tip squealer (6). According to the invention the tip squealer (6) comprises a cavity (16) extending from the tip pocket (9) into the tip squealer (6) such that the cooling passage is divided into first and second portions (17, 18) with an exit hole (11′) in the cavity (16) and an exit hole (11) on the tip crown (8) respectively. In case of a blockage of the exit hole (11) on the tip crown (8) cooling fluid can flow through the additional exit hole (11′) into the tip pocket (9) and cool the squealer (6).

Abstract: An airfoil (1) for a gas turbine comprises a pressure sidewall (2) and a suction sidewall (3), a tip cap (4), and a tip squealer (6). Cooling fluid flows from a hollow space (5) within the airfoil (1) through exit passages (14, 16) extending to the pressure side of the airfoil (1) and to the tip cavity (9) defined by the tip cap (4) and the tip squealer (6). According to the invention, the tip squealer (6) comprises a smooth contour with curved and straight portions (10, 11, 12, 13) allowing an even flow of the cooling fluid (21) about the tip squealer (6) and within the tip cavity (9). Furthermore, the exit passages (14, 16) are partially diffused and oriented at angles with respect to the radial direction. The smooth contour and shape of the exit passages avoids the formation of vortices and enhances film cooling of the tip squealer (6).

Abstract: An improved gas turbine blade is disclosed that utilizes a blade tip shroud with tapered shroud pockets to remove excess weight from the shroud while not compromising shroud bending stresses. Excess shroud weight removed by the tapered pockets reduces overall blade pull while maintaining material necessary to increase shroud stiffness. The resulting rib created between the tapered pockets provides a surface, away from possible areas of stress concentration, for drilling radial cooling holes.

Abstract: In an air-cooled turbine blade (10) which has a shroud-band element (11) at the blade tip, the shroud-band element (11) extending transversely to the blade longitudinal axis, hollow spaces (16, 16′, 17, 17′) for cooling being provided in the interior of the shroud-band element (11), which hollow spaces (16, 16′, 17, 17′) are connected on the inlet side to at least one cooling-air passage (18) passing through the turbine blade (10) to the blade tip and open on the outlet side into the exterior space surrounding the turbine blade (10), the hollow spaces (16, 16′, 17, 17′) and the shroud-band element (11) are matched to one another in shape and dimensions in order to reduce the weight of the shroud-band element (11).

Abstract: A cooling system for cooling of the squealer tip surface region of a high pressure turbine blade used in a gas turbine engine and a method for making a system for cooling of the squealer tip surface region of a high pressure turbine blade used in a gas turbine engine. The method comprises the steps of channeling apertures in a tip cap to a diameter of about 0.004″ to about 0.020″ to allow passage of cooling fluid from a cooling fluid source; applying a bond coat of about 0.0005″ to about 0.010″ in thickness to the tip cap such that the bond coat partially fills the channels; applying a porous TBC layer of at least about 0.003″ in thickness to the bond coat, such that the porous TBC fills the channels; applying a dense ceramic TBC layer over the porous layer; and, passing cooling fluid from a cooling fluid source through the channel into the porous TBC. The density of the dense TBC layer can be varied as needed to achieve desired cooling objectives.

Abstract: A gas turbine engine hollow airfoil has an airfoil outer wall with widthwise spaced apart pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of the airfoil and extending radially from a radially inner base to a radially outer airfoil tip. A cooling circuit within the airfoil has radially extending first, middle, and last channels arranged respectively in series with the first channel in fluid communication with a source of cooling air from outside the airfoil. The last channel is in fluid communication with one of the edges. A refresher passageway extends through a radially inner wall bounding a radially inner portion of the last channel and is in fluid communication with the source of cooling air. The refresher passageway is separate, spaced apart, and independent from the first channel. In one embodiment, a metering plate covers an inlet to the refresher passageway and the metering plate has a metering hole over the inlet and the metering hole is adjustable.

Abstract: A gas turbine engine hollow airfoil includes an airfoil outer wall having width wise spaced apart pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of the airfoil and extending radially outward from a base to a tip. The airfoil includes a single three pass aft flowing serpentine cooling circuit and a straight through leading edge cooling channel extending radially through the airfoil and bounded in part by the leading edge. The exemplary embodiment includes a trailing edge cooling plenum and a straight through single pass trailing edge feed channel for providing cooling air to the trailing edge cooling plenum.

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

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

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

Abstract: In an air-cooled turbine blade which, at the blade tip, has a shroud-band element extending transversely to the longitudinal axis of the blade, a plurality of cooling bores passing through the shroud-band element for the purpose of cooling, which cooling bores are connected on the inlet side to at least one cooling-air passage running through the turbine blade to the blade tip and open on the outlet side into the exterior space surrounding the turbine blade, improved and assured cooling is achieved owing to the fact that the cooling bores run from inside to outside in the shroud-band element at least approximately parallel to the direction of movement of the blade and in each case open upstream of the outer margin of the shroud-band element into a surface recess open toward the exterior space.