| Class / Patent application number | Description | Number of patent applications / Date published |
|---|
| 416097000 | Laminated or porous skin | 9 |
| 20080219854 | TURBINE COMPONENT WITH AXIALLY SPACED RADIALLY FLOWING MICROCIRCUIT COOLING CHANNELS - An airfoil for a gas turbine engine component such as a turbine blade or a vane includes at least one microcircuit cooling channel having a plurality of sub-channels extending along a radial direction of the airfoil. The plurality of channels are axially spaced, and are fed by radially spaced inlets. | 09-11-2008 |
| 20080219855 | Turbine blade with micro-turbine nozzle provided in the blade root - A turbine blade, in the blade root ( | 09-11-2008 |
| 20080226462 | CAST FEATURES FOR A TURBINE ENGINE AIRFOIL - An airfoil for a turbine engine includes a structure having a cooling passage that has a generally radially extending cooling passageway arranged interiorly relative to an exterior surface of the structure. The cooling passageway includes multiple cooling slots extending there from toward the exterior surface and interconnected by a radially extending trench. The trench breaks the exterior surface, and the exterior surface provides the lateral walls of the trench. The airfoil is manufactured by providing a core having multiple generally axially extending tabs and a generally radially extending ligament interconnecting the tabs. The structure is formed about the core to provide the airfoil with its exterior surface. The ligament breaks the exterior surface to form the radially extending trench in the exterior surface of the structure. | 09-18-2008 |
| 20080240927 | Turbine blade for a turbine with a cooling medium passage - A turbine blade for a turbine of a thermal power plant, with a platform for partial delimiting of a flow passage in the turbine, wherein the platform has at least one cooling medium passage, which extends inside the platform, for guiding a cooling medium, is characterized according to the invention in that the at least one cooling medium passage emerges from the platform at at least two connecting openings, and the turbine blade has at least one supplementary component which can be fastened on the platform, with a communicating passage, which is designed for interconnecting the connecting openings in a fluid-guiding manner. | 10-02-2008 |
| 20080260538 | Spar and shell constructed turbine blade - 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 foot 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 foot 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, The shell is made from a high temperature resistant material, such as Molybdenum or Niobium, and is formed from a wife EDM process. | 10-23-2008 |
| 20080273987 | Turbine blade having a convergent cavity cooling system for a trailing edge - A turbine blade including an airfoil defining an airfoil cavity forming a cooling system in the blade. First, second and third ribs are positioned in the airfoil cavity to form first, second and third generally elongated cooling cavities along at least a portion of the span-wise direction of the airfoil in an area adjacent the trailing edge of the airfoil. Each of the ribs includes a plurality of orifices for conveying a cooling fluid into each of the cavities. Each of the cavities includes a pair of converging walls, angling inwardly relative to an outer surface of the airfoil, to increase the impingement of cooling fluid from the orifices onto the cavity walls, and increase the cooling effectiveness within the trailing edge of the airfoil. | 11-06-2008 |
| 20080273988 | Aerofoils - An aerofoil | 11-06-2008 |
| 20080279695 | Enhanced turbine airfoil cooling - The ends of cooling air passages in turbine blades and/or vanes of a gas turbine engine are provided with turbulation promoters to enhance the cooling of such structures as inner and outer shrouds and the like to accommodate thermal loads thereon. | 11-13-2008 |
| 20080279696 | Airfoil for a turbine of a gas turbine engine - An airfoil for a turbine of a gas turbine engine is provided. The airfoil comprises a main body comprising a wall structure defining an inner cavity adapted to receive a cooling air. The wall structure includes a first diffusion region and at least one first metering opening extending from the inner cavity to the first diffusion region. The wall structure further comprises at least one cooling circuit comprising a second diffusion region and at least one second metering opening extending from the first diffusion region to the second diffusion region. The at least one cooling circuit may further comprise at least one third metering opening, at least one third diffusion region and a fourth diffusion region. | 11-13-2008 |
| 20080279697 | Turbine airfoil with enhanced cooling - An airfoil for a turbine of a gas turbine engine is provided comprising an outer wall structure defining at least one inner cavity adapted to receive a cooling fluid. The wall structure comprises at least one cooling fluid path circuit communicating with the at least one inner cavity. The cooling fluid path circuit comprises: at least one metering opening extending from an inner surface of the wall structure such that the metering opening communicates with the at least one inner cavity; at least one intermediate diffusion region communicating with the metering opening; an intermediate metering opening positioned downstream from the intermediate diffusion region and communicating with the intermediate diffusion region; and, an end diffusion region positioned downstream from the intermediate metering opening for communicating with the intermediate metering opening and extending to an exit in an outer surface of the wall structure. | 11-13-2008 |
| 20080286115 | Blade for a gas turbine engine - A main body is provided for a gas turbine engine comprising an outer structure, a first internal partition and a second internal partition. The outer structure and the first internal partition may define an entrance leg of a cooling circuit for receiving a cooling fluid. The second internal partition may include a metering slot. The outer structure, the first internal partition and the second internal partition may define an intermediate leg of the cooling circuit. The intermediate leg may communicate with the entrance leg. The second internal partition and the outer structure may define an exit leg of the cooling circuit. The metering slot meters cooling fluid as it passes from the intermediate leg into the exit leg. | 11-20-2008 |
| 20080286116 | Cooling arrangement - An aerofoil for a gas turbine engine comprising a pressure wall and a suction wall and defining leading and trailing edges, the walls define a passage into which is supplied a cooling fluid, an array of cooling holes is provided through at least one of the walls to allow the cooling fluid to flow from an interior surface to an exterior surface. The array of holes comprise two groups, the holes of each group are angled to intersect the holes of the other group and are characterised in that the holes of at least one of the groups comprises two or more holes at different angles to one another to vary the porosity of the wall to account for otherwise varying wall temperatures. This arrangement also allows either less coolant mass flow to maintain a constant metal temperature, or a lower metal temperature for a given coolant mass flow. | 11-20-2008 |
| 20080310965 | Gas-turbine blade featuring a modular design - A gas-turbine blade has a root | 12-18-2008 |
| 20090004023 | DEVICE FOR COOLING THE SLOTS OF A ROTOR DISK IN A TURBOMACHINE HAVING TWO AIR FEEDS - The invention relates to a device for cooling slots in a turbomachine rotor disk, comprising a rotor disk having a plurality of slots and a flange. The device also comprises blades, each having its root mounted in a respective slot, a retaining annulus having one end mounted against the upstream radial face of the disk, and a flange disposed around the flange of the disk and co-operating therewith to define a space forming a cooling air diffusion cavity that opens out into the bottom of each of the slots, together with air admission orifices opening out into the diffusion cavity at the upstream end thereof, the end of the retaining annulus that is mounted against the upstream radial face of the disk including a plurality of openings that open out radially into the bottom of each of the slots in the disk, at the upstream ends thereof. | 01-01-2009 |
| 20090010765 | Reinforced Airfoils - A reinforced airfoil includes an airfoil body including opposed walls that define a hollow interior space and a reinforcement member provided on at least one of the walls within the interior space, the reinforcement member increasing the thickness of the at least one wall so as to resist deformation of the at least one wall but not extending from one wall to the other. | 01-08-2009 |
| 20090041586 | TURBINE NOZZLE SECTOR - A turbine nozzle sector that comprises an outer platform segment and an inner platform segment between which there extend one or more hollow vanes. Each vane presents a trailing edge cavity for feeding with cooling air and communicating with a plurality of vents distributed along the trailing edge of the vane, these vents serving to exhaust a fraction of the cooling air. Said cavity communicates with an air outlet hole situated level with the outer platform and enabling a fraction of the cooling air to be exhausted. | 02-12-2009 |
| 20090041587 | TURBINE BLADE WITH INTERNAL COOLING STRUCTURE - A rotating blade ( | 02-12-2009 |
| 20090047136 | Angled tripped airfoil peanut cavity - A turbine airfoil comprises a wall portion, a cooling channel, an impingement rib, impingement rib nozzles, turbulators and leading edge cooling holes. The wall portion comprises a leading edge, a trailing edge, an outer diameter end, and an inner diameter end. The cooling channel receives cooling air and extends through an interior of the wall portion between the inner diameter end and the outer diameter end. The impingement rib is positioned within the wall portion forward of the cooling channel and between the outer diameter end and the inner diameter end to define a peanut cavity. The impingement rib nozzles extend through the impingement rib for receiving cooling air from the cooling channel. The turbulators are positioned within the peanut cavity to locally influence the flow of the cooling air. The leading edge cooling holes discharge the cooling air from the peanut cavity to an exterior of the wall portion. | 02-19-2009 |
| 20090068021 | Thermally balanced near wall cooling for a turbine blade - A turbine blade including an airfoil having an airfoil outer wall extending radially outwardly from a blade root to a blade tip. The airfoil outer wall includes a pressure sidewall and a suction sidewall, and the pressure and suction sidewalls are joined together at chordally spaced leading and trailing edges of the airfoil. A pressure side serpentine cooling path extends adjacent the pressure sidewall and a suction side serpentine cooling path extends adjacent the suction sidewall. The pressure side cooling path conducts cooling fluid in a first chordal direction between the leading and trailing edges, and the suction side cooling path conducts cooling fluid in a second chordal direction, opposite the first chordal direction, between the leading and trailing edges. A central partition extends chordally through the airfoil, and a transverse passage extends through the central partition and connects the pressure side cooling path to the suction side cooling path. | 03-12-2009 |
| 20090068022 | Wavy flow cooling concept for turbine airfoils - An airfoil including an outer wall and a cooling cavity formed therein. The cooling cavity includes a leading edge flow channel located adjacent a leading edge of the airfoil and a trailing edge flow channel located adjacent a trailing edge of the airfoil. Each of the leading edge and trailing edge flow channels define respective first and second flow axes located between pressure and suction sides of the airfoil. A plurality of rib members are located within each of the flow channels, spaced along the flow axes, and alternately extending from opposing sides of the flow channels to define undulating flow paths through the flow channels. | 03-12-2009 |
| 20090068023 | Multi-pass cooling for turbine airfoils - An airfoil for a turbine vane of a gas turbine engine. The airfoil includes an outer wall having pressure and suction sides, and a radially extending cooling cavity located between the pressure and suction sides. A plurality of partitions extend radially through the cooling cavity to define a plurality of interconnected cooling channels located at successive chordal locations through the cooling cavity. The cooling channels define a serpentine flow path extending in the chordal direction. Further, the cooling channels include a plurality of interconnected chambers and the chambers define a serpentine path extending in the radial direction within the serpentine path extending in the chordal direction. | 03-12-2009 |
| 20090074589 | Cooling Circuit for Enhancing Turbine Performance - In a gas turbine having a compressor discharge casing, a cooling circuit diverts compressor discharge air toward a high pressure packing (HPP) circuit. The cooling circuit includes an inlet pipe that receives compressor discharge air. One or several cooled cooling air pipes are in fluid communication with the inlet pipe via a pipe manifold, which distributes the discharge air across the cooled cooling air pipes. A seal is disposed upstream of an entrance to the HPP circuit to limit flow into the HPP circuit, and a second seal is disposed downstream of the HPP circuit at turbine wheelspace to limit ingestion and thus the purge flow air required. The circuit serves to reduce required purge flow in the HPP circuit so that an amount of compressor discharge air can be put back to the main flow path, thereby improving turbine performance. | 03-19-2009 |
| 20090081048 | Turbine Blade for a Turbine - The invention relates to a blade for a turbine comprising a blade wall, a first channel for guiding a first medium and a second channel for guiding a second medium that can be supplied to the turbine blade separately from the first medium. In order to combine both media, which are supplied separately, into one mixture, a turbine blade has least one chamber which is arranged in the interior or in the blade wall and said chamber is connected to said channels via a respective connection line. In order to provide a particularly simple component that is economical to produce, the chamber and/or the outlet conduit are a least partially delimited and/or formed by an insert accommodated in the wall. | 03-26-2009 |
| 20090092500 | HOLLOW TURBOMACHINE BLADE - A hollow turbomachine blade including an internal cooling passage, an open cavity situated at the free end of the blade and defined by an end wall and the side wall of at least one rim which extends between the leading edge and the trailing edge of the blade, and at least one cooling channel connecting said internal cooling passage to said open cavity, said cooling channel opening out at the base of the rim and the wall of the rim forming an angle relative to said end wall that is obtuse, being strictly greater than 90°. An indentation may be formed in the wall of the rim at the outlet from said cooling channel. Said blade advantageously does not include a pressure rim. | 04-09-2009 |
| 20090104042 | TURBINE AIRFOIL WITH NEAR WALL MULTI-SERPENTINE COOLING CHANNELS - A turbine airfoil usable in a turbine engine and having at least one cooling system. At least a portion of the cooling system may be positioned in an outer wall of the turbine airfoil and be formed from at least one suction side serpentine cooling chamber and at least one pressure side serpentine cooling chamber. Each of the suction and pressure side serpentine cooling channels may receive cooling fluids from a cooling fluid supply source first before being passed through other components of the cooling system. The cooling fluids may then be passed into a mid-chord cooling chamber to cool internal aspects of the turbine airfoil, yet prevent creation of a large temperature gradient between outer surfaces of the turbine airfoil and inner aspects. | 04-23-2009 |
| 20090123292 | Turbine Blade Tip Cooling System - A turbine blade for a turbine engine having a cooling system in the turbine blade formed from at least one elongated tip cooling chamber forming a portion of the cooling system and at least partially defined by the tip wall proximate to the first end. An inner surface of the tip wall may include a plurality of curved bumper protrusions extending from the inner surface radially inward toward the root. The cooling system may include a plurality of ribs generally aligned with the trailing edge, and the curved bumper protrusions may be offset in a chordwise direction relative to the ribs. A throat section may extend between a first forwardmost curved bumper protrusion and a second immediately adjacent downstream curved bumper protrusion and may be offset radially outward from an inner tip surface, thereby creating a first recessed tip slot with a reduced tip wall thickness. | 05-14-2009 |
| 20090129934 | Turbine Blade Tip Cooling System - A turbine blade for a turbine engine having a cooling system in the turbine blade including a camber-line rib extending radially outward from a tip of the blade and extending from a trailing edge of the blade toward a leading edge. The camber-line rib may form pressure and suction side cooling slots at the tip of the blade. The camber-line rib my include a cooling channel positioned in the camber-line rib and in fluid communication with the at least one cavity forming the cooling system in the blade for cooling aspects of the tip at the trailing edge. | 05-21-2009 |
| 20090148305 | TURBINE BLADES AND METHODS OF MANUFACTURING - A turbine blade includes a convex suction side wall, a concave pressure side wall, a tip wall, an internal cooling circuit, and a plurality of tip edge channels. The tip wall is recessed from a first tip edge of the suction side wall and a second tip edge of the concave pressure side wall to define a suction side wall tip section and a pressure side wall tip section, and the suction side wall tip section is shorter than the pressure side wall tip section. The internal cooling circuit is formed at least partially between the convex suction side wall, the concave pressure side wall, and the tip wall. The plurality of tip edge channels formed through the first tip edge of the convex suction side wall extend to the internal cooling circuit. Methods of manufacturing turbine blades are also provided. | 06-11-2009 |
| 20090169395 | Tungsten shell for a spar and shell turbine vane - The present invention is a vane for us in a gas turbine engine, in which the vane is made of an exotic, high temperature material that is difficult to machine or cast. The vane includes a shell made from Tungsten, and is formed from a wire electric discharge process. The shell is positioned in grooves between the outer and inner shrouds, and includes a central passageway within the spar, and forms a cooling fluid passageway between the spar and the shell. Both the spar and the shell include cooling holes to carry cooling fluid from the central passageway to an outer surface of the vane for cooling. This cooling path eliminates a serpentine pathway, and therefore requires less pressure and less amounts of cooling fluid to cool the vane. | 07-02-2009 |
| 20090175732 | BLADE UNDER PLATFORM POCKET COOLING - Inlets are provided at a front end of inter-blade cavities for allowing coolant to flow therein to cool down the undersurface of the blade platforms as well as the rim of the disc of a rotor assembly. | 07-09-2009 |
| 20090175733 | AIR COOLED TURBINE BLADES AND METHODS OF MANUFACTURING - An air-cooled turbine blade and methods of manufacturing the blade are provided. The blade includes a suction side flow circuit formed within its interior and defined at least by an interior surface of a convex suction side wall, a pressure side flow circuit formed within the blade interior and defined at least by an interior surface of a concave pressure side wall, and a center flow circuit including a first section and a second section, the first section disposed between the suction side flow circuit and the pressure side flow circuit, and the second section in flow communication with the first section and a plurality of openings of a leading edge wall and defined at least partially by an interior surface of the leading edge wall. | 07-09-2009 |
| 20090185913 | Method of Producing a Turbine or Compressor Component, and Turbine or Compressor Component - Disclosed is a turbine or compressor component with an integrated cooling channel, in particular a turbine blade, and a method for producing the same. The aim of the invention is to ensure an improved estimation of the service life of the component and furthermore, if possible, also increased safety during operation and increased service life, even in the presence of constantly variable thermal and mechanical stress. To achieve this, the cooling channel of the component is subjected to internal pressure during a pressure impingement phase, said internal pressure being at a level sufficiently high that it causes the at least semiplastic deformation of the wall regions delimiting the cooling channel. | 07-23-2009 |
| 20090202357 | COOLED PUSHER PROPELLER SYSTEM - A propulsion system and method includes an annular exhaust nozzle about an axis radially outboard of the annular cooling flow nozzle and ejecting an exhaust flow through an annular exhaust nozzle about an axis radially outboard of the annular cooling flow nozzle. | 08-13-2009 |
| 20090202358 | BLADE WITH A COOLING GROOVE FOR A BLADED WHEEL OF A TURBOMACHINE - A blade for a turbomachine bladed wheel, the blade including an airfoil and a platform with at least one air injection passage, the platform including a groove running along the pressure side of the airfoil at least in the vicinity of a downstream portion thereof, and formed between the pressure side and a ridge formed on the platform surface at a short distance from said downstream portion of the pressure side, at least one air injection passage being arranged in the groove. Because of the presence of the groove, the air stream injected via the air injection passage(s) is kept close to the pressure side and thus provides effective cooling of the downstream portion thereof. | 08-13-2009 |
| 20090208343 | SERPENTINE MICROCIRCUITS FOR HOT GAS MIGRATION - A turbine engine component, such as a turbine blade, has an airfoil portion with a pressure side and a suction side. The turbine engine component also has a first cooling circuit within the pressure side for cooling the pressure side of the airfoil portion and a second cooling circuit within the suction side for cooling the suction side of the airfoil portion and for cooperating with a wrap around leading edge cooling circuit for creating a cooling film over the pressure side. | 08-20-2009 |
| 20090208344 | Process For The Electrolytic Treatment Of A Component, And A Component With Through-Hole - There is described a method wherein through holes of a wall are treated on a inside, and wherein one respective pole electrode is assigned to each through hole that is to be processed. | 08-20-2009 |
| 20090214355 | FIXING METHOD FOR A TIP WINGLET AND REDUCED TIP LEAKAGE BLADE - Reducing the tip leakage for an existing compressor or turbine blade ( | 08-27-2009 |
| 20090220349 | Method for Producing a Gas Turbine Component Which is to be Coated, With Exposed Holes, Device for Carrying Out the Method, and Coatable Turbine Blade with Film Cooling Holes - There is described a method with which, on the basis of a three-dimensional recording of a component to be reworked, its surface configuration can be determined and stored temporarily so that, after it has been coated, it can be produced in its original surface form, or in its surface form then required, in certain regions, i.e. locally in the area of film-cooling openings. An especially precise and quick three-dimensional recording can be achieved by the use of the triangulation method. In this case, a reference pattern depicted on the component by a projector is recorded by two camera arranged at an angle. From the images from the cameras, the coordinates describing the surface three-dimensionally can then be determined by a control system using the triangulation method. | 09-03-2009 |
| 20090232660 | Blade for a gas turbine - A blade is provided for a gas turbine. The blade comprises a main body comprising a cooling fluid entrance channel; a cooling fluid collector in communication with the cooling fluid entrance channel; a plurality of side channels extending through an outer wall of the main body and communicating with the cooling fluid collector and a cooling fluid cavity; a cooling fluid exit channel communicating with the cooling fluid cavity; and a plurality of exit bores extending from the cooling fluid exit channel through the main body outer wall. | 09-17-2009 |
| 20090232661 | Turbine blade with multiple impingement cooled passages - A turbine blade with an airfoil wall having a serpentine flow cooling circuit formed within the wall that includes within each channels that flow toward the blade tip of the serpentine a series of impingement holes and impingement chambers such that the cooling air flowing through the channels of the serpentine forms a multiple impingement cooling passages through the channels. Each channel includes a series of slanted ribs that define the impingement chambers, and each slanted rib includes an impingement cooling hole to direct impingement cooling air onto the backside surface of the wall exposed to the hot gas flow. The channel of the serpentine that flows toward the blade root contains no metering holes and is substantially unobstructed to the cooling air flow. The rotation of the blade produces a centrifugal force on the airflow passing through the channels with the metering and impingement holes to aid in the flow towards the blade tip. The return channels are unobstructed in order to minimize the pressure loss on the return channel of the serpentine circuit. | 09-17-2009 |
| 20090238694 | RADIAL SPLIT SERPENTINE MICROCIRCUITS - A turbine engine component, such as a turbine blade has an airfoil portion with an airfoil mean line, a pressure side, and a suction side. A first region on the pressure side of the airfoil portion has a first array of cooling microcircuits embedded in a wall forming the pressure side. A second region on the pressure side has a second array of cooling microcircuits embedded in the wall. The first region is located on a first side of the mean line and the second region is located on a second side of the mean line. | 09-24-2009 |
| 20090238695 | FULL COVERAGE TRAILING EDGE MICROCIRCUIT WITH ALTERNATING CONVERGING EXITS - A turbine engine component has an airfoil portion with a pressure side wall, a suction side wall, and a trailing edge. The turbine engine component further has at least one first cooling circuit core embedded within the pressure side wall, with each first cooling circuit core having a first exit for discharging a cooling fluid, at least one second cooling circuit core embedded within the suction side wall, with each second cooling circuit core having a second exit for discharging a cooling fluid, and the first and second exits being aligned in a spanwise direction of the airfoil portion. | 09-24-2009 |
| 20090252615 | Cooled Turbine Rotor Blade - A cooled turbine rotor blade for a gas turbine which is traversed axially by flow and is equipped with an attachment area and an airfoil profile is provided. Meandering cooling channels with interposed deflecting regions are provided in the interior of the airfoil profile. In the deflecting regions, it is possible to prevent dead water regions, which are generated in the prior art, by virtue of at least one of the ribs running so as to curve towards the leading edge or towards the trailing edge in the region of the airfoil tip. At the same time, an opening is provided in the curvature of the rib. Through this opening a part of the coolant flows in the deflecting region and can pass over into the adjacent cooling duct. | 10-08-2009 |
| 20090297361 | Minimization of fouling and fluid losses in turbine airfoils - Contaminant build-up and cooling airflow looses are reduced in a turbine airfoil by joining root and airfoil cooling air passages thereof with a transition passage. | 12-03-2009 |
| 20090304520 | SERPENTINE COOLING CIRCUIT AND METHOD FOR COOLING TIP SHROUD - A serpentine cooling circuit is formed in a gas turbine blade to cool portions of the tip shroud, primarily the fillet between the airfoil and the tip shroud and the shroud edges. | 12-10-2009 |
| 20090317258 | Rotor blade - Cooling within aerofoils ( | 12-24-2009 |
| 20090324422 | Cascade tip baffle airfoil - A turbine blade includes an airfoil tip with first and second ribs extending along the opposite pressure and suction sides. The ribs extend outwardly from a tip floor and are joined together at opposite leading and trailing edges. A cascade tip baffle transversely bridges the two ribs above the tip floor forward of the maximum width of the tip to partition the tip chordally into corresponding tip pockets on opposite sides of the baffle. | 12-31-2009 |
| 20090324423 | TURBINE AIRFOIL WITH CONTROLLED AREA COOLING ARRANGEMENT - A gas turbine airfoil ( | 12-31-2009 |
| 20090324424 | AIR COOLED BUCKET FOR A TURBINE - A bucket for a turbine is provided. The bucket includes an airfoil having a root portion, a tip portion, an airfoil shape, and a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table I, wherein Z is a distance from a platform on which the airfoil extends outwardly from, and X and Y are coordinates defining the profile at each distance Z from the platform. The bucket also includes a plurality of cooling passages extending between the root portion and tip portion of the airfoil, each of the plurality of cooling passages exiting at the tip portion, the plurality of cooling passages positioned in a camber line pattern. | 12-31-2009 |
| 20090324425 | PARTICLE RESISTANT IN-WALL COOLING PASSAGE INLET - A cooling microcircuit for a turbine engine component has a first cooling passage which has at least one inlet oriented in a radially outward direction for preventing particles from entering the cooling passage and for dislodging particles which may become lodged in the at least one inlet. | 12-31-2009 |
| 20100014985 | SHROUD SEGMENT COOLING CONFIGURATION - An cooling arrangement for a turbine shroud provides an air stream directed from a passage extending through a radial wall to impinge on a surface in an area of a back side of the turbine shroud, the surface defining a plane which improves the attack angle of the air stream to the surface. | 01-21-2010 |
| 20100014986 | SEALING ELEMENT FOR A GAS TURBINE, A GAS TURBINE INCLUDING SAID SEALING ELEMENT AND METHOD FOR COOLING SAID SEALING ELEMENT - A sealing element for a gas turbine, which is provided with at least one rotor ring and at least a plurality of rotor blades radially arranged about the rotor ring and having an end portion fixed to the rotor ring, is provided with a wall, which is adapted to be coupled to the rotor ring and to at least one rotor blade, has an external face adapted to be arranged in contact with a hot working fluid in the gas turbine and an internal face adapted to be arranged in contact with a cooling fluid of the gas turbine, and a gap, which is adapted to be travelled through by the cooling fluid. | 01-21-2010 |
| 20100034662 | COOLED AIRFOIL AND METHOD FOR MAKING AN AIRFOIL HAVING REDUCED TRAIL EDGE SLOT FLOW - An airfoil component having a body having a leading edge and a trailing edge, a ceramic casting insert for making the component and the method for making the component. The component includes an internal cooling passageway and an elongated opening in communication with the internal cooling passageway. The opening is configured with a geometry that provides structural stability during casting and has a cross-section that sufficiently restricts airflow through the opening to provide efficient component operation. The casting insert includes outer edge projections and a web portion corresponding to the geometry of the openings when cast around the insert. The method includes casting the airfoil component around the casting insert and removing the insert to provide the component having the openings. | 02-11-2010 |
| 20100034663 | GAS TURBINE ENGINE ASSEMBLIES WITH VORTEX SUPPRESSION AND COOLING FILM REPLENISHMENT - A gas turbine engine assembly has combustion gases flowing through a gas flow path. The gas turbine engine assembly includes a stator assembly comprising a stator vane that extends into the gas flow path; and a turbine rotor assembly downstream of the stator assembly and comprising a turbine platform and a turbine rotor blade extending from the turbine platform into the mainstream combustion gases flow path. The turbine rotor blade includes a pressure side and a suction side opposing the pressure side that extend from a leading edge to a trailing edge. The combustion gases form horseshoe vortices at a formation area adjacent the leading edge of the turbine rotor blade, and the turbine rotor assembly further includes a first set of holes in the turbine platform for directing first jets into the formation area of the horseshoe vortices. | 02-11-2010 |
| 20100040478 | Cooled Airfoils and Gas Turbine Engine Systems Involving Such Airfoils - Cooled airfoils and gas turbine engine systems involving such airfoils are provided. In this regard, a representative cooled airfoil includes: an exterior surface defining a leading edge, a trailing edge, a suction side and a pressure side; an interior surface defining an interior cavity; trenches in the exterior surface oriented spanwise along the leading edge; and cooling holes communicating between the interior cavity and the trenches such that cooling air provided to the interior cavity flows from the interior cavity though the cooling holes into the trenches, the cooling holes having exterior apertures located in the trenches and interior apertures located at the interior surface. | 02-18-2010 |
| 20100040479 | Gas Turbine Engine Systems Involving Baffle Assemblies - Gas turbine engine systems involving baffle assemblies are provided. In this regard, a representative baffle assembly for a gas turbine engine includes: a cooling plenum defining a cooling air path; and a baffle sized and shaped to extend between surfaces of the cooling plenum such that a cooling air path of reduced cross-section is formed between the baffle and the surfaces, the baffle being operative to increase a flow rate of cooling air as the cooling air directed to the cooling air path is redirected through the cooling air path of reduced cross-section. | 02-18-2010 |
| 20100040480 | Cooling arrangement - With regard to cooling turbine blades in a gas turbine engine a compromise has to be made between convective cooling within the inner cavity defining a flow path for coolant and the blow rates for developing film cooling on an outer surface of the aerofoil. By providing a chamber between the flow cavity and external apertures reconciliation between the necessary flow rates for convective cooling within the cavity defining the pathway for coolant flow within the aerofoil and the necessary coolant blowing rate for film development can be achieved. | 02-18-2010 |
| 20100047078 | Blade - Cooling arrangements have been provided for blades and in particular turbine blades utilising gas turbine engines. Generally for internal strength a leading passage has been separate by a solid wall from a feed passage as impingement apertures may diminish structural strength as centres for stress concentration. However, impingement apertures allow impingement jets which have improved cooling efficiency. By providing a leading passage which is divided at least into a lower section and an upper section the lower section can have a wall which is solid for structural integrity whilst an upper section has impingement apertures for greater cooling efficiency. | 02-25-2010 |
| 20100054953 | AIRFOIL WITH LEADING EDGE COOLING PASSAGE - A turbine engine airfoil includes an airfoil structure having an exterior surface that provides a leading edge. A first cooling passage includes radially spaced legs extending laterally from one side of the leading edge toward another side of the leading edge and interconnecting to form a loop with one another. A trench extends radially in the exterior surface along the leading edge. The trench intersects one of the first and second legs to provide at least one first cooling hole in the trench. | 03-04-2010 |
| 20100054954 | TURBINE BUCKET FOR A TURBOMACHINE AND METHOD OF REDUCING BOW WAVE EFFECTS AT A TURBINE BUCKET - A turbine bucket for a turbomachine includes a main body portion having a base portion and an airfoil portion, the base portion includes a bucket cavity forward region and a shank cavity. The turbine bucket also includes a cooling channel that extends through the main body portion. At least one flow passage extends between one of the cooling channel and the shank cavity, toward the bucket cavity forward region. The at least one flow passage delivers a flow of cooling gas toward the bucket cavity forward region. The flow of cooling gas limits ingestion of hot gases into the bucket cavity forward region. | 03-04-2010 |
| 20100054955 | Blades - A rotary blade, such as a turbine blade for a gas turbine engine, has an aerofoil portion with a tip partly shrouded by winglets. A gutter extends across the radially outer face of the tip to leave upstands. Cooling air feed galleries are drilled into each upstand, from the trailing edge, toward the upper end of a cooling air feed void, which is spaced from the trailing edge. Cooling passages are drilled from the winglet edges to the gallery. Cooling air supplied through the void passes along the gallery, through the passages and leaves the blade at the cooling holes. This allows cooling to be provided near the trailing edge of the tip without requiring the geometry around the trailing edge to be thickened to accommodate a cooling air void. | 03-04-2010 |
| 20100068066 | SYSTEM AND METHOD FOR GENERATING MODULATED PULSED FLOW - A device includes a fluid flow channel having a channel inlet for receiving a pressurized fluid for flow through the fluid flow channel and a channel outlet for discharging the pressurized fluid therefrom. A passive flow element is situated within the fluid flow channel or proximate to the channel inlet. The passive flow element includes an element inlet for receiving the pressurized fluid, and an element outlet. The passive flow element also includes a cavity for receiving the pressurized fluid from the element inlet and generating a periodic flow variation of the pressurized fluid so as to modulate the pressurized fluid flow rate through the element outlet. | 03-18-2010 |
| 20100068067 | Turbine Airfoil Cooling System with Divergent Film Cooling Hole - A cooling system for a turbine airfoil of a turbine engine having at least one divergent film cooling hole positioned in an outer wall defining the turbine airfoil is disclosed. The divergent film cooling hole includes a first section extending from an inner surface of the outer wall into the outer wall and a second section extending the first section and terminating at an outer surface of the outer wall. The divergent film cooling hole may provide a metering capability together with a divergent section that provides a larger film cooling hole breakout and footprint, which creates better film coverage and yields better cooling of the turbine airfoil. The divergent film cooling hole may provide a smooth transition, which allows the film cooling flow to diffuse better in the second section of the divergent film cooling hole. | 03-18-2010 |
| 20100068068 | Turbine Airfoil Cooling System with Diffusion Film Cooling Hole Having Flow Restriction Rib - A cooling system for a turbine airfoil of a turbine engine having at least one diffusion film cooling hole positioned in an outer wall defining the turbine airfoil is disclosed. The diffusion film cooling hole includes first and second sections. The first section may function as a metering section, and the second section may function as a diffusion section. The second section may include flow restriction ribs that direct the flow of cooling fluids in disproportionately larger amounts proximate to the downstream side of the diffusion film cooling hole. | 03-18-2010 |
| 20100068069 | Turbine Blade - A turbine blade, having a plurality of auxiliary cooling channels which branch off from a main cooling channel, formed within a blade body, is provided. The plurality of auxiliary cooling channels open into outlet openings in the leading edge region of the blade body. A heat shield element is attached to the blade body in the leading edge region at a predefined spacing, wherein the heat shield element has a number of outlet channels which are arranged behind one another in the longitudinal direction and extend from the main cooling channel to the outer wall face of the heat shield element. | 03-18-2010 |
| 20100074762 | Trailing Edge Cooling for Turbine Blade Airfoil - A gas turbine engine hollow turbine airfoil having chordwise spaced apart leading and trailing edges, and widthwise spaced apart pressure and suction sidewalls extending chordwise between the leading edge and the trailing edge. A trailing edge rib extends from the trailing edge toward the leading edge, and forms a solid member between the pressure and suction sidewalls. A cooling fluid channel extends in the spanwise direction through the airfoil adjacent to the trailing edge rib. A plurality of fluid chambers are formed in the trailing edge rib. Film cooling holes extend from the fluid chambers to the pressure and suction sidewalls, and trailing edge discharge holes extend from the fluid chambers to the trailing edge. A metering hole is associated with each of the fluid chambers to define a flow restriction connecting the cooling fluid channel to a respective fluid chamber. | 03-25-2010 |
| 20100074763 | Trailing Edge Cooling Slot Configuration for a Turbine Airfoil - A gas turbine engine hollow turbine airfoil having pressure and suction sidewalls extending chordwise between leading and the trailing edges. The trailing edge includes a pressure sidewall lip and a suction sidewall lip, and a breakout distance between the pressure sidewall lip and the suction sidewall lip. A cooling fluid channel extends spanwise through the airfoil for supplying a cooling fluid to the airfoil. Flow channels are provided extending chordwise between the cooling fluid channel and the suction sidewall lip and include a metering section, an internal diffusion section and a breakout slot. The interior diffusion section includes a spanwise dimension and a widthwise dimension perpendicular to the spanwise dimension, wherein the spanwise dimension continuously increases extending in the chordwise direction, and the widthwise dimension continuously decreases extending in the chordwise direction. | 03-25-2010 |
| 20100080711 | TURBINE BLADE WITH IMPROVED DURABILITY TIP CAP - A gas turbine engine has a turbine blade having a tip cap to close off internal cooling passages. The tip cap is formed with a plurality of purge holes, with there being at least two rows of purge holes. The increased number of purge holes spreads the cooling air outwardly across more of the surface area of the tip cap, and provides the tip cap with a better ability to withstand the extreme temperatures it faces in use. | 04-01-2010 |
| 20100098554 | Blade for a rotor - A blade for a rotor, such as a turbine rotor of a gas turbine engine, has a squealer tip comprising a peripheral wall which defines a cavity. A first region of the peripheral wall extends radially, with its outer surface forming a continuation of the adjacent aerofoil surface of the blade. A second region extends obliquely with respect to the radial direction and the adjacent part of the aerofoil surface. The second region defines a winglet, and serves to increase the width of the chamber towards the trailing edge of the blade. | 04-22-2010 |
| 20100111704 | TURBINE BLADE HAVING SQUEALER - A turbine blade of the invention includes an air foil including a plurality of cooling flow passages through which a cooling medium flows from a leading edge region to a trailing edge region, a top plate which forms the apex of the air foil, has a heat-resistant coating applied on the upper surface thereof, and includes a plurality of cooling holes, and a squealer which protrudes radially outward from the blade from the top plate, and is formed so as to extend from a leading edge end to a starting end of the trailing edge region along a suction-surface-side blade wall in a peripheral direction of the blade. | 05-06-2010 |
| 20100119377 | Cooling arrangement - Within components such as high pressure turbine blades and aerofoils in a gas turbine engine it is important to provide cooling such that these components remain within acceptable operational parameters. Typically, film cooling as well as convective cooling is utilised. Film cooling requires holes from a feed passage from which the coolant is presented upon an external surface to develop the film. The holes themselves can create cooling through convective cooling effects. In order to maximise the convective cooling effect holes are created which have an indirect path about a direct line between an inlet and an outlet for the hole. By creating an indirect path in the form of a helix or spiral which in turn may have a variable cross sectional area from the inlet to the outlet control of coolant flow can be achieved. The inlet may have a bell mouth shape whilst the hole may have a slot or elliptical cross section to achieve greater diffusion of the coolant flow in order to create an improved exit blow rate for instant film development. | 05-13-2010 |
| 20100124508 | TURBINE AIRFOIL COOLING SYSTEM WITH PLATFORM EDGE COOLING CHANNELS - A turbine airfoil of a turbine engine having cooling channels positioned on side surface edges of a platform of the airfoil. The platform may include at least one angled side surface that may be aligned with a side surface of a platform of an adjacent turbine blade. The airfoil may include a suction side edge formed from a first surface at an obtuse angle relative to an upper surface of the platform and a second surface at an obtuse angle relative to a bottom surface of the platform. One or more film cooling slots may be positioned in the first surface and may include a diffusion portion. A damper may also be positioned in a groove between the second surface and an adjacent platform. The damper may include cooling slots on a side surface proximate to the second surface of the suction side edge. | 05-20-2010 |
| 20100129231 | METERED COOLING SLOTS FOR TURBINE BLADES - A metered cooling slot disposed in a wall comprising an outer surface that is exposed to a hot gas stream and an inner surface that defines an internal coolant chamber through which a coolant passes, the metered cooling slot comprising: a slot formed within the outer surface elongated in a first direction, the slot comprising a pair of spaced apart, opposing, slot surfaces and a base, the slot surfaces intersecting the outer surface to form a slot outlet opposite the base; and two or more metering apertures formed within the wall, each metering aperture intersecting the inner surface of the wall to form a metering aperture inlet and intersecting one of the pair of slot surfaces to form a metering aperture outlet; wherein: D represents the approximate diameter of at least two of the metering apertures; P represents the approximate distance between the center lines of at least two neighboring metering apertures; and P/D comprises a value within the range of about 4 to 6. | 05-27-2010 |
| 20100135822 | TURBINE BLADE FOR A GAS TURBINE ENGINE - The turbine blade comprises an airfoil having opposite pressure and suction sidewalls extending from a platform to a free end tip and in a chordwise direction from a leading edge to a trailing edge. The blade has a chamfer extending between the pressure sidewall and the tip. The chamfer extends in a chordwise direction, the blade having a plurality of cooling passageways, each extending from an inlet in fluid communication with a pressurized cooling circuit inside the airfoil to an outlet on the chamfer. | 06-03-2010 |
| 20100143153 | TURBINE BLADE - A turbine blade comprising at least one cooling element, a cooling duct for conducting a cooling medium, and a leading edge is provided. The cooling element is located within the flow of the cooling medium and is designed in a cog-shaped manner. The cooling duct is formed within the turbine blade for conducting a cooling air flow and extends along the flow attacking edge in at least some sections. The cooling elements are successively arranged in a stationary manner inside the cooling duct in the longitudinal direction. Each individual cooling element has a cooling capacity that is adapted to a predefined cooling requirement for the leading edge in the vicinity of the cooling element. | 06-10-2010 |
| 20100143154 | SPACER FOR GAS TURBINE BLADE INSERT - A gas turbine blade has: a main hollow body substantially extending along a longitudinal axis and having a leading edge, a trailing edge, opposite to the leading edge, a suction side, and a pressure side, both comprised between the leading edge and the trailing edge; a hollow cooling element extending along the axis, which is equipped with a plurality of cooling holes and is set within the main body so as to define a gap between the main body and the cooling element; and a rib, which is set within the gap for connecting the main body with the cooling element on the suction side, in an area of the main body without holes for communication between the gap and the outside of the blade. | 06-10-2010 |
| 20100150733 | AIRFOIL WITH WRAPPED LEADING EDGE COOLING PASSAGE - A turbine engine airfoil includes an airfoil structure having an exterior surface providing a leading edge. A radially extending first cooling passage is arranged near the leading edge and includes first and second portions. The first portion extends to the exterior surface and forms a radially extending trench in the leading edge. The second portion is in fluid communication with a second cooling passage. In one example, the second cooling passage extends radially, and the first cooling passage wraps around a portion of the second cooling passage from a pressure side to a suction side between the second cooling passage and the exterior surface. In the example, the first portion is arranged between the pressure and suction sides. In one example, the first cooling passage is formed by arranging a core in an airfoil mold. The trench is formed by the core in one example. | 06-17-2010 |
| 20100150734 | TURBINE BLADE - In a turbine blade in which a tabular rib partitioning the blade into a pressure side and a suction side is provided in substantially the midsection of the blade along a center line connecting a leading edge and a trailing edge, and at least two cavities of which a cavity at the suction side and a cavity at the pressure side do not communicate with each other are provided, a pressure adjustment member for cooling air flowing into and out of the cavity at the suction side reduces the amount of cooling air flowing into the suction side, relative to the pressure side. | 06-17-2010 |
| 20100150735 | Multi-Peripheral Serpentine Microcircuits For High Aspect Ratio Blades - A cooling arrangement for a pressure side of an airfoil portion of a turbine engine component is provided. The cooling arrangement comprises a pair of cooling circuits embedded within a wall forming the pressure side. The pair of cooling circuits includes a first serpentine cooling circuit and a second circuit offset from the first serpentine cooling circuit. | 06-17-2010 |
| 20100158700 | TURBINE BLADE ASSEMBLIES AND METHODS OF MANUFACTURING THE SAME - A turbine blade assembly includes an airfoil, a platform, and a first cover plate. A center flow path extends through the platform in communication with an internal cooling circuit of the airfoil, which extends from a first side of the platform. A second side of the platform is located opposite the platform from the first side. An edge of the platform extends between the first and second sides and, a first passage is formed between the first and second sides and includes a first inlet and a first outlet. The first passage extends from the center flow path toward the platform edge, and a first groove is formed on the second side of the platform and extends from the first outlet of the first passage toward the edge of the platform. The first cover plate is disposed over the second side of the platform covering the first groove. | 06-24-2010 |
| 20100158701 | TURBINE BLADES - A cast turbine blade is disclosed and includes an internal cooling passage that passes (e.g., zig-zags or meanders) through the blade from an inlet in the blade root to an outlet in the blade tip. The cooling passage can have a zone at a bend that is at a distance which is remote from the inlet of the cooling passage when the distance from the inlet is measured around the passage, but that is closer to the inlet when the distance from the inlet is measured in a straight line. During casting of the blade, the cooling passage can be defined by a core or cores having a leachable material, the cores being removed after casting by a chemical leaching process. A supplementary passage is also provided for connecting the remote zone to the inlet during the leaching process. The supplementary passage can likewise be defined by a leachable core, or it may be machined into the blade after casting. During the service life of the blade, a plug can be used to obturate the supplementary passage to prevent leakage of cooling air from the cooling passage through the supplementary passage. | 06-24-2010 |
| 20100172762 | Aerofoil - An aerofoil ( | 07-08-2010 |
| 20100183446 | TURBINE BLADE OR VANE WITH IMPROVED COOLING - Disclosed is a turbine blade or vane including a blade or vane body including a leading edge and a trailing edge, a plurality of cooling openings disposed along the trailing edge, a first width of the trailing edge, the first width being disposed across the cooling openings, and a second width of the trailing edge the second width being disposed between the cooling openings, wherein the second width is smaller than the first width. | 07-22-2010 |
| 20100189569 | ROTOR BLADE - A gas turbine engine rotor blade has an airfoil portion containing one or more internal conduits. Each conduit extends to an end of the airfoil portion. The blade has a shroud at the end of the airfoil portion for sealing the blade to a facing stationary engine portion. There is a fillet portion which joins the end to the shroud. The fillet portion eases the transition from the outer surface of the airfoil portion to the outer surface of the shroud and has a cavity which extends from each conduit and expands laterally relative thereto. The area of the cavity on a cross-section through the fillet portion perpendicular to the radial direction of the engine and at an expanding part of the cavity is greater than the area of the conduit, or the combined areas of the conduits, on a parallel cross-section at the end of the airfoil portion. | 07-29-2010 |
| 20100196167 | ROTOR CHAMBER COVER MEMBER HAVING APERTURE FOR DIRT SEPARATION AND RELATED TURBINE - A cover member defining a rotor chamber adjacent to a rotor wheel that supports a rotating blade in a turbine includes a first aperture for introducing a cooling gas stream into the rotor chamber, and a second aperture positioned in a radially outward portion of the cover member for allowing a portion of the cooling gas stream to exit the rotor chamber. The portion of the cooling gas stream exiting the rotor chamber carries dirt particles to purge the rotor chamber. | 08-05-2010 |
| 20100221121 | Turbine airfoil cooling system with near wall pin fin cooling chambers - A cooling system for a turbine airfoil of a turbine engine having a suction side near wall cooling chamber extending from the leading edge to the trailing edge. The suction side near wall cooling chamber may include a plurality of pin fins for increasing the cooling effectiveness of the suction side near wall cooling chamber. The pin fins may be formed in two or more regions having varying sizes and quantities per unit area to accommodate different cooling requirements across the airfoil. In one embodiment, cooling fluids may flow in a counterflow manner through the suction side near wall cooling chamber relative to a pressure side near wall cooling chamber. In another embodiment, the cooling fluids may flow from the leading edge through the suction side near wall cooling chamber and be exhausted through slots in the trailing edge. | 09-02-2010 |
| 20100221122 | Flared tip turbine blade - A turbine blade includes an airfoil terminating in a tip. The tip includes a first rib conforming with a concave pressure side of the airfoil, and a second rib conforming with a convex suction side of the airfoil. The second rib is flared outwardly from the suction side. | 09-02-2010 |
| 20100221123 | TURBINE BLADE COOLING - A turbine blade with a generally hollow airfoil having an outer wall that defines at least one radially extending chamber for receiving the flow of a coolant, the airfoil including a leading edge that resides in an upstream or forward direction, a trailing edge that resides in a downstream or aft direction, a convex suction side, and a concave pressure side, the turbine blade comprising: a plurality of inserts disposed within the chamber that are configured to initially receive at least a portion of the coolant entering the chamber and direct a substantial portion of the coolant through a plurality of insert apertures toward the inner surface of the outer wall; wherein the inserts are configured to form at least one inward bleed channel and a central collector passage into which the inward bleed channel flows. | 09-02-2010 |
| 20100226788 | TURBINE BLADE WITH INCREMENTAL SERPENTINE COOLING CHANNELS BENEATH A THERMAL SKIN - A turbine blade having an internal cooling system with incremental serpentine cooling channels in near walls forming an outer surface of the turbine blade is disclosed. The turbine blade may be formed from an internal structural spar that is covered with a thermal skin. The incremental serpentine cooling channels may be cut into the outer surface of the spar to which the thermal skin may be attached. The incremental serpentine cooling channels may be formed from two or more serpentine cooling channels aligned along an axis that extends generally spanwise throughout the turbine blade. A row of incremental serpentine cooling channels may extend from a root to a tip of the blade, but a single incremental cooling channel does not. | 09-09-2010 |
| 20100226789 | TURBINE BLADE DUAL CHANNEL COOLING SYSTEM - A turbine blade having an internal cooling system with dual serpentine cooling channels in communication with tip cooling channels is disclosed. In at least one embodiment, the cooling system may include first and second tip cooling channels in communication with the first and second serpentine cooling channels, respectively. The first tip cooling channel may extend from the leading edge to the trailing edge and be formed from a first suction side tip cooling channel and a first pressure side tip cooling channel. The second tip cooling channel may extend from a midchord region toward the trailing edge and may be positioned between the pressure and suction sides such that the second tip cooling channel is positioned generally between the first suction side and pressure side tip cooling channels. The first and second tip cooling channels may exhaust cooling fluids through the trailing edge. | 09-09-2010 |
| 20100226790 | TURBINE BLADE LEADING EDGE TIP COOLING SYSTEM - A turbine blade for a turbine engine having a cooling system in the turbine blade formed from at least one leading edge cooling chamber for cooling the leading edge and for exhausting cooling fluids through a tip exhaust outlet. The leading edge cooling channel may include a second section that extends radially inward from the tip exhaust outlet and a first section that extends radially inward from the second section along the leading edge. The second section may be narrower than the first section and may direct cooling fluids to be exhausted from tip exhaust outlet skewed angle towards the suction side of the blade. As such, the leading edge cooling channel and tip exhaust outlet cooperate to exhaust the cooling fluids without forming a separation zone at the upstream side of the intersection between the leading edge cooling channel and the tip exhaust outlet. | 09-09-2010 |
| 20100226791 | BLADE COOLING STRUCTURE OF GAS TURBINE - A blade cooling structure of a gas turbine, which can reduce the pressure loss of a cooling medium without decreasing the heat transfer coefficient, is provided. The blade cooling structure comprises a cooling passage ( | 09-09-2010 |
| 20100232979 | BLADE TIP COOLING GROOVE - An example turbine blade includes a blade having an airfoil profile extending radially toward a blade tip. A shelf is established in the blade tip. A sealing portion of the blade tip extends radially past a floor of the shelf. The sealing portion extends from a blade tip leading edge to a blade tip trailing edge. A groove is established in the blade tip. The groove extends from adjacent the shelf to adjacent the blade tip trailing edge. The groove is configured to communicate a fluid from a position adjacent the shelf to a position adjacent the blade tip trailing edge. | 09-16-2010 |
| 20100239430 | COOLABLE AIRFOIL ATTACHMENT SECTION - A rotor blade suitable for use in a gas turbine engine includes an attachment section which defines at least one internal cooling passage along a passage axis through the attachment section. | 09-23-2010 |
| 20100239431 | Turbine Airfoil Cooling System with Dual Serpentine Cooling Chambers - A cooling system for a turbine airfoil of a turbine engine having dual serpentine cooling channels, an inward serpentine cooling channel and an outward serpentine cooling channel, positioned within the airfoil. The inward serpentine cooling channel may receive cooling fluids from a cooling supply system through the root and exhaust cooling fluids to the outward serpentine cooling channel at the leading edge. The outward serpentine cooling channel may pass the cooling fluids through the outward portion of the serpentine cooling channel and exhaust the cooling fluids through the trailing edge of the airfoil. Such configuration yields a better creep capability for the blade. | 09-23-2010 |
| 20100239432 | Turbine Vane for a Gas Turbine Engine Having Serpentine Cooling Channels Within the Inner Endwall - A turbine vane for a gas turbine engine having an internal cooling system in fluid communication with cooling channels positioned in the inner endwall is disclosed. The cooling system in the inner endwall may include cooling channels extending outwardly from the leading edge, trailing edge, pressure side and suction side toward the edges of the inner endwall. The cooling channels may be serpentine cooling channels and may be two or more serpentine cooling channels coupled together in series. The cooling channels may exhaust cooling fluids from the inner endwall through a plurality of orifices on an outer surface facing the opposing endwall and on the sides surfaces of the endwall. The pressure side and suction side midchord modulus serpentine flow circuits may receive cooling fluids from one pass of an internal midchord cooling channel and may exhaust those cooling fluids into another pass of the midchord cooling channel. | 09-23-2010 |
| 20100247328 | Microcircuit cooling for blades - A turbine engine component, such as a turbine blade, includes at least one cooling circuit having a plurality of legs through which a cooling fluid flows, and a plurality of cooling devices in at least one of the legs. Each of the cooling devices has a heat transfer multiplier in the range of from 1.8 to 2.4 and a reattachment length in the range of from 1.9 to 2.5. | 09-30-2010 |
| 20100247329 | TURBINE BLADE ASSEMBLIES WITH THERMAL INSULATION - A turbine blade assembly for a gas turbine includes a spar with raised ribs, a spacer with a plurality of protrusions mounted around the spar, and an outer shell mounted around the spacer. The protruding portions on the spacer surround the raised ribs on the spar. The protruding portions of the spacer act to space the interior surfaces of the outer shell away from the spar to provide a thermal insulation layer of cooling air. | 09-30-2010 |
| 20100247330 | LINER IN A COOLING CHANNEL OF A TURBINE BLADE - A turbine blade with a blade root, an aerofoil, at least one cooling passage arranged in the turbine blade and extending from the blade root to the aerofoil, and a liner arranged in the at least one cooling passage is provided. The liner protects the cooling passage against corrosion, especially type II hot corrosion. | 09-30-2010 |
| 20100254823 | HOLLOW ROTOR BLADE FOR THE TURBINE OF A GAS TURBINE ENGINE - A hollow blade includes an internal cooling passage, an open cavity located at the tip of the blade and bounded by an end wall and a rim and cooling channels that connect the internal cooling passage to the outer face of the pressure wall. The cooling channels are inclined to the pressure wall in such a way that they emerge on the outer face of the pressure wall near the top of the rim. A reinforcement of material is present between the rim and the end wall of the cavity along at least one portion of the pressure wall, whereby the rim is widened at its base adjacent to the end wall in such a way that the cooling channels emerge near the top of the rim without reducing the mechanical strength of the tip of the blade. | 10-07-2010 |
| 20100254824 | GAS TURBINE AIRFOIL - A turbine component ( | 10-07-2010 |
| 20100290919 | Gas Turbine Blade with Double Impingement Cooled Single Suction Side Tip Rail - A turbine blade is provided comprising: an airfoil including an airfoil outer wall extending radially outwardly from a blade root, a squealer tip section located at an end of the airfoil distal from the root, and cooling structure. The squealer tip section comprises a blade tip surface including pressure and suction edges joined together at chordally spaced-apart leading and trailing edges of the airfoil, and a squealer tip rail. At least a substantial portion of the squealer tip rail is located near the blade tip surface suction edge. The cooling structure directs cooling fluid toward the squealer tip rail to effect impingement cooling of the rail after the cooling fluid has convectively cooled at least a portion of the airfoil outer wall. Cooling fluid is also deflected by the squealer tip rail so as to yield a very small effective flow area above the squealer tip section through which hot working gases may flow. | 11-18-2010 |
| 20100290920 | Turbine Blade with Single Tip Rail with a Mid-Positioned Deflector Portion - A turbine blade is provided comprising: an airfoil including an airfoil outer wall extending radially outwardly from a blade root, a squealer tip section located at an end of the airfoil distal from the root, and cooling structure. The squealer tip section comprises a blade tip surface including pressure and suction edges joined together at chordally spaced-apart leading and trailing edges of the airfoil, and a squealer tip rail. At least a substantial portion of the squealer tip rail is located near the blade tip surface suction edge. The cooling structure directs cooling fluid toward the squealer tip rail to effect impingement cooling of the rail after the cooling fluid has convectively cooled at least a portion of the airfoil outer wall. Cooling fluid is also deflected by the squealer tip rail so as to yield a very small effective flow area above the squealer tip section through which hot working gases may flow. | 11-18-2010 |
| 20100290921 | Extended Length Holes for Tip Film and Tip Floor Cooling - The tip cooling arrangement of the present application reduces large cooling flow requirements which can compromise turbine performance. The tip cooling arrangement of the present application provides convective cooling of a turbine blade tip end, whether a flat tip or a squealer, by extending holes that provide fluid for film cooling the tip end. The holes are thus lengthened and extend from the relatively cooler suction side of the blade to the pressure side of the blade in close proximity to the floor of the tip end. | 11-18-2010 |
| 20100290922 | TURBINE DISK AND GAS TURBINE - In a turbine disk and a gas turbine, the turbine disk is firmly connected to a rotor ( | 11-18-2010 |
| 20100303635 | COOLING ARRANGEMENTS - Providing cooling within hollow blades such as high pressure turbine blades in a gas turbine engine is important to maintain these components within operational margins for the materials from which they are formed. Traditionally, coolant flows in hollow passages have been used along with impingement apertures towards a leading passage for cooling effectiveness. It is known that opposed undulations or ribs can create rotational vortices within the passage. By shaping shaped portions between the opposed undulations and possibly providing undulations upon these shaped portions themselves it is possible to generate stronger more powerful vortices within the passage. These vortices are coupled with the impingement orifices to create proportionally greater impingement jet flow and pressure and therefore cooling effectiveness within the leading passage. | 12-02-2010 |
| 20100310381 | ACTIVE FILM COOLING FOR TURBINE BLADES - A cooling system includes a surface comprising a plurality of orifices and a flow control plasma actuator positioned proximate an orifice to induce cooling air attachment to the surface. In an exemplary embodiment, the plasma actuator includes a power source, a first electrode in contact with a first dielectric layer and connected to the power source, a second electrode in contact with a second dielectric layer and connected to the power source, and a ground electrode. The power source drives the first electrode with a first ac voltage pattern and drives the second electrode with a second ac voltage pattern. The first voltage pattern and the second voltage pattern have a phase difference. In further embodiments, a dc voltage can be used to drive one or more of the electrodes, where the dc voltage can be pulsed in specific embodiments. In another embodiment, a cooling system includes a suction mechanism positioned proximate an orifice to induce cooling air attachment to the surface, the section mechanism being positioned downstream of the orifice. | 12-09-2010 |
| 20100322783 | Rotor or stator blade and method for forming such rotor or stator blade - A rotor blade or a stator blade for rotary machinery is disclosed where the rotor/stator blade comprises a number of thin blade plates. A plurality of the rotor/stator blades are provided with at least one hole forming at least one supply duct when the blade plates are stacked on top of each other to form the rotor/stator blade. The blade plates are joined together by means of sintering such that a solid rotor/stator blade with an outer surface is formed. The rotor/stator blade further comprises a system of distributing micro ducts extending from the at least one supply duct to the outer surface of the blade, fanning out in such a way that the number of micro ducts extending out to the outer surface of the rotor/stator blade is equal to or greater than the number of micro ducts extending from the at least one supply duct, thereby providing cooling liquid to the outer surface of the rotor/stator blade such that the cooling liquid cools the rotor/stator blade by evaporation when the rotor/stator blade is in use. There is also provided a method for manufacturing of such a rotor/stator blade. | 12-23-2010 |
| 20100329887 | COOLABLE GAS TURBINE ENGINE COMPONENT - A gas turbine engine component coupled with a temperature member configured to assist in maintaining a temperature of the gas turbine engine component below a predetermined temperature. The temperature member can take the form of a phase changeable material that can change phase by melting or vaporization, among potential others. A variety of gas turbine engine components can be coupled with the temperature member. | 12-30-2010 |
| 20100329888 | TURBOMACHINERY BLADE HAVING A PLATFORM RELIEF HOLE, PLATFORM COOLING HOLES, AND TRAILING EDGE CUTBACK - A method is disclosed that includes providing a turbomachinery blade having an airfoil connected to a platform in a root region of the turbomachinery blade. The airfoil has a trailing edge extending from the root region to a tip distal from the root region. The method further includes forming a blind relief hole in the platform proximate the trailing edge of the airfoil, and forming a plurality of cooling holes in the platform. | 12-30-2010 |
| 20110008177 | GAS TURBINE VANE WITH IMPROVED COOLING - The disclosure relates to a hollow gas turbine vane that is cooled by an arrangement configured to sequential cooling an endwall of the vane and its airfoil and, at the same time, the two endwalls of the vane. This arrangement can reduce cooling air demand, which can have a positive effect on the turbines efficiency. | 01-13-2011 |
| 20110020137 | Spar and shell constructed turbine blade - 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. The shell is made from a high temperature resistant material, such as Molybdenum or Niobium, and is formed from a wire EDM process. | 01-27-2011 |
| 20110027102 | COOLING STRUCTURE OF TURBINE AIRFOIL - A cooling structure of a turbine airfoil cools a turbine airfoil ( | 02-03-2011 |
| 20110027103 | IMPELLER WHICH INCLUDES IMPROVED MEANS OF COOLING - Impeller for turbine engine which includes a disk provided with blade retention teeth, wherein associated with each tooth there are means for channelling a flow of cooling air which covers the crest of said tooth so that the latter is swept by the air flow. | 02-03-2011 |
| 20110033311 | Turbine Airfoil Cooling System with Pin Fin Cooling Chambers - A cooling system for a turbine airfoil having at least one pin fin with a dimpled outer surface is disclosed. The dimpled outer surface increases the cooling efficiency of the pin fin, which creates numerous efficiencies, including thermal efficiencies, manufacturing efficiencies and the like. The dimples may be formed from shapes including, but not limited to, circular, oval, racetrack, and hemispherical. The dimples may be aligned in a variety of configurations. | 02-10-2011 |
| 20110033312 | COMPOUND COOLING FLOW TURBULATOR FOR TURBINE COMPONENT - Multi-scale turbulation features, including first turbulators ( | 02-10-2011 |
| 20110038735 | Turbine Vane for a Gas Turbine Engine Having Serpentine Cooling Channels with Internal Flow Blockers - A turbine vane for a gas turbine engine with an internal cooling system formed from a serpentine cooling channel with one or more flow blocking ribs is disclosed. The serpentine cooling channels may be configured to receive cooling fluids from internal cooling fluids supply channels. The serpentine cooling channels may include flow blocking ribs to form concurrent flow channels to reduce the cross-sectional area within the midchord region of the airfoil to maintain the internal through flow channel Mach number. The flow blocking ribs may include slots therein and may have any appropriate configuration. In at least one embodiment, the flow blocking ribs may be have a nonuniform taper or a uniformed taper. | 02-17-2011 |
| 20110044822 | GAS TURBINE BLADE AND GAS TURBINE HAVING THE SAME - Provided is a gas turbine blade capable of improving the heat-conducting capacity of a serpentine channel. In a gas turbine blade including a serpentine channel in which a plurality of cooling channels, extending from the base end side to the distal end side of the blade, are provided from the leading edge to the trailing edge of the blade, at least two of these cooling channels being connected in a folded manner at the base end or distal end, the serpentine channel is formed such that the channel cross-sectional area becomes sequentially smaller from the cooling channel provided at the extreme upstream side of the serpentine channel to the cooling channel provided at the extreme downstream side. | 02-24-2011 |
| 20110052412 | HIGH-PRESSURE TURBINE ROTOR, AND METHOD FOR THE PRODUCTION THEREOF - A method for producing a high-pressure turbine rotor, as well as a turbine rotor, is disclosed. The turbine rotor is designed as a blisk (i.e., bladed disk) and forms a radially inward disk and several vanes or blades that project from the disk. The turbine rotor has an internal duct system for air cooling and at least one section of the turbine rotor is manufactured in a generative production process. | 03-03-2011 |
| 20110052413 | COOLED GAS TURBINE ENGINE AIRFLOW MEMBER - In one embodiment, a gas turbine engine turbine flow member is described that contains a passageway having a flow obstruction that forms a tortuous passageway for the passage of a cooling fluid. The flow obstruction can include flow structures disposed toward either side of the passageway. In one non-limiting embodiment a flow structure can have a V-shape, and in another non-limiting embodiment the flow structure can have an elongate shape. As cooling fluid flows through the passageway it is encouraged to flow up, down, and/or along portions of the flow obstruction. | 03-03-2011 |
| 20110058958 | COOLED AEROFOIL BLADE OR VANE - An aerofoil blade or vane ( | 03-10-2011 |
| 20110064585 | COOLING DUCT ARRANGEMENT WITHIN A HOLLOW-CAST CASTING - Described is a cooling passage arrangement inside a hollow-cast cast part, with a flow region, delimited by at least two spaced apart cast-part walls, for a cooling medium (K), which flow region is divided in the flow direction into two cooling passages ( | 03-17-2011 |
| 20110070096 | CURVED ELECTRODE AND ELECTROCHEMICAL MACHINING METHOD AND ASSEMBLY EMPLOYING THE SAME - An electrode for an electrochemical machining process is provided. The electrode comprises a curved, electrically conductive member, and an insulating coating covering at least a portion of a side surface of the curved, electrically conductive member. An electrochemical machining assembly is also provided for machining curved holes in a workpiece. The assembly includes at least one curved electrode and a power supply operatively connected to provide a pulsed voltage to the at least one curved electrode and to the workpiece. The assembly further includes a rotational driver operatively connected to move the at least one curved electrode along a curved path within the workpiece. The assembly is configured to remove material from the workpiece upon application of the pulsed voltage to the at least one curved electrode and to the workpiece. An electrochemical machining method is also provided for forming one or more curved holes in an electrically conductive workpiece. | 03-24-2011 |
| 20110070097 | GAS TURBINE ENGINE COMPONENT COOLING SCHEME - A gas turbine engine includes a compressor section, a combustor section and a turbine section. The turbine section includes components having a platform and an airfoil extending from the platform. The platform includes an outer surface, a cover plate and a cooling channel extending between the outer surface and the cover plate. The cooling channel receives cooling airflow to cool the platform and the airfoil. | 03-24-2011 |
| 20110076155 | GUIDE BLADE FOR A GAS TURBINE - A guide blade for a gas turbine includes an inner and an outer platform, an airfoil extending in a radial direction between the inner and the outer platforms and having a height in the radial direction, and at least one cooling channel disposed in an interior of the airfoil and configured to receive a cooling medium flowing through the at least one cooling channel configured to cool the guide blade, wherein a cross-sectional area of a blade material of the airfoil varies over the height. | 03-31-2011 |
| 20110085915 | BLADE FOR A GAS TURBINE - A blade for a gas turbine includes a leading edge running in a longitudinal direction substantially radially to an axis of the turbine; a trailing edge running in the longitudinal direction; a blade body disposed between the leading edge and the trailing edge so as to define a pressure side and a suction side; an exit slot disposed in the blade body in an area of the trailing edge and running in the longitudinal direction and configured to discharge a cooling medium from an interior of the blade body; and a row of first and second control elements disposed in the exit slot in a distributed manner in the longitudinal direction and configured to control a mass flow of the cooling medium exiting through the exit slot, the first control elements having a first configuration and the second control elements having a second configuration different from the first configuration. | 04-14-2011 |
| 20110103971 | TURBINE BLADE - The amount of cooling air (cooling medium) can be reduced, and low-temperature cooling air is prevented from being blown out through film cooling holes. Part of a cooling medium impingement-cooling an inner circumferential surface of a blade main body located on a ventral side further impingement-cools the inner circumferential surface of the blade main body located on a dorsal side and is blown out through film cooling holes in the blade main body that are located on the dorsal side. | 05-05-2011 |
| 20110110790 | HEAT SHIELD - A heat shield is disclosed. The heat shield may include a base layer and a spacer layer. The spacer layer may be coupled to the base layer. The spacer layer may define a plurality of flow channels. The base layer and the spacer layer may be configured to associate with a hot gas path component. | 05-12-2011 |
| 20110116937 | ONE-STAGE STATOR VANE COOLING STRUCTURE AND GAS TURBINE - A one-stage stator vane cooling structure consisting of: a plurality of linking members that are provided between a plurality of combustors disposed in the circumferential direction of a gas turbine; and cooling holes that are provided in each of the one-stage stator vanes to discharge a cooling gas from the inside to the outside of the one-stage stator vanes in order to cool the one-stage stator vanes on the periphery of the stagnation line of the combustion gas flowing in from the plurality of combustors; in which the cooling holes are formed at positions that are determined according to the relative positions of the one-stage stator vanes and the linking members disposed near the one-stage stator vanes. | 05-19-2011 |
| 20110123351 | TURBINE VANE AND GAS TURBINE - An airfoil part includes a plurality of cooling chambers that are spaces into which the inside of the airfoil part is partitioned, from a leading edge side to a trailing edge side, by a partition wall, that extend in a vane-longitudinal-section direction, and that include division parts on inner walls of a body; insert cylinders that are disposed in the cooling chambers and that have a plurality of impingement holes; and film holes that are provided in the body. The insert cylinders include partitioning parts that extend from the leading edge side to the trailing edge side and that extend in the vane-longitudinal-section direction. The insides of the insert cylinders are partitioned into pressure-surface-side insert spaces close to a pressure surface and suction-surface-side insert spaces close to a suction surface. | 05-26-2011 |
| 20110135497 | TURBINE BLADE - A blade is provided and includes a platform and a root configured to be connected to a blade carrier. Airfoil portions extend from opposite sides of the platform. Each airfoil portion defines an operating surface being the surface facing the other airfoil portion. An operating surface of one of the airfoil portions defines a suction side and the other operating surface of the other airfoil portion defines a pressure side. | 06-09-2011 |
| 20110150666 | TURBINE BLADE - A turbine blade having an internal skeleton having a plurality of internal ribs that form a plurality of open cooling channels; an internal environmental coating applied to the internal skeleton; an outer wall applied about the open cooling channels of the internal skeleton to form a near wall circuit of cooling channels; and an external environmental coating applied to the outer wall wherein the internal environmental coating is different from the external environmental coating. | 06-23-2011 |
| 20110158820 | COMPOSITE GAS TURBINE ENGINE COMPONENT - One embodiment of the present invention is a unique composite gas turbine engine component. In one form, the composite component is an airfoil. Another embodiment is a unique method for manufacturing a composite gas turbine engine component. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations composite gas turbine engine components. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith. | 06-30-2011 |
| 20110176929 | SYSTEM FOR COOLING TURBINE BLADES - A system, in one embodiment, includes a turbine blade having a radial blade tip. The system further includes a trailing edge trench that is formed in the radial blade tip and which extends towards a trailing edge of the turbine blade. The trailing edge trench further includes a first set of cooling passages, each of which includes a first slot formed along a first sidewall of the trailing edge trench, whereby the slot is coupled to a first respective hole extending through a floor of the trailing edge trench. | 07-21-2011 |
| 20110176930 | Turbine vane for a gas turbine and casting core for the production of such - A turbine vane or blade including an interior structure is provided. In addition, turbulence elements connected directly upstream of openings disposed at the rear edge of the blade of the turbine vane or blade are also provided. These are disposed in a sequence, each having a flow side against which a coolant flows and which is at least partially arched in a concave manner. Preferably, the turbulence elements are configured in a crescent-shaped manner. This makes it possible to enlarge the openings without resulting in an increased consumption of coolant. A casting core is also provided. The openings required in the casting core for the production of the webs of a turbine vane or blade may now be placed at further distances than before. | 07-21-2011 |
| 20110182751 | ROTOR DISC - A rotor disc | 07-28-2011 |
| 20110194943 | COOLED SNUBBER STRUCTURE FOR TURBINE BLADES - A turbine blade assembly in a turbine engine. The turbine blade assembly includes a turbine blade and a first snubber structure. The turbine blade includes an internal cooling passage containing cooling air. The first snubber structure extends outwardly from a sidewall of the turbine blade and includes a hollow interior portion that receives cooling air from the internal cooling passage of the turbine blade. | 08-11-2011 |
| 20110194944 | TURBINE BLADE EQUIPPED WITH MEANS OF ADJUSTING ITS COOLING FLUID FLOW RATE - A turbine blade cooled by an automatically variable internal flow of cooling fluid. The blade includes orifices situated under the blade root via which the cooling fluid penetrates, and an adjustment plate including holes situated in register with the orifices, the adjustment plate having a coefficient of expansion that is different from the coefficient of expansion of the blade root. | 08-11-2011 |
| 20110200449 | MEMBER HAVING INTERNAL COOLING PASSAGE - Provided is a member having an internal cooling passage | 08-18-2011 |
| 20110206536 | TURBINE BLADE WITH SHIELDED COOLANT SUPPLY PASSAGEWAY - A turbine blade for a turbine engine includes main coolant passageways which extend through the turbine blade to cool the blade. A tip coolant passageway conveys coolant from a location adjacent the base of the blade directly to the tip of the blade to provide cooling fluid directly to the tip of the blade. This ensures that the coolant arriving at the tip of the blade is at a relatively low temperature and can therefore provide effective cooling of the material located at the tip of the blade. | 08-25-2011 |
| 20110211972 | Small Scale High Speed Turbomachinery - A small scale, high speed turbomachine is described, as well as a process for manufacturing the turbomachine. The turbomachine is manufactured by diffusion bonding stacked sheets of metal foil, each of which has been pre-formed to correspond to a cross section of the turbomachine structure. The turbomachines include rotating elements as well as static structures. Using this process, turbomachines may be manufactured with rotating elements that have outer diameters of less than four inches in size, and/or blading heights of less than 0.1 inches. The rotating elements of the turbomachines are capable of rotating at speeds in excess of 150 feet per second. In addition, cooling features may be added internally to blading to facilitate cooling in high temperature operations. | 09-01-2011 |
| 20110217179 | TURBINE AIRFOIL FILLET COOLING SYSTEM - A cooling system for the fillet of a turbine blade is provided. The blade includes an airfoil transitioning to a platform having a flow path surface. The transition region is defined by a fillet. A cooling passage is formed in the platform and extends about at least a portion of the periphery of the airfoil. The cooling passage is located proximate to the flow path surface and is substantially aligned with at least a portion of the fillet. Coolant is delivered to the passage by a supply hole, which can reduce the temperature in the fillet region. As a result, thermal gradients in the fillet region can be minimized, which can reduce thermal stresses. An exhaust hole extends between the passage and the flow path surface of the platform. Thus, coolant discharged from the exhaust holes enters the flow path of the turbine. | 09-08-2011 |
| 20110217180 | GAS TURBINE BLADE, MANUFACTURING METHOD THEREFOR, AND GAS TURBINE USING TURBINE BLADE - Provided are gas turbine blades in which it is possible to simplify the formation of cooling channels provided inside the turbine blades while simultaneously avoiding loss of turbine blade strength and rigidity due to forming of the cooling channels. In a gas turbine blade, cooling channels provided in the interior thereof include a plurality of straight channel-like base-side elongated holes that extend in a longitudinal direction at a base side of the turbine blade, a plurality of straight channel-like tip-side elongated holes that extend in a longitudinal direction at a tip side of the turbine blade, and a plurality of communicating hollow portions that are interposed at connection portions between the two types of elongated holes to individually allow the two types of elongated holes to communicate with each other and that have larger cross-sectional areas than the channel cross-sectional areas of both elongated holes. In addition, the communicating hollow portions are formed so as to match the position of a platform portion of the turbine blade. | 09-08-2011 |
| 20110217181 | GAS TURBINE BLADE, MANUFACTURING METHOD THEREFOR, AND GAS TURBINE USING TURBINE BLADE - Provided are gas turbine blades in which it is possible to simplify the formation of cooling channels provided inside the turbine blades while simultaneously avoiding loss of turbine blade strength and rigidity due to forming of the cooling channels. In a gas turbine blade ( | 09-08-2011 |
| 20110229343 | APPARATUS FOR COOLING AN AIRFOIL - An apparatus for cooling an airfoil is provided. The airfoil includes an upper airfoil section, a lower airfoil section, at least one cooling passage, and a transition section. The at least one cooling passage is defined at least partially within the lower airfoil section. The at least one cooling passage is configured to flow a cooling medium therethrough, cooling at least a portion of the airfoil. The transition section is disposed between the upper airfoil section and the lower airfoil section and has an outer surface. The outer surface defines at least one cooling hole. The at least one cooling hole is fluidly connected to the at least one cooling passage. At least a portion of the cooling medium is exhausted through the at least one cooling hole. | 09-22-2011 |
| 20110229344 | Apparatus For Cooling A Bucket Assembly - A bucket assembly cooling apparatus is provided. The bucket assembly includes a platform, an airfoil, and a shank. The airfoil may extend radially outward from the platform. The shank may extend radially inward from the platform. The shank may include a pressure side sidewall, a suction side sidewall, an upstream sidewall, and a downstream sidewall. The sidewalls may at least partially define a cooling circuit. The cooling circuit may be configured to receive a cooling medium and provide the cooling medium to the airfoil. The upstream sidewall may at least partially define an interior cooling passage and at least partially define an exterior ingestion zone. The cooling passage may be configured to provide a portion of the cooling medium from the cooling circuit to the ingestion zone of an adjacent bucket assembly. | 09-22-2011 |
| 20110236220 | AIRFOIL COOLING HOLE FLAG REGION - An airfoil is provided and includes a body formed to define a substantially radially extending cooling hole therein, which is configured to be receptive of a supply of a coolant for removing heat from the body, and a flag region therein, which is fluidly communicative with the cooling hole and thereby configured to be receptive of a portion of the supply of the coolant such that the coolant portion is directed to form a vortex within the flag region to increase heat removal from the body beyond that provided by the coolant flow through the cooling hole. | 09-29-2011 |
| 20110236221 | Four-Wall Turbine Airfoil with Thermal Strain Control for Reduced Cycle Fatigue - A turbine airfoil ( | 09-29-2011 |
| 20110236222 | BLADE FOR A GAS TURBINE AND CASTING TECHNIQUE METHOD FOR PRODUCING SAME - A blade for a gas turbine has a leading edge and a trailing edge, and an interior cavity, which is delimited by internal surfaces, for guiding cooling air therethrough. A multiplicity of turbulators or pins, which are formed on the wall, are arranged in a distributed manner in the region of the trailing edge and project from the internal surfaces into the cavity, to improve the transfer of heat between the wall of the blade and the cooling air. An improvement of the internal cooling is achieved by the turbulators or pins extending into the cavity in a direction which can be freely selected within an angular range. | 09-29-2011 |
| 20110236223 | BLADE FOR A GAS TURBINE - A blade for a gas turbine includes an airfoil extending in a longitudinal direction and extending transversely to the longitudinal direction between a leading edge and a trailing edge. The airfoil has a pressure side, a suction side, and a slot-like cooling medium outlet extending along the trailing edge. The cooling medium outlet is configured to discharge cooling medium supplied from an inner space of the blade. A platform extends transversely to the longitudinal direction. An end of the airfoil merges into an underside of the platform and has a transition from the airfoil to the platform at the trailing edge of the airfoil that increases in thickness in a direction toward the underside of the platform. The cooling medium outlet extends into the platform so as to reduce an operating temperature in a region of the transition from the blade airfoil to the platform. | 09-29-2011 |
| 20110243755 | COOLED BLADE FOR A GAS TURBINE, METHOD FOR PRODUCING SUCH A BLADE, AND GAS TURBINE HAVING SUCH A BLADE - A blade for a gas turbine includes an airfoil extending in radial direction of the turbine or longitudinal direction of the blade, respectively, between a platform and a blade tip. The airfoil is bordered across the airfoil by a leading edge and a trailing edge and has a suction side and a pressure side. At the trailing edge a first cooling passage runs parallel to the trailing edge from the platform to the blade tip in the interior of the airfoil. The cooling passage is supplied with a cooling air flow from the platform side, and from which cooling air is discharged through a plurality of cooling holes arranged all over the blade. For such a blade the cooling is optimized by providing a first cooling passage, the passage area of which is tapered in radial direction by between 35% and 59%. | 10-06-2011 |
| 20110243756 | METHOD FOR PRODUCING A BLADE BY CASTING AND BLADE FOR A GAS TURBINE - A method is provided for producing a blade, by casting, for a gas turbine. The blade includes an elongate airfoil which extends in a blade longitudinal direction, merges into a blade root at the lower end, has a shroud segment at the blade tip and is pervaded by a single cooling air channel running in the blade longitudinal direction from the blade root to the blade tip. The method includes, during the casting of the blade, the blade material being fed exclusively from the blade root into the mold provided therefor, and the cooling air channel is formed during the casting of the blade by using a single core body, which is provided, at the blade tip, with a local casting cross section increasing element. | 10-06-2011 |
| 20110255989 | Cooling system of ring segment and gas turbine - In a cooling system of ring segment that cools a ring segment of a gas turbine, the segment body of the ring segment is constituted from a collision plate that has a small hole that blows out cooling air, a cooling space that is enclosed by the collision plate and the main body of the segment body; a first cavity that receives the cooling air from the cooling space; and a first cooling passage, of which one end communicates with the first cavity, and the other end blows out the cooling air from openings that are arranged in the side end portion into combustion gas; the openings of the first cooling passages being arranged so that the arrangement pitch of the openings becomes smaller or the opening area of the openings becomes larger on the upstream in the flow direction of the combustion gas than the openings on the downstream, and are arranged so that the arrangement pitch of the openings becomes larger or the opening area of the openings becomes smaller on the downstream in the flow direction of the combustion gas than the openings on the upstream. | 10-20-2011 |
| 20110255990 | BLADES - A rotor blade for a gas turbine engine has an aerofoil portion and a tip region. The tip region is at the radially outermost end of the blade. The radially outermost surface carries abrasive material (not shown) to interact with an abradable surface. The tip has a recess in which cooling air outlets are formed. The recess is open in a circumferential direction. This allows cooling air outlets to be formed without interference from the abrasive material, and inhibits any tendency for abrasion debris to collect in the recess and interfere with the flow of cooling air. | 10-20-2011 |
| 20110255991 | INTEGRALLY BLADED ROTOR DISK FOR A TURBINE - An integrally bladed rotor disk ( | 10-20-2011 |
| 20110268583 | AIRFOIL TRAILING EDGE AND METHOD OF MANUFACTURING THE SAME - A method for machining an airfoil including a plurality of internal cooling channels is provided. The method includes selectively removing a pressure side section proximate to a trailing edge of the airfoil to expose a portion of the plurality of internal cooling channels proximate to the trailing edge of the airfoil. The method also includes machining the exposed portion of the plurality of internal cooling channels to a predefined shape. | 11-03-2011 |
| 20110268584 | BLADES, TURBINE BLADE ASSEMBLIES, AND METHODS OF FORMING BLADES - Blades, turbine blade assemblies, and methods of forming blades are provided. The blade includes an airfoil including a convex suction side wall, a concave pressure side wall, a leading edge, a trailing edge, a root, and a tip, the convex suction side wall, the concave pressure side wall, and the tip each including interior surfaces that together define an internal cooling circuit, the airfoil including a single crystal superalloy, and a cladding layer disposed over the tip, the cladding layer including a zirconia grain stabilized platinum alloy. | 11-03-2011 |
| 20110274559 | Turbine Airfoil with Body Microcircuits Terminating in Platform - A turbine engine component includes a platform and one or more microcircuit cooling passages embedded within one or more walls of an airfoil portion of the component. Each microcircuit cooling passage terminates within the thickness of the platform so as to provide cooling to the initial 10% span of the airfoil portion. Each microcircuit cooling passage has an inlet for receiving cooling fluid, which inlet is also embedded within the platform. | 11-10-2011 |
| 20110280735 | VENTILATION OF A HIGH-PRESSURE TURBINE IN A TURBOMACHINE - A turbomachine high-pressure turbine including a rotor disc with upstream and downstream annular flanges separating a radially internal annular cavity in which the hub of the disc extends from two radially external annular cavities, of which one is upstream of the disc and receives a ventilation air flow and of which the other is downstream of the disc, the upstream flange of the disc including a mechanism connecting to the upstream eternal cavity and the internal cavity for ventilating the hub of the disc. | 11-17-2011 |
| 20110286857 | CERAMIC CORE TAPERED TRIP STRIPS - A core for creating an airfoil has a ceramic material that forms a body. The body has an outer dimension, a slot extending through the outer dimension and into the body for receiving an insert, the slot disposed at an angle to the outer dimension, and a trip strip having a first portion disposed in the outer dimension. The first portion is in register with the slot wherein a constant dimension such as minimum thickness is maintained between the trip strip and the slot along a length of the slot and wherein said first portion tapers towards said outer dimension. | 11-24-2011 |
| 20110299999 | MULTI-COMPONENT ASSEMBLY CASTING - Multi-component vane segment and method for forming the same. Assembly includes: positioning a pre-formed airfoil component ( | 12-08-2011 |
| 20110300000 | Component Having a Film Cooling Hole - Conventionally coated components with film cooling holes are known, comprising a diffuser, extending through the layers into the substrate. According to the invention, the component is embodied such that the whole diffuser is largely arranged in the layer. | 12-08-2011 |
| 20110305582 | Film Cooled Component Wall in a Turbine Engine - A component wall in a turbine engine. The component wall includes a substrate, a trench, and a plurality of cooling passages. The substrate has a first surface and a second surface opposed from the first surface. The trench is located in the second surface and is defined by a bottom surface between the first and second surfaces, a first sidewall, and a second sidewall spaced from the first sidewall. The first sidewall extends radially outwardly continuously from the bottom surface of the trench to the second surface. The first sidewall includes a plurality of first protuberances extending toward the second sidewall. The cooling passages extend through the substrate from the first surface to the bottom surface of the trench. Outlets of the cooling passages are arranged within the trench such that cooling air exiting the cooling passages is directed toward respective ones of the first protuberances. | 12-15-2011 |
| 20110305583 | COMPONENT WALL HAVING DIFFUSION SECTIONS FOR COOLING IN A TURBINE ENGINE - A film cooling structure formed in a component wall of a turbine engine and a method of making the film cooling structure. The film cooling structure includes a plurality of individual diffusion sections formed in the wall, each diffusions section including a single cooling passage for directing cooling air toward a protuberance of a wall defining the diffusion section. The film cooling structure may be formed with a masking template including apertures defining shapes of a plurality of to-be-formed diffusion sections in the wall. A masking material can be applied to the wall into the apertures in the masking template so as to block outlets of cooling passages exposed through the apertures. The masking template can be removed and a material may be applied on the outer surface of the wall such that the material defines the diffusion sections once the masking material is removed. | 12-15-2011 |
| 20110311369 | GAS TURBINE ENGINE COMPONENTS WITH COOLING HOLE TRENCHES - An engine component includes a body having an interior surface and an exterior surface; a cooling hole formed in the body and extending from the interior surface to the exterior surface; and a concave trench extending from the cooling hole at the exterior surface of the body in a downstream direction. | 12-22-2011 |
| 20110318191 | THERMALLY LOADED, COOLED COMPONENT - A thermally loaded, cooled component thermally coupled to a cooling system configured to receive a gaseous cooling introduced in a forced manner from outside that flows through the cooling system so as to absorb and transport heat away from the component as a result of thermal contact with the component. The component includes at least one Helmholtz resonator configured to improve thermal contact between the cooling medium and the component. | 12-29-2011 |
| 20120014808 | NEAR-WALL SERPENTINE COOLED TURBINE AIRFOIL - A serpentine coolant flow path ( | 01-19-2012 |
| 20120014809 | HIGH PRESSURE TURBINE VANE COOLING HOLE DISTRUBUTION - A turbine vane for a gas turbine engine with an airfoil portion including a perimeter wall having first, second, third and fourth sets of cooling holes defined therethrough, including the holes numbered HA-1 to HA-13, HB-1 to HB-13, PA-1 to PA-9, and SA-1 to SA-3, respectively, and located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y Cartesian coordinate values set forth in Table 3. | 01-19-2012 |
| 20120014810 | TURBINE VANE WITH DUSTING HOLE AT THE BASE OF THE BLADE - A cooled turbine vane for a turbine engine, that includes a blade mounted on a platform carried by a base, the blade including one or more cavities formed therein for cooling air circulation, the cavity extending along the trailing edge and being supplied with cooling air by a supply duct connecting an air intake located in a lower portion of the base and the cavity of the trailing edge by defining a bend within the base. The duct includes, on an axis substantially radial relative to the air intake a bell-shaped niche located under the platform, the niche being open at a top thereof via a dusting hole extending through the platform and being defined at a foot of the base by walls extending substantially radially from the platform to close the platform laterally. | 01-19-2012 |
| 20120034102 | BUCKET ASSEMBLY COOLING APPARATUS AND METHOD FOR FORMING THE BUCKET ASSEMBLY - A bucket assembly and a method for forming the bucket assembly are disclosed. The bucket assembly includes a platform, an airfoil, and a lower body portion. The platform defines a platform cooling circuit configured to flow cooling medium therethrough. The airfoil extends radially outward from the platform. The lower body portion extends radially inward from the platform. The lower body portion defines a root and a cooling passage extending from the root. The cooling passage is configured to flow cooling medium therethrough. The platform and lower body portion further include a ligament between the cooling passage and the platform cooling circuit. The ligament defines a bore hole extending through the ligament between the cooling passage and the platform cooling circuit. | 02-09-2012 |
| 20120057988 | ROTOR FOR A TURBOMACHINE - The invention relates to a rotor ( | 03-08-2012 |
| 20120063916 | TURBINE BLADE PLATFORM COOLING SYSTEMS - The present application provides a turbine blade cooling system. The turbine blade cooling system may include a first turbine blade with a first turbine blade platform having a cooling cavity in communication with a pressure side passage and a second turbine blade with a second turbine blade platform having a platform cooling cavity with a suction side passage. The pressure side passage of the first turbine blade platform is in communication with the suction side passage of the second turbine blade platform. | 03-15-2012 |
| 20120070305 | SHANK CAVITY AND COOLING HOLE - A turbine bucket is provided and includes a shank defining a cavity therein, which is connectible with a rotor such that wheelspace air having an initial pressure is permitted to flow into the cavity and a platform coupled to the shank and defining a cooling hole therein, the shank and the platform each further defining the cavity and the cooling hole, respectively, as being fluidly communicative with one another, such that the wheelspace air permitted to flow into the cavity is deliverable from the cavity and through the cooling hole at a second pressure greater than the initial pressure. | 03-22-2012 |
| 20120070306 | TURBINE COMPONENT COOLING CHANNEL MESH WITH INTERSECTION CHAMBERS | 03-22-2012 |
| 20120070307 | TURBINE BLADES, TURBINE ASSEMBLIES, AND METHODS OF MANUFACTURING TURBINE BLADES - A turbine blade includes a first side wall including a first tip edge, a second side wall opposite the first side wall and including a second tip edge, a tip wall between the first and second side walls, the tip wall recessed from the first tip edge of the first side wall and the second tip edge of the second side wall forming a coolant cavity, a tip recess cavity, a first parapet wall on the first side wall, and a second parapet wall on the second side wall, the coolant cavity defined by the tip wall, and the tip recess cavity defined by the tip wall, and the first and second parapet walls, a step formed between the first tip edge and the tip wall, a cooling hole through the first parapet wall, the step, and the tip wall, the cooling hole including an open and a closed channel section. | 03-22-2012 |
| 20120070308 | COOLED BLADE FOR A GAS TURBINE - A cooled blade for a gas turbine includes an airfoil section which extends in the radial direction of the turbine or in the longitudinal direction of the blade between a platform and a blade tip which is provided with a cap. The airfoil section is bounded transversely with respect to the longitudinal direction by a leading edge and a trailing edge and has a pressure face and a suction face. Cooling channels extend in a radial direction between the platform and the blade tip in an interior of the airfoil section. The cooling channels can be acted upon by a cooling air flow from the platform. The blade tip is cooled by first cooling holes for convection cooling provided on the pressure face of the blade, and second cooling holes for film cooling provided on the suction side of the blade, through the cap of the blade, in the blade tip from the cooling channels, and distributed over the blade width. | 03-22-2012 |
| 20120070309 | BLADE FOR A GAS TURBINE - A blade, for a gas turbine, includes a blade airfoil, having a shroud segment arranged on its upper end. The shroud segment together with shroud segments of other blades of a blade row forming an annular shroud which delimits hot gas passage of the gas turbine, and said shroud segment, on sides on which it adjoins adjacent shroud segments of the annular shroud, is provided with upwardly projecting side rails which extend along a side edge, to improve sealing to the hot gas passage. The side rails include rail-parallel or essentially rail-parallel, upwardly open slots through which cooling air, which is introduced via the shroud segment from an interior of the blade airfoil, discharges into the space above the shroud segment. | 03-22-2012 |
| 20120070310 | AXIAL TURBOMACHINE ROTOR HAVING BLADE COOLING - An axial turbomachine rotor is provided. The rotor includes a rotor disk and a rotor blade ring, which includes a plurality of rotor blades, each of which include a blade root, with which the rotor blade is fixed radially outward on the rotor disk, wherein the blade root is engaged with the rotor disk at the outer edge of the rotor disk in a form-closed manner in such a way that during operation of the rotor, a gap is formed between the rotor blade and the rotor disk at a predetermined surface area of the rotor disk, in which area, a plurality of impingement cooling openings is arranged, through which a cooling medium may flow from the interior of the rotor disk into the gap whereby the rotor blade is cooled using the cooling medium by means of impingement cooling. | 03-22-2012 |
| 20120076665 | COOLED TURBINE BLADES FOR A GAS-TURBINE ENGINE - The present invention relates to a cooled turbine blade for a gas-turbine engine having at least one cooling duct ( | 03-29-2012 |
| 20120082563 | Cooed IBR for a micro-turbine - A micro gas turbine engine in which the turbine rotor blades are formed as an integral bladed rotor with cooling air passages formed within the blades and the rotor disk by an EDM process. an adjacent stator vane includes an air riding seal with an air cushion supplied through the vanes to provide cooling, and where the air cushion is then passed into the turbine blades and rotor disk to provide cooling for the turbine blades. With cooling of the turbine blades, higher turbine inlet temperatures for micro gas turbine engines can be produced. | 04-05-2012 |
| 20120082564 | APPARATUS AND METHODS FOR COOLING PLATFORM REGIONS OF TURBINE ROTOR BLADES - A platform cooling arrangement in a turbine rotor blade having a platform that includes an interior cooling passage formed therein. The platform cooling arrangement may include: a main plenum residing just inboard of the planar topside and extending from an aft position to a forward position within one of the pressure side and the suction side of the platform, the main plenum having a longitudinal axis that is approximately parallel to the planar topside; a supply plenum that extends between the main plenum and the interior cooling passage; and a plurality of cooling apertures, each cooling aperture extending from one of the pressure side and the suction side slashface to a connection with the main plenum. | 04-05-2012 |
| 20120082565 | APPARATUS AND METHODS FOR COOLING PLATFORM REGIONS OF TURBINE ROTOR BLADES - 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. | 04-05-2012 |
| 20120082566 | APPARATUS AND METHODS FOR COOLING PLATFORM REGIONS OF TURBINE ROTOR BLADES - 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. | 04-05-2012 |
| 20120082567 | COOLED ROTOR BLADE - A cooled turbine rotor blade for a gas turbine engine is provided. The engine has an annular flow path for conducting working fluid though the engine. The blade has an aerofoil section for extending across the annular flow path. The blade further has a root portion radially inward of the aerofoil section for joining the blade to a rotor disc of the engine. The blade further has a platform between the aerofoil section and the root portion. The platform extends laterally relative to the radial direction of the engine to form an inner boundary of the annular flow path and to provide a rear overhang portion which projects in use towards a corresponding platform of a downstream nozzle guide vane. The platform contains at least one internal elongate plenum chamber for receiving cooling air. The longitudinal axis of the plenum chamber is substantially aligned with the circumferential direction of the engine. The plenum chamber supplies the cooling air to a plurality of exit holes formed in the external surface of the rear overhang portion to cool that portion. | 04-05-2012 |
| 20120082568 | TURBINE DISC COOLING ARRANGEMENT - A cooling arrangement is provided for a turbine disc of a gas turbine engine. The turbine disc has a plurality of circumferentially spaced disc posts forming fixtures therebetween for a row of turbine blades. Each turbine blade has an attachment formation which engages at a respective fixture, a platform radially outwardly of the attachment formation such that the adjacent platforms of the row form an inner endwall for the working gas annulus of the engine, and an aerofoil which extends radially outwardly from the platform. A respective cavity is formed between an exposed radially outer surface of each disc post and the inner endwall. The cooling arrangement has at each disc post, a cooling plate located in the respective cavity and spaced radially outwardly from the exposed outer surface of the disc post to form a cooling channel between the cooling plate and the exposed outer surface. | 04-05-2012 |
| 20120087803 | CURVED FILM COOLING HOLES FOR TURBINE AIRFOIL AND RELATED METHOD - A turbine bucket includes an airfoil portion at one end thereof; a root portion at an opposite end thereof; a platform portion between the airfoil portion and the root portion; at least one internal cavity within or radially inward of the platform portion having at least one film cooling hole extending between the at least one cavity and an external surface of the platform portion. The film cooling hole is curved along a length dimension of the film cooling hole. | 04-12-2012 |
| 20120100008 | ANNULAR FLOW CHANNEL SECTION FOR A TURBOMACHINE - An annular flow channel section for a turbomachine is provided. The annular flow channel includes a guide vane ring having a number of guide vanes which are arranged next to each other in the circumferential direction, each guide vane including a vane base, a platform and an airfoil that projects into the flow channel in a radiating pattern. The flow channel is delimited on the platform side by shielding elements, each of which being disposed between two immediately adjacent airfoils, wherein the shielding elements are arranged on the platforms for creating a particularly space-saving flow channel section while creating a gap, and impingement-cooling openings are provided in the platform. | 04-26-2012 |
| 20120107134 | APPARATUS AND METHODS FOR COOLING PLATFORM REGIONS OF TURBINE ROTOR BLADES - 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. | 05-03-2012 |
| 20120107135 | APPARATUS, SYSTEMS AND METHODS FOR COOLING THE PLATFORM REGION OF TURBINE ROTOR BLADES - 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. The platform cooling arrangement may include a linear plenum residing just inboard of the planar topside and linearly extending through the platform from either the pressure side slashface or the suction side slashface to a connection with the interior cooling passage, the linear plenum having a longitudinal axis that is approximately parallel to the planar topside; and a plurality of cooling apertures linearly extending from a topside outlet formed on the topside of the platform to a connection with the linear plenum, wherein the cooling apertures are configured such that each forms an acute angle with the topside of the platform. | 05-03-2012 |
| 20120107136 | SEALING PLATE AND ROTOR BLADE SYSTEM - A rotor blade system, for example, of a gas turbine is provided. The rotor blade system includes a plurality of rotor blades which are arranged annularly on a rotor disk. A plurality of sealing plates are arranged on a side surface of the rotor disk. An individual sealing plate is formed from a plurality of metal sheets, wherein two of the metal sheets are arranged opposite each other a distance apart and parallel to a plane of the sealing plate, forming a gap for guiding of cooling air. | 05-03-2012 |
| 20120121435 | ROTOR FOR A TURBO MACHINE - The invention relates to a rotor ( | 05-17-2012 |
| 20120121436 | ROTOR FOR A TURBO MACHINE - The invention relates to a rotor ( | 05-17-2012 |
| 20120121437 | ROTOR FOR A TURBO MACHINE - The invention relates to a rotor ( | 05-17-2012 |
| 20120128504 | Rotor section for a rotor of a turbomachine, and rotor blade for a turbomachine - A rotor section for a rotor of a gas turbine is provided. The rotor section includes a rotor disk with rotor blades inserted in retaining grooves thereupon, the rotor blades being secured against displacement along the retaining grooves by means of a sealing element which is arranged on the end face. In order to disclose a reliable construction which can be designed in a straightforward and simple manner for the circumferential fixing of the sealing element, each sealing element is secured by means of a blocking element in each case which in this case engages in a hole which is arranged in the rotor blade root on the end face. | 05-24-2012 |
| 20120134845 | BLADE FOR A GAS TURBINE, METHOD FOR MANUFACTURING A TURBINE BLADE, AND GAS TURBINE WITH A BLADE | 05-31-2012 |
| 20120134846 | WIND TURBINE GENERATOR - Provided is a wind turbine generator that prevents an increase of weight of a wind turbine blade and a reduction in ventilation efficiency and that realizes good ventilation cooling of a rotor head with a simple structure. The wind turbine generator includes: a hollow interior space formed in a wind turbine blade; an exhaust path provided through a wind-turbine-blade forming member at a tip end of the wind turbine blade or in a vicinity of the tip end, so as to allow the interior space to communicate with an outside of the wind turbine blade; and an opening that allows the interior space to communicate with a rotor head main body. Ventilation cooling inside the rotor head main body is performed by using a pressure difference between the tip end of the wind turbine blade and the rotor head main body, produced by a rotation of the wind turbine blade. | 05-31-2012 |
| 20120156055 | APPARATUS AND METHODS FOR COOLING PLATFORM REGIONS OF TURBINE ROTOR BLADES - 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. | 06-21-2012 |
| 20120163992 | DRILL TO FLOW MINI CORE - A core for forming a cooling microcircuit has at least one row of metering/tripping features configured to form at least one row of protrusions in the cooling microcircuit, a plurality of teardrop features configured to form forming a plurality of fluid passageways in the cooling microcircuit, and a terminal edge. The plurality of teardrop features includes a central teardrop feature having a trailing edge which is spaced from the terminal edge and a first teardrop feature located on a first side of and spaced from the central teardrop feature. The first teardrop feature has a longitudinal axis and is non-symmetrical about the longitudinal axis. A process of using the core and a turbine engine component formed thereby are described. | 06-28-2012 |
| 20120163993 | LEADING EDGE AIRFOIL-TO-PLATFORM FILLET COOLING TUBE - A turbine engine component includes an airfoil portion having a leading edge, a platform, a leading edge airfoil to platform fillet, and a cooling tube located within said fillet. The cooling tube has a flared entrance end and a flared exit end. | 06-28-2012 |
| 20120163994 | GAS TURBINE ENGINE AND AIRFOIL - One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique turbine engine airfoil. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and airfoils. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. | 06-28-2012 |
| 20120163995 | TURBINE BLADE - 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. | 06-28-2012 |
| 20120171046 | APPARATUS AND METHODS FOR COOLING PLATFORM REGIONS OF TURBINE ROTOR BLADES - 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. | 07-05-2012 |
| 20120171047 | TURBOMACHINE AIRFOIL COMPONENT AND COOLING METHOD THEREFOR - An airfoil component for use in a turbomachine, and method of promoting the heat transfer characteristics within the component. The component includes an airfoil portion having a span-wise direction delimited by an airfoil root and airfoil tip, and a chord-wise direction delimited by leading and trailing edges. A chamber within the airfoil portion contains a permeable foam member. The chamber is fluidically connected to a cooling fluid source and to a cooling hole through first and second passages, respectively, within the airfoil portion. The chamber is located relative to the first and second passages so as to be offset in the chord-wise direction therefrom so that cooling fluid entering the chamber through the first passage is diverted by the foam member in the chord-wise direction before exiting the airfoil portion through the cooling hole. | 07-05-2012 |
| 20120183412 | CURVED COOLING PASSAGES FOR A TURBINE COMPONENT - A turbine component having a curved cooling passage is disclosed. The 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 and a suction side extending between a leading edge and a trailing edge. An airfoil cooling circuit may be at least partially disposed within the airfoil and may be configured to direct a cooling medium through the airfoil. The curved cooling passage may generally be in flow communication with the airfoil cooling circuit such that the cooling medium flowing through the airfoil cooling circuit may be directed into the cooling passage. Additionally, the curved cooling passage may generally extend lengthwise within the airfoil between the leading and trialing edges along at least a portion of one of the pressure side and the suction side of the airfoil. | 07-19-2012 |
| 20120201695 | TURBINE BLADE SQUEALER TIP RAIL WITH FENCE MEMBERS - A turbine blade includes an airfoil, a blade tip section, a squealer tip rail, and a plurality of chordally spaced fence members. The blade tip section includes a blade tip floor located at an end of the airfoil distal from the root. The blade tip floor includes a pressure side and a suction side joined together at chordally spaced apart leading and trailing edges of the airfoil. The squealer tip rail extends radially outwardly from the blade tip floor adjacent to the suction side and extends from a first location adjacent to the airfoil trailing edge to a second location adjacent to the airfoil leading edge. The fence members are located between the airfoil leading and trailing edges and extend radially outwardly from the blade tip floor and axially from the squealer tip rail toward the pressure side. | 08-09-2012 |
| 20120207614 | INTEGRATED AXIAL AND TANGENTIAL SERPENTINE COOLING CIRCUIT IN A TURBINE AIRFOIL - A continuous serpentine cooling circuit forming a progression of radial passages ( | 08-16-2012 |
| 20120207615 | Cooling of a Gas Turbine Component Designed as a Rotor Disk or Turbine Blade - A gas turbine component for example a turbine blade or a rotor disk is provided. In order to extend the service life of the corresponding component by reducing the thermally or mechanically induced stress concentration in the direct surroundings of a duct opening onto a surface, at least one groove-like recess is provided near the effective zone of the opening. | 08-16-2012 |
| 20120207616 | Castings, Casting Cores, and Methods - The pattern has a pattern material and a casting core combination. The pattern material has an airfoil. The casting core combination is at least partially embedded in the pattern material. The casting core combination comprises a metallic casting core and at least one additional casting core. The metallic casting core has opposite first and second faces. The metallic core and at least one additional casting core extend spanwise into the airfoil of the pattern material. In at least a portion of the pattern material outside the airfoil of the pattern material, the metallic casting core is bent transverse to the spanwise direction so as to at least partially surround an adjacent portion of the at least one additional casting core. | 08-16-2012 |
| 20120224975 | PROCESS FOR THE VAPOR PHASE ALUMINIZATION OF A TURBOMACHINE METAL PART AND DONOR LINER AND TURBOMACHINE VANE COMPRISING SUCH A LINER - An aluminization process by vapor phase deposition for high-temperature oxidation protection of a metal turbomachine part. The part including a cavity into which a metal component is introduced and assembled from an opening in the part. A halide is formed by reaction between a halogen and a metal donor containing aluminum, then the halide is transported by a carrier gas in order to come into contact with the metal part, wherein the metal component has first, before the implementation of the process, been surface-enriched with aluminum in order to serve as an aluminum donor. | 09-06-2012 |
| 20120230838 | TURBINE BLADE AND GAS TURBINE - Provided is a turbine blade including: a platform; a blade body including a cooling flow channel including a meandering serpentine cooling flow channel; a fillet portion provided in a joint surface between the blade body and the platform; and a base portion including a cooling flow channel communicated with the serpentine cooling flow channel. The cooling flow channel includes a bypass flow channel that is branched off from a high-pressure part of the cooling flow channel, passes through along the inside of the fillet portion, and is connected to a low-pressure part of the cooling flow channel. | 09-13-2012 |
| 20120237359 | COOLED PUSHER PROPELLER SYSTEM - A propulsion system and method includes an annular exhaust nozzle about an axis radially outboard of the annular cooling flow nozzle and ejecting an exhaust flow through an annular exhaust nozzle about an axis radially outboard of the annular cooling flow nozzle. | 09-20-2012 |
| 20120251331 | Turbine Blade Platform Undercut - A system and method of extending the useable life of a gas turbine blade is disclosed in which the gas turbine blade includes an undercut configuration designed to relieve mechanical and thermal stress imparted into the pedestal region of the airfoil trailing edge. The embodiments of the present invention include turbine blade configurations having different trailing edge undercut configurations as well as additional cooling supplied to the internal passages of the trailing edge region of the turbine blade. | 10-04-2012 |
| 20120263603 | TURBINE BLADE AND GAS TURBINE - A turbine blade includes: a base portion fixed to a rotor; a platform that is fixed to the base portion and includes a cooling flow channel; a blade shape portion that extends from the platform to an outer side in a radial direction; and a fillet portion provided in a joint surface between the blade shape portion and the platform, and the turbine blade further includes: a main pipe that is branched off from the cooling flow channel and has an opening at a side end portion of the platform; and a branch pipe that is branched off from the main pipe, extends along the fillet portion so as to be close thereto on an inner side in the radial direction, and includes a film cooling hole opened on a surface of the platform. | 10-18-2012 |
| 20120269647 | COOLED AIRFOIL IN A TURBINE ENGINE - An airfoil in a gas turbine engine includes an outer wall and an inner wall. The outer wall includes a leading edge, a trailing edge opposed from the leading edge in a chordal direction, a pressure side, and a suction side. The inner wall is coupled to the outer wall at a single chordal location and includes portions spaced from the pressure and suction sides of the outer wall so as to form first and second gaps between the inner wall and the respective pressure and suction sides. The inner wall defines a chamber therein and includes openings that provide fluid communication between the respective gaps and the chamber. The gaps receive cooling fluid that provides cooling to the outer wall as it flows through the gaps. The cooling fluid, after traversing at least substantial portions of the gaps, passes into the chamber through the openings in the inner wall. | 10-25-2012 |
| 20120269648 | SERPENTINE COOLING CIRCUIT WITH T-SHAPED PARTITIONS IN A TURBINE AIRFOIL - A serpentine cooling circuit (AFT) in a turbine airfoil ( | 10-25-2012 |
| 20120269649 | TURBINE BLADE WITH IMPROVED TRAILING EDGE COOLING - A turbine blade ( | 10-25-2012 |
| 20120269650 | TURBINE WHEEL FOR A TURBINE ENGINE - A turbine wheel for a turbine engine, comprising a disk carrying blades, each having a platform carrying an airfoil and connected by a tang to a root, and sealing and damping sheets housed in the inter-tang spaces, the platforms including projections on their radially internal faces against which the sheets bear radially in operation, in order to define radial clearance and create at least one space between the sheets and the platforms, and the sheets including holes for feeding air to the or each space. | 10-25-2012 |
| 20120275926 | WELDED ROTOR OF A GAS TURBINE ENGINE COMPRESSOR | 11-01-2012 |
| 20120282107 | TURBINE AIRFOIL COOLING SYSTEM WITH HIGH DENSITY SECTION OF ENDWALL COOLING CHANNELS - A cooling system for a turbine airfoil of a turbine engine having a trailing edge cooling region formed from endwall cooling channels having a higher density of cooling channels than other areas in order to cool the material forming the intersection between the trailing edge of the airfoil and the endwall to prevent premature cracking. The increased density of cooling channels in the endwall at the trailing edge forms a heat sink that draws heat from the airfoil, thereby lowering the temperature of the airfoil and increasing the useful life of the airfoil. | 11-08-2012 |
| 20120282108 | TURBINE BLADE WITH CHAMFERED SQUEALER TIP AND CONVECTIVE COOLING HOLES - A squealer tip formed from a pressure side rib and a suction side rib extending radially outward from a tip of the turbine blade is disclosed. The pressure and suction side ribs may be positioned along the pressure side and the suction side of the turbine blade, respectively. The pressure and suction side ribs may include chamfered leading edges with film cooling holes having exhaust outlets positioned therein. The film cooling holes may be configured to be diffusion cooling holes with one or more tapered sections for reducing the velocity of cooling fluids and increasing the size of the convective surfaces. | 11-08-2012 |
| 20120282109 | Blade, Integrally Bladed Rotor Base Body and Turbomachine - A blade for a turbomachine is disclosed. At least one cooling channel is configured in the blade neck, whose inlet is disposed near a platform projection on the high-pressure side and whose outlet is disposed in the region of a platform projection on the low-pressure side. An integrally bladed rotor base body having a plurality of these types of blades as well as turbomachine with such a rotor base body is also disclosed. | 11-08-2012 |
| 20120282110 | INNER VENTILATION BLADE - A blade, for example of a high-pressure turbine, is designed with an internal partition ( | 11-08-2012 |
| 20120301319 | Curved Passages for a Turbine Component - 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. | 11-29-2012 |
| 20120308399 | TURBINE NOZZLE SLASHFACE COOLING HOLES - A turbine vane or blade segment includes at least one airfoil extending radially outwardly from a radially inner band. The radially inner band is formed with substantially axially-extending side edges, and the at least one airfoil is formed with a leading edge, a trailing edge, a pressure side, a suction side, and an internal cooling plenum. The pressure side faces one of the substantially axially-extending side edges and the leading edge is located proximate to that one edge. A plurality of cooling passages are formed in the radially inner band in fluid communication with the internal plenum and exiting the inner band via a plurality of exit holes in the one of the substantially axially-extending side edges. The plurality of exit holes are confined to a region along the one of the substantially axially-extending side edges where static pressure P | 12-06-2012 |
| 20120315150 | TURBINE ROTOR BLADE - A concave (a relief portion) | 12-13-2012 |
| 20120328450 | COOLING SYSTEM FOR TURBINE AIRFOIL INCLUDING ICE-CREAM-CONE-SHAPED PEDESTALS - A turbine airfoil comprises a wall portion, a cooling channel, a plurality of trip strips and a plurality of pedestals. The wall portion comprises a leading edge, a trailing edge, a pressure side and a suction side. The cooling channel is for receiving cooling air and extends radially through an interior of the wall portion between the pressure side and the suction side. The plurality of trip strips line the wall portion inside the cooling channel along the pressure side and the suction side. Each of the pedestals is an elongate, tapered pedestal having a curved leading edge. The plurality of pedestals is interposed within the trip strips and connects the pressure side with the suction side. | 12-27-2012 |
| 20120328451 | PLATFORM COOLING PASSAGES AND METHODS FOR CREATING PLATFORM COOLING PASSAGES IN TURBINE ROTOR BLADES - 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. | 12-27-2012 |
| 20130004331 | TURBINE BLADE OR VANE WITH SEPARATE ENDWALL - A turbine engine airfoil structure including an airfoil adapted to be supported to extend across a gas passage for a hot working gas in a turbine engine. The airfoil structure further includes a platform structure located at one end of the airfoil and positioned at a location forming a boundary of the gas passage. The platform structure includes a platform member including a gas side surface extending generally perpendicular from the airfoil at a junction with the airfoil, and providing a structural connection to the airfoil. The platform structure further includes a separately formed platform cover attached to the platform member at the gas side surface. The platform cover extends from a location adjacent to one of the sidewalls of the airfoil, and includes an outer surface located for contact with the hot working gas passing through the gas path. | 01-03-2013 |
| 20130004332 | GAS TURBINE BLADE AND METHOD FOR PRODUCING A BLADE | 01-03-2013 |
| 20130039777 | AIRFOIL INCLUDING TRENCH WITH CONTOURED SURFACE - A airfoil comprises a wall, a cooling channel, a trench and a plurality of leading edge cooling holes. The wall has a leading edge, a trailing edge, a pressure side, a suction side, an outer diameter end and an inner diameter end to define an interior. The cooling channel extends radially through the interior of the wall between the pressure side and the suction side and along the leading edge. The trench extends radially along an exterior of the wall at the leading edge and is recessed axially into the leading edge to form a back wall. The back wall is contoured to include at least one undulation. The plurality of leading edge cooling holes extends through the back wall of the trench to connect the interior of the wall at the cooling channel to the exterior. | 02-14-2013 |
| 20130045111 | TURBINE BLADE COOLING SYSTEM WITH BIFURCATED MID-CHORD COOLING CHAMBER - A cooling system for a turbine blade of a turbine engine having a bifurcated mid-chord cooling chamber for reducing the temperature of the blade. The bifurcated mid-chord cooling chamber may be formed from a pressure side serpentine cooling channel and a suction side serpentine cooling channel with cooling fluids passing through the pressure side serpentine cooling channel in a direction from the trailing edge toward the leading edge and in an opposite direction through the suction side serpentine cooling channel. The pressure side and suction side serpentine cooling channels may flow counter to each other, thereby yielding a more uniform temperature distribution than conventional serpentine cooling channels. | 02-21-2013 |
| 20130064680 | TURBINE ENDWALL WITH GROOVED RECESS CAVITY - A vane assembly for a gas turbine engine including an endwall and an airfoil extending from the endwall. An inner rail extends radially inwardly of the endwall, and an overhang portion extends axially from a location of the inner rail to a downstream edge. A recess cavity is defined in the overhang portion between the inner rail and the downstream edge. The recess cavity extends radially into the overhang portion and defines a cavity surface. A plurality of grooves extend radially into the cavity surface and have an elongated dimension extending in a direction from the inner rail toward the downstream edge. A plurality of cooling passages extend axially through the overhang portion, and are located between the grooves. | 03-14-2013 |
| 20130064681 | TRAILING EDGE COOLING SYSTEM IN A TURBINE AIRFOIL ASSEMBLY - An airfoil in a gas turbine engine includes an outer wall, a cooling fluid cavity, and a plurality of cooling fluid passages. The outer wall has a leading edge, a trailing edge, a pressure side, a suction side, and radially inner and outer ends. The cooling fluid cavity is defined in the outer wall, extends generally radially between the inner and outer ends of the outer wall, and receives cooling fluid for cooling the outer wall. The cooling fluid passages are in fluid communication with the cooling fluid cavity and include zigzagged passages that include alternating angled sections, each section having both a radial component and a chordal component. The cooling fluid passages extend from the cooling fluid cavity toward the trailing edge of the outer wall and receive cooling fluid from the cooling fluid cavity for cooling the outer wall near the trailing edge. | 03-14-2013 |
| 20130071255 | Gas Turbine Blade - A gas turbine blade having a film cooling structure can reduce stress and strain that occur around the cooling holes of the film cooling structure. For example, in the gas turbine blade, a plurality of cooling holes thoroughly connected to the cooling pass formed inside the gas turbine blade are arranged in the span direction in the leading edge portion of the gas turbine blade, and the direction of the longitudinal axis of the cooling holes is made identical to the direction of principal strain occurring in the leading edge portion of the turbine blade within a range of 15 degrees. | 03-21-2013 |
| 20130078110 | OFFSET COUNTERBORE FOR AIRFOIL COOLING HOLE - 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. | 03-28-2013 |
| 20130084191 | TURBINE BLADE WITH IMPINGEMENT CAVITY COOLING INCLUDING PIN FINS - A turbine blade including an airfoil having a cavity defined between pressure and suction side walls, and first and second cooling circuits located within the cavity. The second cooling circuit includes a first upstream impingement cavity, a second upstream impingement cavity and a discharge impingement cavity. A plurality of pin fins are located within each of the impingement cavities, extending between the pressure and suction side walls to provide cooling to the pressure and suction side walls, and to restrict flow through the second cooling circuit and to direct an increased flow of cooling fluid into the first cooling circuit. | 04-04-2013 |
| 20130089434 | METHODS AND SYSTEMS FOR USE IN REGULATING A TEMPERATURE OF COMPONENTS - 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. | 04-11-2013 |
| 20130094971 | HOT GAS PATH COMPONENT FOR TURBINE SYSTEM - A hot gas path component for a turbine system is disclosed. The hot gas path component includes a shell having an exterior surface and an interior surface. The hot gas path component further includes a porous medium having an exterior surface and an interior surface, the exterior surface positioned adjacent to the interior surface of the shell. The porous medium is configured for flowing a cooling medium therethrough. | 04-18-2013 |
| 20130101436 | FAN ROTOR WITH COOLING HOLES - A disc for a fan rotor (with a pilot to connect to a rotating shaft, a hub and a plurality of blades) includes a flat circular portion connecting to the pilot at an inner edge and to the hub at an outer edge; a plurality of first circular cooling holes of a first diameter located around the inner edge of the disc; and a plurality of second circular cooling holes of a second diameter located around the outer edge of the disc, wherein the second diameter is larger than the first diameter. | 04-25-2013 |
| 20130108469 | AIRFOIL AND METHOD OF FABRICATING THE SAME | 05-02-2013 |
| 20130108470 | ROTOR BLADE WITH BONDED COVER | 05-02-2013 |
| 20130115100 | BUCKET ASSEMBLY FOR TURBINE SYSTEM - A bucket assembly for a turbine system is disclosed. The bucket assembly includes a main body having an exterior surface and defining a main cooling circuit, and a platform surrounding the main body and at least partially defining a platform cooling circuit. The bucket assembly further includes a passage defined in the main body extending from the exterior surface, the passage connecting the main cooling circuit and the platform cooling circuit, and a plug at least partially disposed in the passage. The plug includes a head and a plunger. The head is configured for preventing a flow through at least a portion of the passage. The plunger has a continuous exterior surface and is configured for allowing a flow between the main cooling circuit and the platform cooling circuit. | 05-09-2013 |
| 20130115101 | BUCKET ASSEMBLY FOR TURBINE SYSTEM - A bucket assembly for a turbine system is disclosed. The bucket assembly includes a main body having an exterior surface and defining a main cooling circuit, and a platform surrounding the main body and at least partially defining a platform cooling circuit. The platform includes a forward portion and an aft portion each extending between a pressure side slash face and a suction side slash face. The platform further includes a forward face, an aft face, and a top face. The bucket assembly further includes a passage defined in the aft portion of the platform. The passage is in fluid communication with one of the main cooling circuit or the platform cooling circuit. | 05-09-2013 |
| 20130115102 | BUCKET ASSEMBLY FOR TURBINE SYSTEM - A bucket assembly for a turbine system is disclosed. The bucket assembly includes a main body having an exterior surface and defining a main cooling circuit, and a platform surrounding the main body and at least partially defining a platform cooling circuit. The platform includes a forward portion and an aft portion each extending between a pressure side slash face and a suction side slash face. The platform further includes a forward face, an aft face, and a top face. The bucket assembly further includes a passage defined in the platform generally between the platform cooling circuit and the pressure side slash face and in fluid communication with one of the main cooling circuit or the platform cooling circuit. | 05-09-2013 |
| 20130115103 | FILM HOLE TRENCH - An article is disclosed that comprises a thermal material having a first surface and a second surface. The thermal material defines a film hole between the first surface and the second surface, and the film hole includes a metering portion adjacent the first surface and a diffuser portion adjacent the second surface. The metering portion defines a metering hole axis, and the diffuser portion defines a trench. The trench extends substantially parallel to a metering hole axis. | 05-09-2013 |
| 20130142665 | HYBRID VAPOR AND FILM COOLED TURBINE BLADE - A cooling system for cooling a fluid reaction apparatus of a gas turbine engine includes a vapor cooling subsystem and a film cooling subsystem. The vapor cooling subsystem has a vaporization section and a condenser section for cooling a portion of the fluid reaction apparatus. The condenser section is cooled by a fluid. The film cooling subsystem is configured for cooling a portion of the fluid reaction apparatus by discharging fluid out of openings defined in the fluid reaction apparatus. At least a portion of the fluid used to cool the condenser section of the vapor cooling subsystem is discharged out of the openings of the film cooling subsystem. | 06-06-2013 |
| 20130142666 | TURBINE BLADE INCORPORATING TRAILING EDGE COOLING DESIGN - A turbine blade ( | 06-06-2013 |
| 20130142667 | TURBINE BLADE AND GAS TURBINE HAVING THE SAME - A turbine blade of the present invention includes a blade body mounted on a rotor body so as to extend outward from the rotor body in a radial direction of the rotor body and a chip shroud fixed to the outside of the blade body in the radial direction. A cooling passage which extends in the radial direction of the rotor body and in which a cooling medium circulates is formed in the blade body. The chip shroud includes a shroud body where a recess opened to the outside in the radial direction and communicating with the cooling passage is formed on an outer peripheral end face, and a plug that includes a plurality of plug pieces closing an opening of the recess in cooperation with each other by being inserted into mounting grooves formed on side surfaces of the recess. | 06-06-2013 |
| 20130142668 | COOLED TURBINE BLADE FOR GAS TURBINE ENGINE - A cooled turbine blade for a gas turbine engine including a pressure surface wall, a suction surface wall and a distal wall connecting the pressure surface wall and the suction surface wall, arranged so as to create in the region of the distal end of the blade at least one external cavity forming a bathtub-shaped cavity and at least one internal cavity separated by the distal wall, the blade having at least one opening for the introduction of a flow of cooling air into the external cavity, wherein, on the one hand, at least one part of the distal wall is inclined relative to the verticals of the pressure surface wall and, on the other hand, the opening is created in the vicinity of the distal wall so that the flow of cooling air is directed towards the distal end of the pressure surface wall. | 06-06-2013 |
| 20130149169 | COMPONENT HAVING COOLING CHANNEL WITH HOURGLASS CROSS SECTION - A cooling channel ( | 06-13-2013 |
| 20130156600 | COMPONENTS WITH MICROCHANNEL COOLING - A component includes a substrate having an outer surface, an inner surface and a tip. The inner surface defines at least one hollow, interior space. The outer surface defines one or more grooves, where each groove extends at least partially along the outer surface of the substrate and has a base. The component further includes a coating disposed over at least a portion of the outer surface of the substrate. The coating includes at least a structural coating that extends over the groove(s), such that the groove(s) and the structural coating together define one or more channels for cooling the component. The tip comprises a tip cap enclosing the hollow, interior space(s), and a tip rim disposed at a radially outer end of the substrate. The tip rim at least partially defines at least one discharge channel in fluid communication with at least one cooling channel. | 06-20-2013 |
| 20130156601 | GAS TURBINE ENGINE AIRFOIL COOLING CIRCUIT - An airfoil for a gas turbine engine includes an airfoil body and a cooling circuit defined within the airfoil body. The cooling circuit includes at least a first cavity in fluid communication with a second cavity. A first portion of the first cavity extends between an outer diameter and an inner diameter of the airfoil body and a second portion of the first cavity extends across a space between a leading edge and a trailing edge of the airfoil body. | 06-20-2013 |
| 20130156602 | FILM COOLED TURBINE COMPONENT - Film cooling of turbine component surfaces, such as high-lift airfoil surfaces, is achieved by a flow of cooling air through film cooling holes of generally constant cross-sectional area, which extend from the surfaces to a radial cooling passage within the airfoil. The film cooling holes are angularly offset from that portion of the radial cooling passage immediately upstream of the film cooling holes by an acute angle which effects a radial reversal of the flow of cooling air into the film cooling holes from the radial cooling passage to reduce the momentum of airflow through the film cooling holes to reduce separation of cooling air film from the surface. | 06-20-2013 |
| 20130156603 | AEROFOIL BLADE OR VANE - An aerofoil blade or vane for the turbine of a gas turbine engine is provided. The blade or vane includes an aerofoil portion which, in use, extends radially across a working gas annulus of the engine. A coolant inlet is formed at an end of the aerofoil portion for the entry of a cooling air flow into the portion. A corresponding coolant exhaust is formed at the trailing edge of the aerofoil portion for spent cooling air to flow from the portion. A passage within the aerofoil portion connects the inlet to the exhaust. A fence within the aerofoil portion extends radially and forwardly from a start position at the end of the aerofoil portion adjacent the trailing edge to an end position. The passage forms a loop which extends along one side of the fence, wraps around the end position, and extends along the other side of the fence. | 06-20-2013 |
| 20130171003 | TURBINE ROTOR BLADE PLATFORM COOLING - A platform cooling arrangement for a turbine rotor blade having a platform at an interface between an airfoil and a root, the root including attachment means and a shank, wherein the platform comprises a suction side that includes a topside extending from an airfoil base to a suction side slashface, and wherein the platform overhangs a shank cavity. The platform cooling arrangement may include: a pocket formed in an underside region of the platform, the pocket comprising a mouth that fluidly communicates with the shank cavity; a manifold extending from the suction side slashface to a pressure side slashface, the manifold including a connection to the pocket; and cooling apertures that extend from connections made with the pocket and manifold to ports. | 07-04-2013 |
| 20130171004 | TURBINE ROTOR BLADE PLATFORM COOLING - A platform cooling arrangement in a turbine rotor blade having a platform positioned between an airfoil and a root. The rotor blade, along a side that coincides with a pressure side of the airfoil, includes a pressure side of the platform includes a topside extending from an airfoil base to a pressure side slashface. The platform cooling arrangement includes: a main plenum residing just inboard of the topside in the pressure side of the platform, the main plenum extending through the platform from an upstream end having an aft position to a downstream end having a forward position; and cooling apertures. Near the upstream end, the main plenum includes an aft switchback, and, between the aft switchback and the downstream end, a forward arc. Each of the cooling apertures extends from the main plenum to a port formed on the pressure side slashface. | 07-04-2013 |
| 20130171005 | TURBINE ROTOR BLADE PLATFORM COOLING - 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 formed therein, wherein, in operation, the interior cooling passage comprises a high-pressure coolant region and a low-pressure coolant region, and wherein a suction side of the platform comprises a topside extending circumferentially from the airfoil to a suction side slashface, and wherein the suction side of the platform comprises an aft edge. The platform cooling arrangement may include: a manifold positioned within an aft side of the suction side of the platform; a high-pressure connector that connects the manifold to the high-pressure coolant region of the interior cooling passage; a low-pressure connector that connects the manifold to the low-pressure coolant region of the interior cooling passage; and heat transfer structure positioned within the manifold. | 07-04-2013 |
| 20130177446 | SYSTEM AND METHOD FOR COOLING TURBINE BLADES - A system includes a turbine blade, which includes at least one cooling slot configured to convey a coolant in a flow direction from an interior to an exterior of the turbine blade. The cooling slot includes an entrance coupled to the interior and a converging section downstream from the entrance. The converging section includes a first cross-sectional area that decreases in the flow direction. The cooling slot also includes an exit disposed along a trailing edge of the turbine blade. | 07-11-2013 |
| 20130177447 | Turbine Nozzle Assembly Methods - The present application provides a method of installing an impingement cooling assembly in an inner platform of an airfoil of a turbine nozzle. The method may include the steps of positioning an insert within a cavity of the airfoil, positioning a core exit cover about an opening of the cavity, positioning an impingement plenum within a platform cavity, inserting an unfixed spoolie through an assembly port of the impingement plenum and into an airflow cavity of the insert, and closing the assembly port. | 07-11-2013 |
| 20130177448 | CORE FOR A CASTING PROCESS - A core for a casting process includes a core body and a first cooling hole forming portion that extends from the core body. The core body includes a varying thickness along a length of the core body. The core body can include an undulating shaped section and can be a ceramic core body. | 07-11-2013 |
| 20130183165 | AIRFOIL - An airfoil includes a platform and an exterior surface connected to the platform. A plurality of trench segments are on the exterior surface, and a single cooling passage in each trench segment supplies a cooling media to the exterior surface. | 07-18-2013 |
| 20130183166 | AIRFOIL - An airfoil includes a platform and an exterior surface connected to the platform. A plurality of trench segments are on the exterior surface, and each trench segment extends less than 50% of a length of the exterior surface. A cooling passage in each trench segment supplies a cooling media to the exterior surface. | 07-18-2013 |
| 20130195675 | CERAMIC CORE TAPERED TRIP STRIPS - A core die for creating an airfoil includes a first section, a second section mating with the first section, and an insert for creating a slot. The first section and the second section define a body having an outer dimension. The insert is disposed at an angle to the outer dimension. A trip strip includes a first portion disposed in the second section. The first portion is in register with the insert and a thickness is maintained between the first portion and the insert along a length of the insert. The first portion tapers towards the outer dimension and the thickness is filled by the ceramic material between the slot and the first portion. | 08-01-2013 |
| 20130209269 | GAS TURBINE ENGINE COMPONENT WITH MULTI-LOBED COOLING HOLE - A component for a gas turbine engine includes a wall and a cooling hole extending through the wall. The wall has a first surface and a second surface. The cooling hole includes a metering section extending downstream from an inlet in the first surface of the wall and a diffusion section extending from the metering section to an outlet in the second surface of the wall. The diffusion section includes a first plurality of lobes diverging longitudinally and laterally from the metering section on a first side of a centerline axis of the cooling hole and a second plurality of lobes diverging longitudinally and laterally from the metering section on a second side of the centerline axis. | 08-15-2013 |
| 20130209270 | METHOD FOR RECONDITIONING A BLADE OF A GAS TURBINE AND ALSO A RECONDITIONED BLADE - A method is provided for reconditioning a blade of a gas turbine. The blade includes a blade airfoil, with a pressure side and a suction side, which extends in the blade longitudinal direction between a platform and a blade tip, and has a leading edge and a trailing edge, and is outwardly delimited by a pressure-side wall and a suction-side wall which converge at the trailing edge of the blade airfoil, forming discharge openings for cooling air which are arranged in a distributed manner along the trailing edge between the walls. | 08-15-2013 |
| 20130209271 | GAS TURBINE BLADE, MANUFACTURING METHOD THEREFOR, AND GAS TURBINE USING TURBINE BLADE - Gas turbine blades which simplify the formation of cooling channels provided inside the turbine blades while simultaneously avoiding loss of turbine blade strength and rigidity due to forming of the cooling channels. Cooling channels provided in the interior of a gas turbine blade include a plurality of straight channel-like base-side elongated holes extending in a longitudinal direction at a base side of the turbine blade, a plurality of straight channel-like tip-side elongated holes extending in a longitudinal direction at a tip side of the turbine blade, and a plurality of communicating hollow portions interposed at connection portions between the two types of elongated holes to allow the two types of elongated holes to communicate with each other, and have larger cross-sectional areas than the channel cross-sectional areas of both elongated holes. The communicating hollow portions are formed to match the position of a platform portion of the turbine blade. | 08-15-2013 |
| 20130216395 | GAS TURBINE BLADE - A gas turbine blade including a root, an airfoil with a leading edge and a trailing edge, a radial outer tip, and a pressure side and a suction side between the leading edge and the trailing edge, and a cooling air channel system extending from an air inlet opening in the root throughout the airfoil to a plurality of air outlets at the pressure side and the leading edge of the top of the tip of the airfoil. | 08-22-2013 |
| 20130230407 | Turbine Bucket with a Core Cavity Having a Contoured Turn - The present application thus provides a turbine bucket. The turbine bucket may include a platform, an airfoil extending from the platform at an intersection thereof, and a core cavity extending within the platform and the airfoil. The core cavity may include a contoured turn about the intersection so as to reduce thermal stress therein. | 09-05-2013 |
| 20130230408 | Turbine Bucket with Contoured Internal Rib - The present application thus provides a turbine bucket. The turbine bucket may include a platform and an airfoil extending from the platform. The airfoil may include an internal rib with a number of through holes positioned along a number of hole spaces and a number of in-between spaces. The in-between spaces may include a first depth, the hole spaces may include a second depth, and wherein the first depth is less than the second depth. | 09-05-2013 |
| 20130236330 | TURBINE AIRFOIL WITH AN INTERNAL COOLING SYSTEM HAVING VORTEX FORMING TURBULATORS - A turbine airfoil usable in a turbine engine and having at least one cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels having a plurality of turbulators protruding from an inner surface and positioned generally nonorthogonal and nonparallel to a longitudinal axis of the airfoil cooling channel. The configuration of turbulators may create a higher internal convective cooling potential for the blade cooling passage, thereby generating a high rate of internal convective heat transfer and attendant improvement in overall cooling performance. This translates into a reduction in cooling fluid demand and better turbine performance. | 09-12-2013 |
| 20130251539 | TRAILING EDGE OR TIP FLAG ANTIFLOW SEPARATION - An airfoil includes a leading edge, a trailing edge region, a suction surface, a pressure surface, a cooling passageway, and a column of flow separators. The suction surface and the pressure surface both extend axially between the leading edge and the trailing edge region, as well as radially from a root section of the airfoil to a tip section of the airfoil to define a central cavity of the airfoil. The cooling passageway is located within the central cavity at the trailing edge region. The column of flow separators is located in the cooling passageway adjacent the trailing edge and includes a first flow separator having a first longitudinal axis and a second flow separator having a second longitudinal axis. The first longitudinal axis is offset at an angle with respect to the first longitudinal axis. | 09-26-2013 |
| 20130259703 | TURBINE NOZZLE - A nozzle arrangement for a gas turbine engine comprising a first housing member and a second housing member. The nozzle arrangement may further include a first nozzle and a second nozzle. Each of the first nozzle and second nozzle may extend between the first housing member and the second housing member so as to form a doublet. A plurality of cooling apertures may be arranged on at least one of the first nozzle, the second nozzle, the first housing member, or the second housing member so as to provide a different degree of first order cooling across the doublet. | 10-03-2013 |
| 20130259704 | TURBINE COOLING APPARATUS - A turbine blade for a gas turbine engine is disclosed. The turbine blade can include at least one internal cooling path, and an internal vane disposed in the at least one internal cooling path. The internal vane can include a central portion, a first leg extending in a first direction from the central portion, and a second leg extending in a second direction from the central portion. The central portion can have a thickness greater than a thickness of the first leg or a thickness of the second leg. | 10-03-2013 |
| 20130259705 | TURBINE AIRFOIL TRAILING EDGE COOLING SLOTS - A turbine airfoil includes pressure and suction sidewalls extending along a span from a base to a tip. Spanwise spaced apart trailing edge cooling holes in the pressure sidewall end at corresponding spanwise spaced apart trailing edge cooling slots extending chordally substantially to the trailing edge. Each cooling hole includes a plug extending downstream through at least a portion of a spanwise diverging section leading into the slot. The plug may be spanwise centered in the hole and may include a plug dome rising up from a plug base extending along a suction sidewall surface of the suction sidewall. The cooling hole may further include an inlet leading to a metering section with a constant area and constant width flow cross section upstream of spanwise diverging section. Lands may be disposed between the trailing edge cooling slots forming slot floors between lands. | 10-03-2013 |
| 20130266454 | TURBINE AIRFOIL TIP SHELF AND SQUEALER POCKET COOLING - An airfoil comprises leading and trailing edges with pressure and suction surfaces defining a chord length therebetween. The pressure and suction surfaces extend from a root section of the airfoil to a tip section. A tip shelf is formed along the tip section between the pressure surface and a tip shelf wall spaced between the pressure surface and the suction surface. A squealer pocket is formed along the tip section between the tip shelf wall and a squealer tip wall extending from the suction surface. The tip shelf extends from within 10% of the cord length measured from the leading edge to within 10% of the chord length measured from the trailing edge. The squealer pocket extends from within 10% of the chord length measured from the leading edge to terminate between 10% and 25% of the chord length measured from the trailing edge. | 10-10-2013 |
| 20130272896 | IMPINGEMENT COOLING OF GAS TURBINE BLADES OR VANES - A turbine component includes a hollow aerofoil and an impingement tube located within the hollow aerofoil. The impingement tube is formed from at least two separate sections each extending span wise through the hollow aerofoil. Adjacent sections of the impingement tube are connected together by a locking device. The locking device is insertable into the hollow aerofoil and locks the impingement tube into place in the hollow aerofoil. The locking device is a roll pin being located in an axially direction between said sections and has a main extension which extends in a radial direction of the hollow aerofoil. | 10-17-2013 |
| 20130280092 | AIRFOIL COOLING ENHANCEMENT AND METHOD OF MAKING THE SAME - 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. | 10-24-2013 |
| 20130280093 | GAS TURBINE ENGINE CORE PROVIDING EXTERIOR AIRFOIL PORTION - A core has a body that includes a cooling passage portion with a film cooling passage portion extending there from to a film cooling hole portion. An exterior airfoil portion is connected to the film cooling hole portion and is spaced apart from the cooling passage portion to provide a space surrounding the film cooling hole portion that corresponds to an exterior airfoil wall. | 10-24-2013 |
| 20130280094 | GAS TURBINE BLADE AND GAS TURBINE HAVING THE SAME - Provided is a gas turbine blade capable of improving the heat-conducting capacity of a serpentine channel. In a gas turbine blade ( | 10-24-2013 |
| 20130302176 | TURBINE AIRFOIL TRAILING EDGE COOLING SLOT - A turbine airfoil includes pressure and suction sidewalls extending along a span from a base to a tip. Spanwise spaced apart trailing edge cooling holes in pressure sidewall end at a trailing edge cooling slot extending chordally substantially to an airfoil trailing edge. Each cooling hole includes a curved inlet, a metering section with a constant area and constant width flow cross section, and a spanwise diverging section leading into slot. Axial partitions extend chordally between and radially separate cooling holes along span. Aft ends of partitions include swept boat tails. Upper and lower deck sidewalls spanwise bound a deck of slot and extend outward to an external surface of pressure sidewall. Fillets between sidewalls and deck have fillet radii substantially the same size as bottom corner radii of flow cross section of diverging sections adjacent bottom corner radii. | 11-14-2013 |
| 20130302177 | TURBINE AIRFOIL TRAILING EDGE BIFURCATED COOLING HOLES - A gas turbine engine turbine airfoil having pressure and suction sidewalls extending outwardly along a span and chordwise between opposite leading and trailing edges. A spanwise row of spaced apart bifurcated trailing edge cooling holes encased in the pressure sidewall end at corresponding trailing edge cooling slots extending chordally substantially to the trailing edge. Axially extending inter-hole partitions separate the cooling holes. An inlet between adjacent pairs of the inter-hole partitions includes a divergent inlet section. An axial intra-hole partition bifurcates the cooling hole into diverging upper and lower diverging sections downstream and aft of the divergent inlet section. A forward end of the intra-hole partition divides an aft end of the divergent inlet section into upper and lower inlet flowpaths leading to the upper and lower diverging sections leading into the trailing edge cooling slots. | 11-14-2013 |
| 20130302178 | ASYMMETRICALLY SHAPED TRAILING EDGE COOLING HOLES - A turbine airfoil includes pressure and suction sidewalls extending along a span from a base to a tip. Spanwise spaced apart trailing edge cooling holes in pressure sidewall end at corresponding spanwise spaced apart trailing edge cooling slots extending chordally substantially to trailing edge. Each cooling hole includes an asymmetric flow cross section through one or more asymmetric intermediate sections leading into slot. Flow cross section is asymmetric with respect to a mid-plane extending axially and spanwise through intermediate sections. Different trailing edge cooling holes may include different asymmetric flow cross sections. Lands may extend between the cooling slots. A raised floor may extend away from at least one of pressure or suction sidewalls at least partially through one or more asymmetric intermediate sections and optionally at least partially through cooling slot. Raised floor may include up and down ramps and a flat transition section between ramps. | 11-14-2013 |
| 20130302179 | TURBINE AIRFOIL TRAILING EDGE COOLING HOLE PLUG AND SLOT - A turbine airfoil includes pressure and suction sidewalls extending along a span from a base to a tip. Spanwise spaced apart trailing edge cooling holes disposed in pressure sidewall. All or a plurality of cooling holes end at a trailing edge cooling slot extending chordally substantially to an airfoil trailing edge. Each cooling hole includes a curved inlet, a metering section with a constant area and constant width flow cross section, and a spanwise diverging section leading into slot. Axial partitions extend chordally between and radially separate cooling holes along span. Aft ends of partitions include swept boat tails. Upper and lower deck sidewalls spanwise bound a deck of slot and extend outward to an external surface of pressure sidewall. Fillets between sidewalls and deck have fillet radii substantially the same size as bottom corner radii of flow cross section of diverging sections adjacent bottom corner radii. | 11-14-2013 |
| 20130315748 | COOLING STRUCTURES IN THE TIPS OF TURBINE ROTOR BLADES - A turbine rotor blade used in a gas turbine engine, which includes an airfoil having a tip at an outer radial edge, is described. The airfoil includes a pressure sidewall and a suction sidewall that join together at a leading edge and a trailing edge of the airfoil, the pressure sidewall and the suction sidewall extending from a root to the tip. The tip includes a tip plate and, disposed along a periphery of the tip plate, a rail. The rail includes a microchannel connected to a coolant source. | 11-28-2013 |
| 20130315749 | COOLING STRUCTURES IN THE TIPS OF TURBINE ROTOR BLADES - A turbine rotor blade for a gas turbine engine is described. The turbine rotor blade includes an airfoil that includes a tip at an outer radial end. The tip includes a rail that defines a tip cavity; and the rail includes a circumscribing rail microchannel. The circumscribing rail microchannel is a microchannel that extends around at least a majority of the length of the inner rail surface. | 11-28-2013 |
| 20130323078 | TURBINE BLADE ROOT WITH MICROCIRCUIT COOLING PASSAGES - A method of fabricating an airfoil includes the steps of fabricating a first core including a first plurality of ribs defining a first plurality of passages of a completed airfoil, and fabricating as second core including a second plurality of ribs defining a second plurality of passages of the completed airfoil. The second plurality of ribs includes a plurality of standoffs. The plurality of standoffs set a spacing between the first plurality of ribs and the second plurality of ribs to define a spacing between the first plurality of channels and the second plurality of channels of the completed airfoil. The airfoil is then molded about the core assembly. Once completed, the core assembly is removed to provide a completed airfoil incorporating multiple microcircuits with a desired stability and structural integrity. | 12-05-2013 |
| 20130323079 | TURBOMACHINERY COMPONENT COOLING SCHEME - A turbomachinery component includes a surface exposed to hot working fluid flow. The surface has an undulating contour formed from a series of alternating protuberances and troughs. A set of three cooling outlets is associated with each trough. | 12-05-2013 |
| 20130323080 | VORTEX GENERATORS FOR IMPROVED FILM EFFECTIVENESS - An airfoil includes a suction surface, a pressure surface, a first showerhead cooling hole and a second showerhead cooling hole. The suction surface and the pressure surface both extend axially between a leading edge and a trailing edge, as well as radially from a root section to a tip section. The first showerhead cooling hole and the second showerhead cooling hole both extend into pressure surface near the leading edge. The first showerhead cooling hole and the second showerhead cooling hole are angled in opposing directions. | 12-05-2013 |
| 20140003960 | AIRFOIL | 01-02-2014 |
| 20140003961 | GAS TURBINE ENGINE COMPONENT HAVING PLATFORM COOLING CHANNEL | 01-02-2014 |
| 20140003962 | TURBINE BLADE | 01-02-2014 |
| 20140010666 | AIRFOIL COOLING CIRCUITS - An airfoil includes leading and trailing edges; first and second sides extending from the leading edge to the trailing edge, each side having an exterior surface; a core passage located between the first and second sides and the leading and trailing edges; and a wall structure located between the core passage and the exterior surface of the first side. The wall structure includes a plurality of cooling fluid inlets communicating with the core passage for receiving cooling fluid from the core passage, a plurality of cooling fluid outlets on the exterior surface of the first side for expelling cooling fluid and forming a cooling film along the exterior surface of the first side, and a plurality of cooling passages communicating with the plurality of cooling fluid inlets and the plurality of cooling fluid outlets. At least a portion of one cooling passage extends between adjacent cooling fluid outlets. | 01-09-2014 |
| 20140023517 | NOZZLE FOR TURBINE SYSTEM - A nozzle for a turbine system is disclosed. The nozzle includes an airfoil, an inner sidewall, and an outer sidewall. The airfoil includes exterior surface defining a pressure side and a suction side extending between a leading edge and a trailing edge. The airfoil further defines a tip and a root. The inner sidewall is connected to the airfoil at the tip. The outer sidewall is connected to the airfoil at the root. The inner sidewall and outer sidewall each includes a peripheral edge defining a pressure side slash face, a suction side slash face, a leading edge face, and a trailing edge face. At least one of the inner sidewall pressure side slash face, the inner sidewall suction side slash face, the outer sidewall pressure side slash face, or the outer sidewall suction side slash face has a generally curvilinear profile. | 01-23-2014 |
| 20140037458 | COOLING STRUCTURES FOR TURBINE ROTOR BLADE TIPS - A rotor blade for a turbine of a combustion turbine engine having an airfoil that includes a pressure and a suction sidewall defining an outer periphery and a tip portion defining an outer radial end. The tip portion includes a rail that defines a tip cavity. The airfoil includes an interior cooling passage configured to circulate coolant. The rotor blade further includes: a slotted portion of the rail; and at least one film cooling outlet disposed within at least one of the pressure sidewall and the suction sidewall of the airfoil. The film cooling outlet includes a position that is adjacent to the tip portion and in proximity to the slotted portion of the rail. | 02-06-2014 |
| 20140037459 | AIRFOIL DESIGN HAVING LOCALIZED SUCTION SIDE CURVATURES - An airfoil for a gas turbine engine comprises a radially extending body having a transverse cross-section. The transverse cross-section comprises a leading edge, a trailing edge, a pressure side and a suction side. The pressure side extends between the leading edge and the trailing edge with a predominantly concave curvature. The suction side extends between the leading edge and the trailing edge with a predominantly convex curvature. The suction side includes an approximately flat portion flanked by forward and aft convex portions. In another embodiment, the suction side includes a series of local curvature changes that produce inflection points in the convex curvature of the suction side spaced from the trailing edge. | 02-06-2014 |
| 20140037460 | COOLED BLADE FOR A GAS TURBINE - The invention relates to a cooled blade for a gas turbine that includes a radially extending aerofoil with a leading edge, a trailing edge, a suction side and a pressure side, and wherein a lip overhang is provided on the suction side of the trailing edge The blade also includes a plurality of radial internal flow channels connected via flow bends to form a multi-pass serpentine for a coolant flow, whereby a trailing edge ejection region is provided for cooling said trailing edge, said trailing edge ejection region comprising a trailing edge passage of said multi-pass serpentine running essentially parallel to said trailing edge and being connected over its entire length with a pressure side bleed. An optimized cooling is achieved by mainly determining the cooling flow from the trailing edge passage to the pressure side bleed by means of a staggered field of pins, which is provided between said pressure side bleed and said trailing edge passage, with the lateral dimension of said pins increasing in coolant flow direction. | 02-06-2014 |
| 20140037461 | COOLING SYSTEM IN A TURBINE AIRFOIL ASSEMBLY INCLUDING ZIGZAG COOLING PASSAGES INTERCONNECTED WITH RADIAL PASSAGEWAYS - An airfoil in a gas turbine engine includes an outer wall, a cooling fluid cavity, a plurality of cooling fluid passages, and a plurality of radial passageways. The outer wall has leading and trailing edges, pressure and suction sides, and radially inner and outer ends. The cooling fluid cavity is defined in the outer wall and receives cooling fluid for cooling the outer wall. The cooling fluid passages are in fluid communication with the cooling fluid cavity and include alternating angled sections, each section having both a radial component and a chordal component. The cooling fluid passages extend from the cooling fluid cavity toward the trailing edge of the outer wall and receive cooling fluid from the cooling fluid cavity for cooling the outer wall near the trailing edge. The radial passageways interconnect radially adjacent cooling fluid passages. | 02-06-2014 |
| 20140044556 | LAST STAGE BLADE INCLUDING A PLURALITY OF LEADING EDGE INDENTATIONS - Aspects of the invention provide for a last stage blade of a steam turbine. In one embodiment, the last stage blade of a steam turbine includes: a blade leading edge for receiving a through flow, the blade leading edge including a plurality of indentations at a top portion of the steam turbine blade; and a blade trailing edge for exhausting the through flow. | 02-13-2014 |
| 20140044557 | TURBINE BLADE AND METHOD FOR COOLING THE TURBINE BLADE - A turbine blade includes an outer surface, and a rim surrounds at least a portion of the outer surface. A trench in the rim extends around at least a portion of the outer surface. | 02-13-2014 |
| 20140056717 | GAS TURBINE ENGINE AIRFOIL INTERNAL COOLING FEATURES - An airfoil for a gas turbine engine includes spaced apart pressure and suction walls joined at leading and trailing edges to provide an airfoil. Intermediate walls interconnect the pressure and suction walls to provide cooling passageways. The cooing passageways have interior pressure and suction surfaces that are respectively provided on the pressure and suction walls. Trip strips include a chevron-shaped trip strip that is provided on at least one of the interior pressure and suction surfaces. | 02-27-2014 |
| 20140056718 | INLET SHROUD ASSEMBLY - An inlet shroud assembly includes a hub and a shroud for a ram air fan. The hub includes a center bore, a first frustoconical section, a plurality of slotted cooling holes located on the first frustoconical section, a second frustoconical section, a plurality of circular cooling holes located on the second frustoconical section, a rim, a central cavity, and an annular cavity. The shroud includes a central disk with a center bore, a circular flange located on a radially outer edge of the central disk, and an outer disk with a web portion and a tip portion, the tip portion having a curved lip. | 02-27-2014 |
| 20140056719 | Gas Turbine, Gas Turbine Blade, and Manufacturing Method of Gas Turbine Blade - A gas turbine blade includes a hollow-blade-form portion formed by a leading edge on an upstream side of an working fluid of a gas turbine in a flow direction, a trailing edge on a downstream side of the working fluid in the flow direction, and a suction surface and a pressure surface reaching the trailing edge from the leading edge, and a shank portion for supporting the blade form portion. The blade also includes a partition for connecting the suction surface and the pressure surface in a hollow region of the blade-form portion, coolant paths formed by the partition, the suction surface, and the pressure surface, an impingement cooling hole formed in the partition for dividing the first path that is a flow path nearest to the leading edge side among the coolant paths and the second path adjacent to the first path, and first and second converter portions. | 02-27-2014 |
| 20140064983 | AIRFOIL AND METHOD FOR MANUFACTURING AN AIRFOIL - An airfoil includes a pressure side, a suction side opposed to the pressure side, a cavity inside the airfoil between the pressure and suction sides, and a trailing edge downstream from the cavity between the pressure and suction sides. A first set of cooling passages through the trailing edge provide fluid communication from the cavity through the trailing edge. A first divider across each cooling passage in the first set of cooling passages extends from the pressure side to the suction side at the trailing edge. | 03-06-2014 |
| 20140064984 | COOLING ARRANGEMENT FOR PLATFORM REGION OF TURBINE ROTOR BLADE - A platform cooling arrangement in a turbine rotor blade having a platform at an interface between an airfoil and a root. The platform may include a pressure side slashface and a suction side slashface. The platform cooling arrangement may include: a cooling channel formed within the interior of the platform, the cooling channel extending from a first end toward one of the pressure side slashface and the suction side slashface. At a second end, the cooling channel may include a pocket. The pocket may include an abrupt increase in cross-sectional flow area just before the cooling channel reaches the slashface. | 03-06-2014 |
| 20140072448 | SYSTEM AND METHOD FOR AIRFOIL COVER PLATE - A system includes at least one gas turbine blade that includes at least one recess disposed in a surface of the at least one gas turbine blade. The system also includes a porous insert that includes multiple pores that enable a cooling fluid to flow through the porous insert. The system further includes a cover plate, wherein the porous insert is disposed within the at least one recess, and the cover plate is disposed over the porous insert to enclose the porous insert within the at least one recess. | 03-13-2014 |
| 20140093387 | METHOD OF MANUFACTURING A COOLED TURBINE BLADE WITH DENSE COOLING FIN ARRAY - 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. | 04-03-2014 |
| 20140093388 | COOLED TURBINE BLADE WITH LEADING EDGE FLOW DEFLECTION AND DIVISION - A cooled turbine blade having a base and an airfoil, the base including cooling air inlet and an internal cooling air passageway, and the airfoil including an internal heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a “skin” that encompasses a tip wall, an inner spar, a leading edge rib, and a leading edge air deflector. The leading edge rib is configured to form a leading edge chamber in conjunction with the leading edge of the skin. The leading edge air deflector is at least partially intersected by the inner spar. | 04-03-2014 |
| 20140093389 | COOLED TURBINE AIRFOIL STRUCTURES - In accordance with an exemplary embodiment, disclosed is an air-cooled turbine blade having an airfoil shape, including a convex suction side wall, a concave pressure side wall, the walls including an interior surface that defines an interior with the blade, a suction side flow circuit formed within the blade interior, a pressure side flow circuit formed within the blade interior; and a trailing edge pin bank positioned aft of the suction side and pressure side flow circuits. The turbine blade includes a wishbone-shaped architecture at a transition point between the suction side flow circuit and the pressure side flow circuit and the trailing edge pin bank. | 04-03-2014 |
| 20140093390 | COOLED TURBINE BLADE WITH LEADING EDGE FLOW REDIRECTION AND DIFFUSION - A cooled turbine blade having a base and an airfoil, the base including cooling air inlet and an internal cooling air passageway, and the airfoil including an internal heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a “skin” that encompasses a tip wall, an inner spar, a leading edge rib, and a leading edge air deflector. The leading edge rib is configured to form a leading edge chamber in conjunction with the leading edge of the skin. The leading edge air deflector is shaped and positioned such that cooling air leaving the leading edge chamber is both turned and diffused. | 04-03-2014 |
| 20140093391 | COOLED TURBINE BLADE WITH TRAILING EDGE FLOW METERING - A cooled turbine blade having a base and an airfoil, the base including cooling air inlet and an internal cooling air passageway, and the airfoil including an internal heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a “skin” that encompasses a tip wall, an inner spar, a plurality of trailing edge cooling fins, and a perforated first and second trailing edge rib configured to meter cooling air passing there thorough. | 04-03-2014 |
| 20140093392 | GAS TURBINE ENGINE COMPONENT - Described is a gas turbine engine component ( | 04-03-2014 |
| 20140112799 | COOLING ARRANGEMENT FOR A GAS TURBINE COMPONENT - A cooling arrangement ( | 04-24-2014 |
| 20140119944 | Film Cooling Channel Array with Multiple Metering Portions - A cooling channel array for a gas turbine engine is provided. The cooling channel array is carried by a component wall having an inner surface and an outer surface and comprises at least two metering portions that communicate with a diffusion cavity. | 05-01-2014 |
| 20140140860 | AEROFOIL COOLING - An aerofoil component of a gas turbine engine is provided. The component has a longitudinally extending aerofoil portion which spans, in use, a working gas annulus of the engine. The aerofoil portion contains an internal chamber for a flow of coolant. The chamber includes a helical passage which spirals in a plurality of turns around an axis that extends in the length direction of the aerofoil portion. | 05-22-2014 |
| 20140154096 | TURBINE BLADE AIRFOILS INCLUDING SHOWERHEAD FILM COOLING SYSTEMS, AND METHODS FOR FORMING AN IMPROVED SHOWERHEAD FILM COOLED AIRFOIL OF A TURBINE BLADE - 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. | 06-05-2014 |
| 20140186190 | TIP LEAKAGE FLOW DIRECTIONALITY CONTROL - An airfoil according to an exemplary aspect of the present disclosure includes, among other things, a suction sidewall and a pressure sidewall, each sidewall extending spanwise from an airfoil base and extending chordwise between a leading edge and a trailing edge. A tip wall extends chordwise from the leading edge to the trailing edge and joining respective outer spanwise ends of the suction and pressure sidewalls. A tip leakage control channel is recessed into the tip wall, the tip leakage control channel including a control channel floor that extends between a first control channel vane sidewall and a second control channel vane sidewall established by a corresponding tip leakage control vane. An internal cooling cavity is disposed between the suction sidewall and the pressure sidewall and a channel cooling aperture feeds airflow from the internal cooling cavity to the tip leakage control channel. | 07-03-2014 |
| 20140193274 | FUEL-COOLED BLADED ROTOR OF A GAS TURBINE ENGINE - Liquid fuel from a rotary fluid trap is atomized by a sharp edge on an inside surface of an aft cavity of a bladed rotor of a gas turbine engine and directed into each of a plurality of associated hollow blades through corresponding blade inlet ducts that are in fluid communication with corresponding aft hollow interior portions of each blade. A radially-extending central rib within each blade partitions the hollow interior thereof into aft and forward hollow interior portions that are in fluid communication through an associated opening in the central rib and through a radially-extending gap between the central rib and the interior surface of the blade. A blade outlet duct provides for fluid communication between the forward hollow interior portion and a forward cavity of the bladed rotor, and a rotor outlet duct provides for discharging the fuel from a radially-inboard portion of the forward cavity. | 07-10-2014 |
| 20140199177 | AIRFOIL AND METHOD OF MAKING - An airfoil includes leading and trailing edges, a first exterior wall extending from the leading edge to the trailing edge and having inner and outer surfaces, a second exterior wall extending from the leading edge to the trailing edge generally opposite the first exterior wall and having inner and outer surfaces, and cavities within the airfoil. A first cavity extends along the inner surface of the first exterior wall and a first inner wall and has an upstream end and a downstream end, and a feed cavity is located between the first inner wall and the second exterior wall. | 07-17-2014 |
| 20140219815 | MULTI-LOBED COOLING HOLE - A gas turbine engine component subjected to a flow of high temperature gas includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe, a second lobe and a transition region. The first lobe diverges longitudinally and laterally from the metering section and has a first shape. The second lobe is generally opposite the first lobe and diverges longitudinally and laterally from the metering section and has a second shape different from the first shape. The transition region is positioned between the first and second lobes and includes a downstream end adjacent the outlet. | 08-07-2014 |
| 20140219816 | HIGH PRESSURE TURBINE BLADE COOLING HOLE DISTRIBUTION - A turbine blade for a gas turbine engine with an airfoil portion defined by a perimeter wall surrounding at least one enclosure, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the at least one enclosure and a gaspath of the gas turbine engine. The plurality of cooling holes includes at least one row of holes selected from the group consisting of a first row, a second row, a third row and a fourth row, with the first, second, third and fourth rows of holes respectively including the holes numbered PA-1 to PA-10, PB-1 to PB-3, HA-1 to HA-9, and SA-1 to SA-8 located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3. | 08-07-2014 |
| 20140219817 | HIGH PRESSURE TURBINE BLADE COOLING HOLE DISTRIBUTION - A turbine blade for a gas turbine engine with an airfoil portion defined by a perimeter wall surrounding at least one enclosure, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the at least one enclosure and a gaspath of the gas turbine engine. The plurality of cooling holes includes at least one row of holes selected from the group consisting of a first row, a second row, a third row and a fourth row, with the first, second, third and fourth rows of holes respectively including the holes numbered PA-1 to PA-10, PB-1 to PB-3, HA-1 to HA-9, and SA-1 to SA-8 located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3. | 08-07-2014 |
| 20140219818 | Turbine Component Cooling Channel Mesh with Intersection Chambers | 08-07-2014 |
| 20140234121 | FILM COOLING STRUCTURE AND TURBINE BLADE - The film cooling structure includes: a wall surface along which a heating medium flows; and at least one pair of film cooling holes that open at the wall surface and that blow a cooling medium. The pair of film cooling holes are arranged to be adjacent to each other in a main flow direction of the heating medium. In addition, perforation directions of the pair of film cooling holes are set such that directions of swirls of the cooling medium formed by blowing are opposite to each other, a swirl of the cooling medium on a downstream side in the main flow direction is mixed and merged with another swirl of the cooling medium on an upstream side in the main flow direction, and the merged cooling medium flows along the wall surface in a direction intersecting with the main flow direction. | 08-21-2014 |
| 20140255207 | TURBINE ROTOR BLADES HAVING MID-SPAN SHROUDS - A rotor blade for use in a turbine of a combustion turbine engine is described. The rotor blade may include an airfoil that extends from a connection with a root. The rotor blade may further include a mid-span shroud configured to engage a corresponding mid-span shroud on at least one neighboring rotor blades during operation. Outboard of the mid-span shroud, the airfoil may include an outboard region that is substantially hollow, and inboard of the mid-span shroud, the airfoil may include an inboard region that is substantially solid. | 09-11-2014 |
| 20140255208 | Gas Turbine Blade - The gas turbine blade includes a cooling channel formed therein, and a partition disposed on its tip side for isolating the cooling channel from the outside. The partition is integrally formed with a blade portion in a position on its inner side in the radial direction with respect to the tip of the gas turbine blade. Reinforcements are provided on the outer side of the partition in the radial direction and on the inner side of the tip end wall extended from the blade portion to connect the partition with the tip end wall. Outer surface cooling holes are formed to extend from the cooling channel into communication with an outer surface of the tip end wall, and inner surface cooling holes are formed to extend from the cooling channel into an inner surface of the tip end wall through the partition. | 09-11-2014 |
| 20140271226 | Turbine Blade Tip With Tip Shelf Diffuser Holes - A turbine blade having a tip including a tip shelf through which pass one or more diffuser cooling holes, the diffuser cooling holes directing the flow of cooling gas along the tip shelf, spreading the flow of cooling gas more evenly along the tip shelf and enhancing the formation of a curtain of cooling air along the tip shelf, leading to lower blade tip temperatures, reduced diversion of cooling air, and greater turbine blade life. | 09-18-2014 |
| 20140271227 | MANUFACTURE OF HOLLOW AEROFOIL - A method of manufacturing a hollow aerofoil component for a gas turbine engine includes using a capping panel to cover a pocket in a pocketed aerofoil body. During manufacture, the outer surface of the capping panel is located relative to the pocketed aerofoil body. This ensures that the outer surface of the capping panel is located as accurately as possible. This means that the capping panel can be made to be as thin as possible, which in turn reduces weight and material wastage. Once the capping panel has been located in position, it may be welded to the aerofoil body in order to produce the hollow aerofoil component. | 09-18-2014 |
| 20140271228 | TURBINE BLADE - A turbine blade has hollowness, and is provided with a back-side wall of which a portion of the inner wall surface is exposed at the rear edge portion, with cooling air flown along the inner wall surface at the exposed region; and a recess provided in the inner wall surface at the exposed region. The contour of the recess ( | 09-18-2014 |
| 20140271229 | TURBINE BLADE - The present invention is a turbine blade ( | 09-18-2014 |
| 20140286790 | Dust Mitigation for Turbine Blade Tip Turns - A dust mitigation system for airfoils includes a plurality of contoured tip turns which curve about at least two axes. This inhibits recirculation areas common within airfoils and further inhibits dust build up within the cooling flow path of the airfoil. | 09-25-2014 |
| 20140286791 | Component Cooling Channel - A cooling channel ( | 09-25-2014 |
| 20140294598 | TURBINE BLADE - The cooling effectiveness of a turbine blade that a gas turbine engine or the like is provided with is further increased by providing a convex portion that is arranged in the inner portion of a cooling air hole and that is provided projecting out from the inner wall surface of the cooling air hole. | 10-02-2014 |
| 20140322028 | GAS TURBINE BLADE WITH TIP SECTIONS OFFSET TOWARDS THE PRESSURE SIDE AND WITH COOLING CHANNELS - A hollow blade including an airfoil extending along a longitudinal direction, a root, a tip, an internal cooling passage, and an open cavity defined by an end wall and a rim, together with cooling channels connecting the internal cooling passage to a pressure side. The cooling channels slope relative to the pressure side. A stack of airfoil sections of the blade at a level of the rim of the tip of the blade are offset towards the pressure side. The pressure side wall of the airfoil includes a projecting portion and cooling channels arranged in the projecting portion to open out into a terminal face of the projecting portion. | 10-30-2014 |
| 20140348665 | PIN-FIN ARRAY - The present application provides an airfoil with a cooling flow therein. The airfoil may include an internal cooling passage, a number of cooling holes in communication with the internal cooling passage, and a number of pin-fins positioned within the internal cooling passage. The pin-fins are arranged with one or more turning openings and one or more guiding openings so as to direct the cooling flow towards the cooling holes. | 11-27-2014 |
| 20150017018 | TURBINE COMPONENT AND METHODS OF ASSEMBLING THE SAME - A turbine component is provided. The turbine component includes an airfoil having a first surface and a second surface. A thermal barrier coating is coupled to the second surface, wherein the thermal barrier coating includes a first portion, a second portion and a trench defined between the first and second portions. A channel is coupled in flow communication to the first surface and the trench, wherein the channel includes a first sidewall and a second sidewall opposite of the first sidewall. The first and second sidewalls extend from the first surface and toward the trench at an angle. The turbine component includes a cover coupled to the second surface, wherein the cover includes a first end coupled to the first portion and a second end extending into the trench and spaced from the second portion. | 01-15-2015 |
| 20150030460 | METHODS FOR MODIFYING COOLING HOLES WITH RECESS-SHAPED MODIFICATIONS AND COMPONENTS INCORPORATING THE SAME - Methods for modifying a plurality of cooling holes of a component include disposing a recess-shaped modification in a recess of the component comprising a plurality of cooling hole outlets, wherein the recess-shaped modification is formed to substantially fill the recess and comprising a plurality of modified cooling holes passing there through. The method further includes aligning the plurality of modified cooling holes of the recess-shaped modification with the plurality of cooling hole outlets of the component, and, bonding the recess-shaped modification disposed in the recess to the component, wherein the plurality of modified cooling holes of the recess-shaped modification is fluidly connected with the plurality of cooling holes of the component. | 01-29-2015 |
| 20150030461 | IMPINGEMENT COOLING OF TURBINE BLADES OR VANES - A turbine assembly is provided having a hollow aerofoil having a cavity with an impingement tube insertable inside the cavity and used for impingement cooling of an inner surface of the cavity, and a platform arranged at a radial end of the hollow aerofoil, and a cooling chamber used for cooling of the platform which is arranged relative to the hollow aerofoil on an opposed side of the platform. The cooling chamber is limited at a first radial end from the platform and at an opposed radial second end from a cover plate. The impingement tube is formed from a leading piece and a trailing piece. The leading piece extends in span wise direction at least completely through the cooling chamber from the platform to the cover plate and the trailing piece terminates in span wise direction at the platform. | 01-29-2015 |
| 20150037167 | TURBINE BLADE AND TURBINE WITH IMPROVED SEALING - The disclosure pertains to a turbine with a gas turbine blade and a rotor heat shield for separating a space region through which hot working medium flows from a space region inside a rotor arrangement of the turbine. The rotor heat shield includes a platform which forms an axial heat shield section and which is arranged substantially parallel to the surface of a rotor and a radial heat shield section at the upstream end of the axial heat shield section, which is extending in a direction away from the surface of the axial heat shield section towards the hot gas. Further the turbine comprises a blade rear cavity which is delimited by the downstream end of the platform and/or the downstream end of the blade foot, the radial heat shield section. The disclosure further refers to a gas turbine blade and a rotor heat shield designed for such a turbine. | 02-05-2015 |
| 20150044059 | AIRFOIL FOR A TURBINE SYSTEM - An airfoil includes a main portion formed of a base material. Also included is a trailing edge region of the main portion. Further included is a trailing edge supplement structure comprising at least one pre-sintered preform (PSP) material operatively coupled to the base material proximate the trailing edge region. | 02-12-2015 |
| 20150064020 | TURBINE BLADE OR VANE WITH SEPARATE ENDWALL - A turbine engine airfoil structure including an airfoil adapted to be supported to extend across a gas passage for a hot working gas in a turbine engine. The airfoil structure further includes a platform structure located at one end of the airfoil and positioned at a location forming a boundary of the gas passage. The platform structure includes a platform member including a gas side surface extending generally perpendicular from the airfoil at a junction with the airfoil, and providing a structural connection to the airfoil. The platform structure further includes a separately formed platform cover attached to the platform member at the gas side surface. The platform cover extends from a location adjacent to one of the sidevvalls of the airfoil, and includes an outer surface located for contact with the hot working gas passing through the gas path. | 03-05-2015 |
| 20150078916 | TURBINE BLADES WITH TIP PORTIONS HAVING CONVERGING COOLING HOLES - A turbine rotor blade is provided with for a turbine section of an engine that includes a shroud surrounding the rotor blade. The rotor blade includes a platform and an airfoil extending from the platform into a mainstream gas path. The airfoil includes a pressure side wall, a suction side wall joined to the pressure side wall at a leading edge and a trailing edge, a tip cap extending between the suction side wall and the pressure side wall, a first squealer tip extension extending from the pressure side wall at a first angle relative to the pressure side wall, the first squealer tip extension defining a first cooling hole that converges between an inlet and an outlet; an internal cooling circuit configured to deliver cooling air to a gap between the pressure side squealer tip extension and the shroud via the first cooling hole. | 03-19-2015 |
| 20150086381 | INTERNALLY COOLED AIRFOIL - An internally cooled airfoil for a gas turbine engine has a hollow airfoil body including pressure and suction sidewalls defining a cooling passage therebetween. A combination of pedestal and trip-strips are used in the cooling passage to enhance heat transfer while minimizing the coolant pressure drop across these features. | 03-26-2015 |
| 20150093252 | INTERNALLY COOLED AIRFOIL - An internally cooled airfoil for a gas turbine engine has a hollow airfoil body defining a core cavity. An insert is mounted in the core cavity. A cooling gap is provided between the insert and the hollow airfoil body. A plurality of standoffs project across the cooling gap. Trip-strips projecting laterally between adjacent standoffs. The trip-strips and the standoffs may be integrated into a unitary heat transfer feature. | 04-02-2015 |
| 20150104326 | TURBINE ROTOR BLADES WITH IMPROVED TIP PORTION COOLING HOLES - A turbine rotor blade is provided for a turbine section of an engine. The turbine rotor blade includes a platform and an airfoil extending from the platform into a mainstream gas path of the turbine section. The airfoil includes a first side wall; a second side wall joined to the first side wall at a leading edge and a trailing edge; a tip cap extending between the first side wall and the second side wall; a first parapet wall extending from the first side wall; and a first cooling hole through the tip cap and the first parapet wall configured to deliver cooling air. The first cooling hole has a closed channel section and an open channel section. The open channel section forms a slot. | 04-16-2015 |
| 20150104327 | TURBINE ROTOR BLADES WITH TIP PORTION PARAPET WALL CAVITIES - In accordance with an exemplary embodiment, a turbine rotor blade is provided for a turbine section of an engine. The turbine rotor blade includes a platform and an airfoil extending from the platform into a mainstream gas path of the turbine section. The airfoil includes a first side wall; a second side wall joined to the first side wall at a leading edge and a trailing edge; a tip cap extending between the first side wall and the second side wall; a first parapet wall extending from the first side wall; a first parapet wall cavity formed at least partially within the first parapet wall; and a first cooling hole extending between the first parapet wall cavity and a first surface of the first parapet wall such that cooling air flows through the first parapet wall cavity, through the first cooling hole, and out of the first parapet wall. | 04-16-2015 |
| 20150110641 | TURBINE BUCKET BASE HAVING SERPENTINE COOLING PASSAGE WITH LEADING EDGE COOLING - Various embodiments of the invention include turbine buckets and systems employing such buckets. Various particular embodiments include a turbine bucket having: a base including: a casing having at least one exhaust aperture on an outer surface of the casing; and a core within the casing, the core having: a serpentine cooling passage; and at least one outlet passage fluidly connected with the serpentine cooling passage and the exhaust aperture, wherein the at least one outlet passage permits flow of a coolant from the serpentine cooling passage to the at least one exhaust aperture on the outer surface of the casing; and an airfoil connected with the base at a first end of the airfoil, the airfoil including: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side | 04-23-2015 |
| 20150118063 | TURBINE AIRFOIL TIP SHELF AND SQUEALER POCKET COOLING - An airfoil comprises pressure and suction surfaces extending from a root section to a tip section of the airfoil. The airfoil also comprises a leading edge and trailing edge defining a chord length therebetween. A tip shelf is formed along the tip section between the pressure surface and a tip shelf wall, the tip shelf wall being spaced between the pressure surface and the suction surface. A squealer pocket is formed along the tip section between the tip shelf wall and a squealer tip wall extending from the suction surface. The tip shelf extends from within 10% of the cord length measured from the leading edge to within 10% of the chord length measured from the trailing edge. The squealer pocket extends from within 10% of the chord length measured from the leading edge to terminate less than 85% of the chord length measured from the trailing edge. | 04-30-2015 |
| 20150118064 | GAS TURBINE ENGINE AIRFOIL TRAILING EDGE PASSAGE AND CORE FOR MAKING SAME - An airfoil has a body that includes leading and trailing edges that adjoin pressure and suction sides to provide an exterior airfoil surface. A cooling passage extends in a radial direction from a root to a tip. A trailing edge cooling passage interconnects the cooling passage to the trailing edge. The trailing edge cooling passage includes first and second pedestals of different sizes that are arranged in a repeating pattern with respect to pedestals of the same size and with respect to pedestals of different sizes. | 04-30-2015 |
| 20150125309 | WIND TURBINE BLADE - A wind turbine blade having an elongated blade body extending along a longitudinal axis and having an upper skin and a lower skin, the lower skin spaced from the upper skin in a thickness direction of the blade body, the upper skin and/or lower skin having a laminated layer, the laminated layer having an outer layer wherein the outer layer forms part of the upper and/or lower skin respectively, an inner layer spaced from the outer layer in the thickness direction; and an intermediate layer sandwiched between the outer layer and inner layer, the intermediate layer having a plurality of heat transfer paths within the intermediate layer for transferring heat. | 05-07-2015 |
| 20150125310 | Gas Turbine Airfoil - A gas turbine airfoil is provided that is superior in the cooling performance of an end wall and in the thermal efficiency of a gas turbine. A gas turbine airfoil includes an airfoil portion having a cooling passage therein; and an end wall portion located at an inner band end portion of the airfoil portion in the turbine-radial direction. Cooling holes are disposed in the leading edge side hook portion of the end wall portion. The plurality of cooling holes are arranged at different distance of intervals in the circumferential direction of the gas turbine. Cooling air that has flowed in the cooling passage is configured to flow from the cooling holes toward the leading edge of the end wall portion. | 05-07-2015 |
| 20150139813 | TURBINE - Provided is a turbine including a rotor; a blade provided on the rotor and comprising a cooling flow path through which a cooling fluid flows; and a shroud surrounding an exterior of the blade, wherein the blade includes: at least one rib turbulator protruding into the cooling flow path; and at least one subsidiary protrusion protruding from an outer surface of the at least one rib turbulator. | 05-21-2015 |
| 20150139814 | Gas Turbine Blade - The invention provides a gas turbine blade that is capable of reducing the temperature difference between the pressure side and the suction side even if the trailing-edge cooling channel is narrow, thereby lessening thermal stress as well. | 05-21-2015 |
| 20150292334 | TURBINE AIRFOIL TIP SHELF AND SQUEALER POCKET COOLING - An airfoil includes a pressure surface and a suction surface extending from a root section of the airfoil to a tip section of the airfoil. The airfoil also includes a leading edge and a trailing edge defining a chord length of the airfoil therebetween. The airfoil further includes a tip shelf formed along the tip section of the airfoil between the pressure surface and a tip shelf wall. The tip shelf wall is spaced between the pressure surface and the suction surface and the tip shelf extends from within 10% of the chord length measured from the leading edge to within 10% of the chord length measured from the trailing edge. | 10-15-2015 |
| 20150292335 | ROTOR BLADE - A turbine blade has a trailing end and a leading end and a tip. A gutter is formed in the tip and extends to an exit defined in a region of the trailing end of the blade. The gutter is defined at least in part by a floor, and the floor defines a decreased depth in a region proximal to the exit than in a region distal to the exit. | 10-15-2015 |
| 20150345304 | GAS TURBINE ENGINE COMPONENT HAVING VASCULAR ENGINEERED LATTICE STRUCTURE - A component according to an exemplary aspect of the present disclosure includes, among other things a wall and a vascular engineered lattice structure formed inside of the wall. The vascular engineered lattice structure defines a hollow vascular structure configured to communicate a fluid through the vascular engineered lattice structure. The vascular engineered lattice structure has at least one inlet hole and at least one outlet hole that communicates the fluid into and out of the hollow vascular structure. | 12-03-2015 |
| 20150354368 | MICRO JET GAS FILM GENERATION APPARATUS - A micro jet gas film generation apparatus aims to eject a gas to a work object which is desired for cooling or insulating from heat. The micro jet gas film generation apparatus has a spout formed at a diameter of 5-100 μm to generate a gas film on the work object. As the diameter of the spout of the micro jet gas film generation apparatus is small, the micro jet gas film generated from the spout cannot produce a large eddy due to the lack of sufficient energy, hence can maintain a thin film after a long distance ejection to improve cooling and heat insulation performance. Moreover, due to the small diameter of the spout, it also consumes less amount of gas and can reduce the amount of gas required. | 12-10-2015 |
| 20150361800 | AIRFOIL WITH VARIABLE LAND WIDTH AT TRAILING EDGE - An internally cooled airfoil, such as a turbine blade, has an airfoil section extending between a tip and a root. The interior of the airfoil includes a distribution of lands at the trailing edge in the span direction. A width of each of the lands is a widest dimension in the span direction of the land in the interior of the airfoil. A pitch is a distance in the span direction between centerlines of two adjacent lands. The pitch is constant throughout the distribution of the lands. The distribution of the lands includes at least two different widths. | 12-17-2015 |
| 20160003056 | GAS TURBINE ENGINE SHAPED FILM COOLING HOLE - A component for a gas turbine engine includes a wall that adjoins an interior cooling passage and provides an exterior surface. A film cooling hole fluidly connects the interior cooling passage and the exterior surface. The film cooling passage includes inlet and outlet passages that fluidly interconnect and adjoin one another in a misaligned non-line of sight relationship. | 01-07-2016 |
| 20160047251 | COOLING HOLE HAVING UNIQUE METER PORTION - A gas turbine engine component has a cooling hole with a metering section. The metering section includes a convex surface and a concave surface, with a first arcuate channel connecting an end of the convex surface and an end of the concave surface. The end of the convex surface and the end of the concave surface define a dimension that is smaller than a diameter of the arcuate channel. | 02-18-2016 |
| 20160108739 | FILM HOLE WITH PROTRUDING FLOW ACCUMULATOR - A wall of a gas turbine engine is provided. The wall may comprise an external surface adjacent a gas path and an internal surface adjacent an internal flow path. A film hole may have an inlet at the internal surface and an outlet at the external surface. A flow accumulator adjacent the inlet may protrude from the internal surface. A method of making an engine component is also provided and comprises the step of forming a component wall comprising an accumulator on an internal surface and a film hole defined by the component wall. The film hole may include an opening adjacent the accumulator and defined by the internal surface. | 04-21-2016 |
| 20160115796 | TURBINE BLADE COOLING STRUCTURE - In a structure for internally cooling a turbine blade, a cooling medium passage is provided in the turbine blade. The cooling medium passage has a shape in which a plurality of cylindrical spaces, each having substantially cylindrical shape, extending in parallel with each other partially overlap each other. A cooling medium supply passage that supplies a cooling medium to the cooling medium passage is connected to a portion of the cooling medium passage that includes a peripheral wall, in a direction that forms an acute angle with respect to a longitudinal direction of the cooling medium passage. | 04-28-2016 |
| 20160251966 | ENGINE COMPONENT | 09-01-2016 |
| 20190145265 | FILLET OPTIMIZATION FOR TURBINE AIRFOIL | 05-16-2019 |
| 20130280091 | GAS TURBINE ENGINE AIRFOIL IMPINGEMENT COOLING - An airfoil has a body that includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface. A leading edge wall provides the exterior airfoil surface at the leading edge. An impingement wall is integrally formed with the leading edge wall to provide an impingement cavity between the leading edge wall and the impingement wall and multiple impingement holes are provided in the impingement wall. The impingement holes are spaced laterally across the impingement wall. A method of manufacturing an airfoil includes the steps of depositing multiple layers of powdered metal onto one another, joining the layers to one another with reference to CAD data relating to a particular cross-section of an airfoil, and producing the airfoil. | 10-24-2013 |
| 20140255206 | TURBINE BLADE COOLING CHANNEL FORMATION - Embodiments of the invention relate generally to turbine blades and, more particularly, to the formation of cooling channels on a surface of a turbine blade and turbine blades including such cooling channels. In one embodiment, the invention provides a method of forming a cooling channel along a surface of a turbine blade, the method comprising: applying a first mask material to a first portion of a surface of a turbine blade; forming a first barrier layer atop the first mask material and atop a second portion of the surface of the turbine blade; removing the first mask material and the barrier layer atop the first mask material to expose the first portion of the surface of the turbine blade; and etching the first portion of the surface of the turbine blade to form a cooling channel along the surface of the turbine blade. | 09-11-2014 |
| 20140356188 | TURBINE BLADE AIRFOILS INCLUDING FILM COOLING SYSTEMS, AND METHODS FOR FORMING AN IMPROVED FILM COOLED AIRFOIL OF A TURBINE BLADE - Turbine blade airfoils, film cooling systems thereof, and methods for forming improved film cooled components are provided. The turbine blade airfoil has an external wall surface and comprises leading and trailing edges, pressure and suction sidewalls both extending between the leading and the trailing edges, an internal cavity, one or more isolation trenches in the external wall surface, a plurality of film cooling holes arranged in cooling rows, and a plurality of span-wise surface connectors interconnecting the outlets of the film cooling holes in the same cooling row to form a plurality of rows of interconnected film cooling holes. Each film cooling hole has an inlet connected to the internal cavity and an outlet opening onto the external wall surface. The span-wise surface connectors in at least one selected row of interconnected film cooling holes are disposed in the one or more isolation trenches. | 12-04-2014 |
| 20140369852 | COOLED TURBINE BLADE WITH DOUBLE COMPOUND ANGLED HOLES AND SLOTS - A turbine blade for a gas turbine engine. The turbine blade includes a base having a blade root, a platform, a cooling air inlet, and a base air passageway. The turbine blade also includes an airfoil section adjoined to the base and having an outer wall, an airfoil air passageway, a plurality of trailing edge slots in fluid communication with the airfoil air passageway and a plurality of directional film holes through the outer wall in fluid communication with the airfoil air passageway. The plurality of directional film holes includes a first portion configured to discharge the cooling air toward a tip end, and a second portion configured to discharge the cooling air toward the platform. | 12-18-2014 |
| 20150064019 | Gas Turbine Components with Porous Cooling Features - The present application provides a hot gas path component for use with a gas turbine engine. The hot gas path component may include an airfoil, an internal cooling cavity, and a porous section created by a direct metal laser melting technique. The porous section may be built into the airfoil or the airfoil may be built separately and attached to the airfoil. | 03-05-2015 |
| 20150093251 | Cooling Module Design and Method for Cooling Components of a Gas Turbine System - A cooling arrangement in a gas turbine system ( | 04-02-2015 |
| 20150345302 | TRANSPIRATION-COOLED ARTICLE HAVING NANOCELLULAR FOAM - A transpiration-cooled article includes a body wall that has first and second opposed surfaces. The first surface is adjacent a passage that is configured to receive a pressurized cooling fluid. At least a portion of the body wall includes a nanocellular foam through which the pressurized cooling fluid from the passage can flow to the second surface. The article can be an airfoil that includes an airfoil body that has an internal passage and an outer gas-path surface. At least a portion of the airfoil body includes a nanocellular foam through which cooling fluid from the internal passage can flow to the gas-path surface. | 12-03-2015 |
| 20160069192 | DOUBLE-JET FILM COOLING STRUCTURE AND METHOD FOR MANUFACTURING SAME - A double-jet film cooling structure includes: an injection port, formed on a wall surface facing a high-temperature gas passage; a main passage as a straight round hole to supply cooling medium to the injection port; a pair of branch passages as straight round holes; and communication passages connecting the main passage to the branch passages. The main passage and the branch passages have same constant inner diameters. Each communication passages has an envelope surface obtained by continuously arranging straight round holes each passing a branch point and having the constant inner diameter. Transverse injection angles of the branch passages relative to gas flow along the wall surface are oriented in opposite directions. An angle between axial direction of the main passage and the wall surface is greater than an angle formed between axial direction of branch passage and the wall surface. | 03-10-2016 |
| 20160160655 | CONTROLLING EXIT SIDE GEOMETRY OF FORMED HOLES - A component includes a structural member and an outer wall covering the structural member with a gap between the outer wall and the structural member. The outer wall includes an array of holes each of the array of holes extending from an exterior surface of the outer wall to an interior surface of the outer wall. The outer wall includes an array of recesses on the interior surface of the outer wall, each hole in the array of holes terminating within one of the array of recesses of the outer wall. | 06-09-2016 |
