Class / Patent application number | Description | Number of patent applications / Date published |
416096000 | Blade inserts | 7 |
20080260537 | Turbine Blade with an Impingement Cooling Insert - The invention relates to a turbine blade, particularly a guide blade for a gas turbine, respectively comprising a blade leg, a platform area and a hollow blade for receiving a metal plate-type impact cooling insert which consists of at least two sections which overlap in an overlap area. The aim of the invention is to provide a turbine blade wherein les coolant is required. The two sections thus have a wave-shaped cross-section in the overlap area in order to seal said overlap area. | 10-23-2008 |
20080267784 | Van Wheel of Turbine Comprising a Vane and at Least One Cooling Channel - There is described a vane wheel of a turbine comprising at least one vane, the footing thereof being held on a wheel disk. At least one cooling channel is arranged between the wheel disk and the vane footing. Said vane wheel having a plurality of turbulators is embodied on at least one of the walls of the cooling channel, said turbulators being configured in such a way that the turbulence and thus the heat transfer of a cooling fluid flowing through the cooling channel are increased. | 10-30-2008 |
20090074588 | AIRFOIL WITH COOLING HOLE HAVING A FLARED SECTION - An airfoil is provided for a gas turbine engine. The airfoil may comprise a main body comprising a leading edge having a leading edge outer surface, a trailing edge having a trailing edge outer surface, a suction side having a suction side outer surface and a pressure side having a pressure side outer surface. The main body may further comprise at least one interior cooling passage and a plurality of cooling holes extending from the cooling passage to at least one of the leading edge outer surface, the trailing edge outer surface, the suction side outer surface and the pressure side outer surface. Preferably, at least one of the cooling holes includes a proximal metering section having a first dimension extending transverse to an axis extending in a flow direction of a cooling fluid passing through the one cooling hole, a flared section and an exit opening having a second dimension transverse to the axis which is larger than the first dimension. The flared section is preferably curvilinear as it extends from the proximal metering section towards the exit opening. | 03-19-2009 |
20090110561 | TURBINE ENGINE COMPONENTS, TURBINE ENGINE ASSEMBLIES, AND METHODS OF MANUFACTURING TURBINE ENGINE COMPONENTS - Turbine engine components are provided for use with a plurality of blades, where each blade has an attachment portion defined by a wall and a surface. The components include a disk having a first side, a second side, and an outer radial section with a rim and an overhang. Disk slots extend radially inwardly from the disk rim, and each is configured to receive the attachment portion of a corresponding blade such that when the blade is disposed therein, the wall of the blade attachment portion is substantially flush with an adjacent portion of the first side of the disk and each disk slot defines a cooling air passage with the surface of the blade attachment portion, where the cooling air passage extends axially from the disk first side to the disk second side. A cooling air slot is formed in the disk first side and extends radially outward from the overhang of the disk to the cooling air passage. | 04-30-2009 |
20090136358 | Blade cooling - In order to couple coolant air flow presented through a coolant gallery | 05-28-2009 |
20090162210 | IMPELLER AND COOLING FAN INCORPORATING THE SAME - An impeller ( | 06-25-2009 |
20090169394 | METHOD OF FORMING COOLING HOLES AND TURBINE AIRFOIL WITH HYBRID-FORMED COOLING HOLES - A method of forming a cooling hole in a workpiece that includes the steps of laser-forming a blind, inwardly-tapering transition opening into a first side of the workpiece, and EDM-forming a generally cylindrical through hole to a second, opposing side of the workpiece communicating with the inwardly-tapering transition opening to form a through cooling hole communicating with the first and second sides of the workpiece. An airfoil having cooling holes formed by use of both laser and EDM is also disclosed. | 07-02-2009 |
20090269210 | METHOD FOR FORMING TURBINE BLADE WITH ANGLED INTERNAL RIBS - A die for forming a lost wax ceramic core allows the formation of non-parallel separating spaces between adjacent portions of the core. The core will eventually form cooling channels in an airfoil. The die for forming the core includes a plurality of moving parts having rib extensions. At least some rib extensions are non-parallel to form the non-parallel spaces. The die includes two main die halves that come together to form several of the spaces. Inserts move with those die components and come together to form other spaces. At least one of the inserts contacts surfaces on one of the die halves, such that the non-parallel spaces are formed. | 10-29-2009 |
20090324421 | Turbine Blade - A turbine blade is provided. The turbine blade includes a support structure and a shell which surrounds the support structure and which is connected to and at a distance from the support structure by at least one spacing element. For example the spacing element can be a solder globule in order to form a space through which a cooling medium can flow between the support structure and the shell. A method for the production of a turbine blade having a support structure and a shell which surrounds the support structure and which is connected to and at a distance from the support structure is also provided. The shell is soldered to the support structure in at least at one place of the support structure in order to connect the shell to and at a distance from the support structure. | 12-31-2009 |
20100003142 | AIRFOIL WITH TAPERED RADIAL COOLING PASSAGE - A turbine engine airfoil includes an airfoil structure having an exterior surface and an end portion. A cooling passage extends a length radially within the structure in a direction toward the end portion. The cooling passage provides a convection surface along the length adjacent to the exterior surface. The convection surface includes a generally uniform width along the length. The cooling passage has generally decreasing cross-sectional areas along the length in the direction. | 01-07-2010 |
20100028163 | Injection Molded Component - An intermediate component includes a first wall member, a leachable material layer, and a precursor wall member. The first wall member has an outer surface and first connecting structure. The leachable material layer is provided on the first wall member outer surface. The precursor wall member is formed adjacent to the leachable material layer from a metal powder mixed with a binder material, and includes second connecting structure. | 02-04-2010 |
20100054952 | Turbine Blade - A turbine blade comprising a plurality of ribs arranged one after the other in a cooling channel extending along a leading edge is provided. The plurality of ribs is split into pairs of ribs formed by two ribs arranged in the form of a skating step. | 03-04-2010 |
20100284822 | Turbine Airfoil with a Compliant Outer Wall - A turbine airfoil usable in a turbine engine with a cooling system and a compliant dual wall configuration configured to enable thermal expansion between inner and outer layers while eliminating stress formation in the outer layer is disclosed. The compliant dual wall configuration may be formed a dual wall formed from inner and outer layers separated by a support structure. The outer layer may be a compliant layer configured such that the outer layer may thermally expand and thereby reduce the stress within the outer layer. The outer layer may be formed from a nonplanar surface configured to thermally expand. In another embodiment, the outer layer may be planar and include a plurality of slots enabling unrestricted thermal expansion in a direction aligned with the outer layer. | 11-11-2010 |
20110038734 | Turbine Blade Having a Constant Thickness Airfoil Skin - A turbine blade is provided for a gas turbine comprising: a support structure comprising a base defining a root of the blade and a framework extending radially outwardly from the base, and an outer skin coupled to the support structure framework. The skin has a generally constant thickness along substantially the entire radial extent thereof. The framework and the skin define an airfoil of the blade. | 02-17-2011 |
20110058957 | BLADE FOR A GAS TURBINE - A blade is provided for a gas turbine, especially for the low-pressure turbine of a gas turbine with sequential combustion, and is produced in accordance with a casting process and has a blade airfoil which extends in the radial direction between an inner platform and an outer platform, and in the interior of which extends a cooling passage, bypassing the platforms, and through which flows a cooling medium, especially cooling air, for cooling the blade. In the outer and/or inner platform there are core outlet openings which arise from the use of a casting core and which connect the cooling passage to the outside space and are sealed off by means of a sealing element. Optimum cooling is ensured by the sealing elements being formed and inserted into the core outlet openings so that they align with the wall surface of the cooling passage in a flush manner. | 03-10-2011 |
20110070095 | AEROFOIL STRUCTURE - An aerofoil structure ( | 03-24-2011 |
20110081253 | GAS TURBINE ENGINE BALANCING - An apparatus and method for balancing a gas turbine engine rotor includes a plurality of balancing weights adapted to be selectively attached to at least one of inlets or outlets of a cooling passage of the rotor. The weights include cooling access which permits coolant to communicate with the cooling passage. | 04-07-2011 |
20110103970 | STEAM TURBINE WITH RELIEF GROOVE ON THE ROTOR - A steam turbine is provided having a relief groove which is arranged in the region of the equalizing piston and extends in the circumferential direction of the rotor. The relief groove, with regard to an inlet passage, is arranged in the axial upstream direction so that it is arranged on the rotor outside a region in which the steam flow enters the bladed flow path via the inlet passage. The relief groove, with regard to the first blade row, is arranged in a region in which the greatest thermal stresses can arise in the rotor. As an option, the relief groove has a cover for reducing vortex flows, and also devices for reducing heating of the groove or devices for active cooling. The steam turbine allows an increased number of risk-free running up and running down operations of the steam turbine with minimum detriment to the turbine performance. | 05-05-2011 |
20110135496 | COOLING OF THE TIP OF A BLADE - A turbine blade including an open cavity at its distal tip, the cavity being defined by a bottom wall and a side wall extending along the perimeter of the distal tip in an extension of the upper and lower walls of the blade, the side wall of the cavity including an opening in a vicinity of the leading edge of the blade opening into the cavity. A deflector extends at least in the middle portion of the cavity between the leading edge and the trailing edge. | 06-09-2011 |
20110158819 | INTERNAL REACTION STEAM TURBINE COOLING ARRANGEMENT - A rotor of a turbomachine includes a rotor drum located at a central axis and a plurality of buckets secured to the rotor drum. A first reaction stage includes axial entry dovetailed buckets. An axial passage for cooling flow is provided along a mating surface between the bucket dovetail and the dovetail slot in the rotor drum. Cool steam at taken between a first stage bucket and a second stage nozzle and passed through the axial passage to a low pressure sink at an upstream end of the rotor. | 06-30-2011 |
20110200448 | TURBINE DISK AND BLADE ARRANGEMENT - A turbine disk and blade arrangement for a gas turbine engine has a plurality of turbine blades mounted circumferentially around a disk. Each blade has a fir tree root which provides, on circumferentially spaced sides thereof, a series of fore-to-aft-extending projections and grooves. The disk has a plurality of circumferentially spaced, radially extending posts which define fir tree recesses therebetween. Each fir tree recess also provides, on circumferentially spaced sides thereof, a series of fore-to-aft-extending projections and grooves. Each fir tree root is slidable into a respective fir tree recess with the projections and grooves on facing sides of the fir tree root and the fir tree recess interengaging with each other. On each pair of facing sides of the fir tree root and the fir tree recess, one or more of the projections of the fir tree roots and/or the fir tree recesses are truncated to form fore-to-aft-extending cooling passages between facing sides of the fir tree roots and fir tree recesses. The arrangement has lock plates which cover forward surfaces of the posts and of the fir tree roots and/or cover rearward surfaces of the posts and of the fir tree roots. The lock plates and the posts define feed channels therebetween which feed cooling air to the cooling passages, each cooling passage extending to a respective feed channel. | 08-18-2011 |
20110223036 | BLADE FOR A GAS TURBINE - The blade for a gas turbine includes a blade airfoil having a leading edge and a trailing edge and extending in the blade longitudinal direction up to a blade tip, and at the blade tip the blade airfoil merges into a shroud segment, wherein on the shroud segment a first rib, projecting upwards, is arranged in the flow direction, extending transversely to the flow direction, and upstream of the first rib, in the region of the leading edge of the blade airfoil, a winglet is formed on the shroud segment for guiding of the hot gas flow in this region. With such a blade, a longer service life is achieved by provision being made for direct cooling of the winglet. | 09-15-2011 |
20110250078 | TURBINE BUCKET HAVING A RADIAL COOLING HOLE - A turbine bucket is provided and includes a shank interconnectable with a rotor and formed to accommodate coolant therein and an airfoil blade coupled to a radially outward portion of the shank and including a body formed to define a substantially radially extending cooling hole therein, which is disposed to be solely receptive of the coolant accommodated within the shank for removing heat from the body, the cooling hole being further defined as having a substantially non-circular cross-sectional shape at a predefined radial position of the body. | 10-13-2011 |
20110268582 | COOLED BLADE FOR A GAS TURBINE - A blade for a gas turbine, is provided and includes a blade airfoil which extends in a blade longitudinal direction and at the lower end merges into a shank which terminates in a blade root for fastening the blade on a blade carrier, particularly on a rotor disk. Devices for cooling the blade, which are supplied with a cooling medium, especially cooling air, via a feed hole arranged on the shank at the side, are arranged inside the blade airfoil. In the region of the feed hole provision is made for a planar stiffening element which reaches beyond the immediate region of the feed hole for reducing peaks of mechanical stress. | 11-03-2011 |
20120034101 | TURBINE BLADE SQUEALER TIP - A turbine blade having a squealer tip coupled to a radially outer end of the turbine blade that is usable in a gas turbine engine is disclosed. The squealer tip may require less cooling air and may therefore be more efficient than conventional configurations. The squealer tip may be formed from one or more materials such as oxide dispersion strengthened alloys and FeCrAl alloys. The squealer tip may be formed from a plurality of segmented tips extending radially outward and spaced apart from each other. For example, the squealer tip may be formed from two rails extending radially outward and spaced apart from each other. The two rails may be formed from outer and inner rails that each form a continuous ring. The squealer tip may be attached to the tip with a transient liquid phase bond or an additive manufacturing process, such as, a selective laser melting process. | 02-09-2012 |
20120107133 | DEICING DEVICE FOR PROPFAN-TYPE PROPELLER BLADES - A deicing device for propfan-type aircraft propulsion unit blades, wherein the propulsion unit includes a turbomachine that drives in rotation at least one rotor including a plurality of blades arranged around an annular crown moving with the blades, which forms with its outer wall part of an outer envelope of the propulsion unit, the outer envelope being subjected to atmospheric conditions outside the propulsion unit, the turbomachine generating a flow of hot gases that exit via an annular vein, which is concentric with the moving annular crown, and defined for part of its surface by the inner wall of the moving annular crown. The deicing device includes: a mechanism capturing thermal energy from the annular vein, within the moving annular part; a mechanism transferring thermal energy towards the rotor blades; and a mechanism distributing the thermal energy onto at least a part of the surface of the blades. | 05-03-2012 |
20120121434 | TURBINE BLADE ASSEMBLY INCLUDING A DAMPER - A damper for a turbine rotor assembly of a gas turbine engine is disclosed. The damper may have a forward plate. The damper may further have an aft plate including a larger surface area than the forward plate. The aft plate may have at least one aperture for regulating a flow of gas through the aft plate. The damper may also have a longitudinal structure connecting the forward plate and the aft plate. | 05-17-2012 |
20120156054 | TURBINE COMPONENT WITH NEAR-SURFACE COOLING PASSAGE AND PROCESS THEREFOR - A process for creating a near-surface cooling passage in an air-cooled turbomachine component. The process entails forming a channel in a surface of a surface region of the component so that the channel is open at the surface and fluidically connected to a first cooling passages within the component. A metallic layer is then deposited on the surface and over the channel without filling the channel. The metallic layer closes the channel at the surface of the surface region to define therewith a second cooling passage within the component that is fluidically connected to the first cooling passages. A coating system is then deposited on the metallic layer to define an outermost surface of the component. The second cooling passage is closer to the outermost surface of the component than the first cooling passages. | 06-21-2012 |
20120177503 | COMPONENT COOLING CHANNEL - A cooling channel ( | 07-12-2012 |
20120201694 | TURBINE BLADE - Disclosed is a turbine blade capable of being cooled by a coolant gas supplied to a hollow region, wherein a plurality of meandering flow paths that guide the coolant gas between the suction wall surface and the pressure wall surface while causing the coolant gas to repeatedly meander are continuously arranged from the hub side toward the tip side of the turbine blade, and the meandering flow paths adjacent to each other cause the coolant gas to meander in different repetitive patterns. | 08-09-2012 |
20120230837 | COOLING SLEEVE - A cooling sleeve includes a first end that extends to a second end, and at least one coolant inlet member. The cooling sleeve also includes a second sleeve portion. The second sleeve portion includes a first end section that extends to a second end section, and a coolant outlet member. The first and second ends of the first sleeve portion are operatively connected to corresponding ones of the first and second end sections of the second sleeve portion to form a continuous cooling zone. The coolant passing into the inlet member circulates through the cooling zone to create a localized temperature reduction. | 09-13-2012 |
20120244010 | ELECTRODE AND ELECTROCHEMICAL MACHINING PROCESS FOR FORMING NON-CIRCULAR HOLES - An electrode for an electrochemical machining process is provided. The electrode includes an electrically conductive member defining at least one passage and an insulating coating partially covering a side surface of the electrically conductive member. The insulating coating does not cover at least one of first and second exposed sections of the electrically conductive member, where the first and second exposed sections are separated by approximately 180 degrees and extend substantially along a longitudinal axis of the electrically conductive member. The insulating coating also does not cover an exposed front end of the electrically conductive member. An electrochemical machining method is also provided, for forming a non-circular hole in a workpiece using the electrode. | 09-27-2012 |
20130108466 | ASYMETRICALLY SLOTTED ROTOR FOR A GAS TURBINE ENGINE | 05-02-2013 |
20130108467 | TURBINE WHEEL FOR A TURBINE ENGINE | 05-02-2013 |
20130156598 | BLADE FOR A TURBO MACHINE - A blade for a turbomachine, for example a gas turbine, is provided. The blade is arranged on a turbine rotor of the gas turbine. The blade includes a root portion having two narrow sides and two broad sides, a cooling air supply passage in the root portion, and a cooling air bleed which is arranged in the root portion and is in fluid connection with the cooling air supply passage. The cooling air bleed includes a nozzle on one of the narrow sides of the root portion, wherein the nozzle is formed by a hole and wherein an axial direction of the hole is inclined upward between 92° and 135° with respect to a longitudinal direction of the blade. | 06-20-2013 |
20130156599 | TURBINE BLADE FOR A GAS TURBINE - A turbine blade for a gas turbine is provided. The quantity of coolant flowing off the rear edge thereof is set relatively simply and exactly directly upon casting the turbine blade, without reworking the cast turbine blade with respect to the setting of coolant consumption being necessary. Raised areas are situated on the inner surfaces of the intake side wall or pressure side wall, between which a throttle element is present, by means of which the quantity of coolant flowing out is set. This arrangement allows a core tool to be produced simply, by means of which the casting cores required for casting the turbine blade are produced having the desired precision in great quantities. | 06-20-2013 |
20130209268 | GAS TURBINE BLADE - A gas turbine blade including a root and an air foil with a leading edge and a trailing edge, a cooling air channel system extending from a cooling air opening in the root via a winding serpentine channel to a trailing edge channel at the trailing edge including an air outlet at the trailing edge is provided. For efficiently cooling the trailing edge of the blade it is proposed that the cooling air channel system includes an air bypass channel connecting the cooling air opening in the root with the trailing edge channel bypassing the serpentine channel. | 08-15-2013 |
20130243606 | TURBINE BLADE TIP COOLING - A turbine blade includes a blade portion, the blade portion comprising a tip outer wall and a trailing edge, an internal cooling circuit, the internal cooling circuit being configured for directing cooling air within the blade portion, and a tip trailing edge slot positioned adjacent to the tip outer wall and the trailing edge, the tip trailing edge slot being fluidly connected to the internal cooling circuit. The tip outer wall is recessed at the tip trailing edge slot such that the tip outer wall is not provided over the trailing edge slot, thereby allowing cooling air to flow from the cooling circuit, into the trailing edge slot, and radially over the tip outer wall. | 09-19-2013 |
20130294927 | SYSTEM AND METHOD FOR COVERING A BLADE MOUNTING REGION OF TURBINE BLADES - A system includes a cover segment configured to mount in first and second grooves circumferentially along a blade mounting region of a turbine rotor. The cover segment includes a cover body and at least one fastener. The cover body includes first and second lips extending along first and second circumferential portions of the cover body at an offset from one another. The first lip is configured to mount circumferentially along the first groove and the second lip is configured to mount circumferentially along the second groove. The at least one fastener is configured to retain the cover body to the turbine rotor or at least one blade mounted in the blade mounting region. | 11-07-2013 |
20140044555 | TRAILING EDGE COOLING CONFIGURATION FOR A GAS TURBINE ENGINE AIRFOIL - An airfoil for a gas turbine engine includes pressure and suction surfaces provided by pressure and suction walls that extend in a radial direction and are joined at a leading edge and a trailing edge. A cooling passage is arranged between the pressure and suction walls and extends to the trailing edge. Elongated pedestals are arranged in the cooling passage and interconnect the pressure and suction walls. The elongated pedestals are spaced apart from one another in the radial direction and extend from a plane to the trailing edge. A metering pedestal includes at least a portion that is arranged between the plane and the trailing edge. The portion is provided between adjacent elongated pedestals in the radial direction. | 02-13-2014 |
20140056716 | BICAST TURBINE ENGINE COMPONENTS - A turbine blade assembly includes a turbine blade having a pressure sidewall and an opposed suction sidewall and a first snubber assembly associated with one of the pressure sidewall and the suction sidewall. The first snubber assembly includes a first base portion extending outwardly from the one of the pressure sidewall and the suction sidewall, and a first snubber portion. The first base portion is integrally cast with the turbine blade and includes first connection structure. The first snubber portion is bicast onto the first base portion and includes second connection structure that interacts with the first connection structure to substantially prevent separational movement between the first base portion and the first snubber portion. | 02-27-2014 |
20140072447 | CERAMIC AND REFRACTORY METAL CORE ASSEMBLY - A core assembly for forming a cast component includes a refractory metal core and a ceramic core element. The refractory metal core includes first and second ends and sides extending from the first end to the second end. The ceramic core element includes a slot positioned between first and second lands, each land having an inner surface facing the slot and an adjacent outer surface. The first end of the refractory metal core is secured within the slot with an adhesive, and the refractory metal core extends from the ceramic core element in both a longitudinal and a transverse direction. The slot, lands, and refractory metal core form a core assembly providing access paths to the sides of the refractory metal core. Surplus adhesive is removed from the refractory metal core via the access paths. Investment casting provides the component with an internal passage and an internal cooling circuit. | 03-13-2014 |
20140093386 | COOLED TURBINE BLADE WITH INNER SPAR - 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 inner spar cooling fins extending from the inner spar to the skin, and a plurality of trailing edge cooling fins extending from the pressure side of the skin to the lift side of the skin aft of the inner spar. | 04-03-2014 |
20140099210 | SYSTEM FOR GAS TURBINE ROTOR AND SECTION COUPLING - A system, including a first turbomachine rotor disk, a first annular protrusion with a first spline coupled to the first turbomachine rotor disk, a second turbomachine rotor disk, and a second annular protrusion with a second spline coupled to the second turbomachine rotor disk, wherein the first and second splines are coupled to one another. | 04-10-2014 |
20140112798 | GAS TURBINE AND TURBINE BLADE FOR SUCH A GAS TURBINE - A gas turbine includes a rotor concentrically surrounded by a casing, with an annular hot gas channel axially extending between the rotor and the casing. The rotor is equipped with a plurality of blades, which are arranged on the rotor in an annular fashion. Each of the blades is mounted with a root in a respective axial slot on a rim of the rotor radially extending with an airfoil into said hot gas channel and adjoining with an axially oriented root surface to an annular rim cavity. Cooling means are provided at the root of each of said blades to receive cooling air being injected into said rim cavity through stationary injecting means. An optimized cooling is achieved by providing the root surface to be an essentially plane surface and the cooling means including a scoop for capturing and redirecting at least part of the injected cooling air, which scoop is designed as a recess with respect to the root surface. | 04-24-2014 |
20140119943 | TURBINE ROTOR ASSEMBLY - A turbine rotor assembly may include a turbine rotor with a plurality of turbine blade slots and a plurality of turbine blades. A root structure of each turbine blade may include a portion shaped to be received in a corresponding turbine blade slot of the rotor. The rotor assembly may also include a seal plate attached to the forward end of the rotor. The seal plate may extend upwards from a first end below the inner end of the blade slots to a second end located between an outermost lobe of the blade slots and the outer rim of the rotor. | 05-01-2014 |
20140140859 | UBER-COOLED MULTI-ALLOY INTEGRALLY BLADED ROTOR - Uber-cooled multi-alloy integrally bladed rotors (IBR) are made having blades with internal cooling passages with a cavity in the root portion attached to a disk having a protrusion on the periphery of the disk. The blades are put on the protrusion and the blade and disk are forced together, followed by locally heating the blade cavity/disk protrusion to a temperature between the blade and disk material softening temperatures, causing the protrusion to deform against the blade cavity, and holding them in place until bonding occurs. Subsequent to bonding the portion of the blade that defines the cavity is removed. | 05-22-2014 |
20140147287 | TRAILING EDGE AND TIP COOLING - A gas turbine engine component includes an airfoil extending radially from a root section to a tip section and having a trailing edge cooling passageway and first, second and third flow dividers in the cooling passageway. The first, second and third flow dividers have longitudinal axes that are angled based upon a position of the flow divider relative to the tip section of the airfoil. | 05-29-2014 |
20140161625 | TURBINE COMPONENT HAVING COOLING PASSAGES WITH VARYING DIAMETER - Systems and devices configured to cool turbine components in a turbine by passing a cooling flow through the turbine component via a cooling passage with a variable diameter are disclosed. In one embodiment, a turbine component includes: at least one elongated cooling passage extending from a root of the bucket to a tip of the bucket, wherein the elongated cooling passage has a variable diameter along a length of the bucket. | 06-12-2014 |
20140161626 | METHOD FOR MANUFACTURING AN OXIDE/OXIDE COMPOSITE MATERIAL TURBOMACHINE BLADE PROVIDED WITH INTERNAL CHANNELS - An oxide/oxide composite material turbomachine blade including a fiber reinforcement obtained by weaving a first plurality of threads and a second plurality of threads, with the threads of said first plurality of threads being arranged in successive layers and extending in the longitudinal direction of the fiber blank corresponding to the longitudinal direction of the blade is disclosed. The reinforcement is densified by a matrix, with the blade further including one or several internal channels having a coiled shape extending in the longitudinal direction of the blade. | 06-12-2014 |
20140169981 | UBER-COOLED TURBINE SECTION COMPONENT MADE BY ADDITIVE MANUFACTURING - A gas turbine airfoil having internal cooling passages is formed by additive manufacturing. Layers of superalloy powder are fused by an energy beam using a two-dimensional pattern providing unmelted areas forming passageways therein. Layers of the powder are added and fused using sufficient two-dimensional patterns to form the entire airfoil with the desired pattern of internal cooling passages. After completion of the formation of the airfoil, it may be hot isostatic pressed, directionally recrystallized, bond coated, and covered with a thermal barrier layer. | 06-19-2014 |
20140178207 | TURBINE BLADE - A rotor stage of a turbine has a rotational axis, a shroud and radially inward thereof a turbine blade defined partly by a pressure side wall, a suction side wall and a tip portion. The tip portion has a pressure side tip rib and a tip cavity floor defining a tip cavity. The pressure side tip rib has a width W | 06-26-2014 |
20140193272 | Gas Turbine Engine Cooling Systems and Methods Incorporating One or More Cover Plate Assemblies Having One or More Apertures Therein - Turbine cooling systems and methods are disclosed herein. The turbine cooling system may include a rotor disk having a bucket attached thereto. The bucket may include a shank cavity. The turbine cooling system may also include a cover plate positioned at least partially about the rotor disk and the shank cavity. The cover plate may be configured to separate a first flow of cooling fluid from a second flow of cooling fluid. At least one aperture may be disposed in the cover plate about the shank cavity. The at least one aperture may be configured to provide the first flow of cooling fluid to the shank cavity. | 07-10-2014 |
20140193273 | INTERIOR CONFIGURATION FOR TURBINE ROTOR BLADE - An interior cooling configuration formed within an airfoil of a blade of a combustion turbine engine is provided. The interior cooling configuration may include a first flow passage and a second flow passage that have a side-by-side configuration for a segment, and multiple lateral crossover passages extending between and fluidly connecting the first flow passage to the second flow passage. The crossover passages may be staggered. | 07-10-2014 |
20140212298 | GAS TURBINE ENGINE TURBINE BLADE TIP COOLING - An airfoil for a gas turbine engine includes pressure and suction walls spaced apart from one another and joined at leading and trailing edges to provide an airfoil having an exterior surface that extends in a radial direction to a tip. A camber line at the tip extends from the leading edge to the trailing edge. Pressure and suction side shelves are arranged in the exterior surface on opposing sides of the camber line respectively in the pressure and suction side walls. A plateau is proud of and separates the pressure and suction side shelves. The plateau is arranged along the camber line and extends to the leading edge. | 07-31-2014 |
20140219813 | GAS TURBINE ENGINE SERPENTINE COOLING PASSAGE - A gas turbine engine component includes a structure having a cooling passage providing upstream and downstream portions separated from one another by an inner wall and fluidly connected by a bend. The downstream portion includes an outer wall opposite the inner wall to provide a downstream region extending between the inner and outer walls. A turbulence promoter extends from the outer wall adjacent to the bend in the downstream portion. The turbulence promoter is absent from a stagnation region adjoining the inner wall adjacent to the bend in the downstream portion | 08-07-2014 |
20140219814 | FILM-COOLED TURBINE BLADE FOR A TURBOMACHINE - A turbine blade for a turbomachine has an outer wall which delimits an inner cavity. Cooling fluid flows in the inner cavity. A through duct is arranged in the outer wall through which the cooling fluid flows from the inner cavity to an outside of the turbine blade. The through duct is inclined with respect to a trailing edge of the turbine blade, wherein a marginal portion of an entrance of the through duct is designed on an upstream side to be sharp-edged in relation to other marginal portions of the entrance such that a separation zone of a cooling fluid flow is formed in the through duct. A pair of swirls builds up in the through duct, wherein velocity vectors of the cooling fluid flow between the swirl centers point toward a downstream side of the through duct. | 08-07-2014 |
20140271225 | INTERIOR COOLING CIRCUITS IN TURBINE BLADES - A rotor blade comprising an airfoil portion and a root portion, and an internal cooling circuit having flow passages in the root portion and the airfoil portion, wherein the internal cooling circuit includes: a first flow passage; and a non-integral plug. The plug may include a plug channel configured to correspond to a desired level of coolant flow through the first cooling passage. The plug may be connected to the rotor blade in a fixed blocking position relative to the first flow passage. | 09-18-2014 |
20140294597 | COOLING FOR THE RETAINING DOVETAIL OF A TURBOMACHINE BLADE - A turbomachine assembly in which a foil is configured to cover mainly one of bulbs of a disc and to be held, radially with respect to the disc, by the bulb of the disc and a pocket for a blade that can collaborate therewith, when these two are effectively collaborating, and the bulb of the disc includes at least one longitudinal cavity configured to form, with the foil, when the foil is covering the bulb of the disc, a secondary passage through which a secondary cooling air flow can pass. | 10-02-2014 |
20140314581 | METHOD FOR FORMING SINGLE CRYSTAL PARTS USING ADDITIVE MANUFACTURING AND REMELT - A method of forming a metal single crystal turbine component with internal passageways includes forming a polycrystalline turbine blade with internal passageways by additive manufacturing and filling the passageways with a core ceramic slurry. The ceramic slurry is then treated to harden the core and the turbine component is encased in a ceramic shell which is treated to form a ceramic mold. The turbine component in the mold is then melted and directionally solidified in the form of a single crystal. The outer shell and inner ceramic core are then removed to form a finished single crystal turbine component with internal passageways. | 10-23-2014 |
20140322027 | WIND TURBINE ROTOR BLADE AND A METHOD FOR DEICING A WIND TURBINE ROTOR BLADE - There is provided a wind power installation rotor blade comprising a rotor blade nose, a rotor blade trailing edge, a rotor blade root region for fixing the rotor blade to a hub of a wind power installation and a rotor blade tip. The rotor blade extends from the rotor blade root region along a longitudinal axis to the rotor blade tip. The rotor blade further has an air distribution unit having an adjusting member for directing an air flow into the rotor blade nose region and/or a rotor blade trailing edge region. | 10-30-2014 |
20140348664 | IMPINGEMENT-COOLED TURBINE ROTOR - An integral turbine includes a forward hub section and an aft hub section. The forward hub section and the aft hub section are metallurgically coupled to one another along an annular interface that resides within a plane generally orthogonal to a rotational axis of the axially-split turbine. The turbine further includes an airfoil blade ring metallurgically coupled to a radial outer surface of the coupled forward and aft hub sections and an impingement cavity formed within an interior portion of the coupled forward and aft hub sections. The impingement cavity includes an interior surface that is positioned proximate to the radial outer surface of the coupled forward and aft hub sections. Further, an impingement cooling air flow impinges against the interior surface of the impingement cavity to provide convective and conductive cooling to the radial outer surface of the coupled forward and aft hub sections. | 11-27-2014 |
20140356186 | Method for Manufacturing Gas Turbine Blade, and Gas Turbine Blade - This method is a method for manufacturing a gas turbine blade, including:
| 12-04-2014 |
20140356187 | DE-ICING OF A WIND TURBINE BLADE - A heating assembly for a wind turbine: generator, the assembly comprising: a heat reservoir mounted within a blade of the wind turbine generator; a heat source for supplying heat to the heat reservoir; a plurality of thermal conductors projecting front said heat reservoir to a surface of said blade. | 12-04-2014 |
20150023800 | GAS TURBINE ARRANGEMENT ALLEVIATING STRESSES AT TURBINE DISCS AND CORRESPONDING GAS TURBINE - A turbine arrangement is provided, particularly a gas turbine arrangement, having at least one rotor blade and a turbine disc, the rotor blade having a root portion, the turbine disc having at least one slot in which the root portion of the rotor blade is secured. The slot has a plurality of opposite pairs of slot lobes and a plurality of opposite pairs of slot fillets, and a slot bottom of the slot. The slot bottom is arranged to have a first convex surface section. Furthermore the root portion of the rotor blade has a root bottom with a first concave surface section corresponding to the first convex surface section of the slot bottom. Additionally, the first convex surface section is pierced by an outlet of a cooling duct through the turbine disc. | 01-22-2015 |
20150050159 | DUAL ELEMENT TURBINE BLADE - A turbine blade includes a core element having a base portion, a tip portion, and an intermediate portion extending between the base portion and the tip portion. The intermediate portion includes a non-uniform cross-section and is a high-strength fiber material. The turbine blade further includes a shell disposed around the core element, and the volume between the core element and the shell forms a void. | 02-19-2015 |
20150050160 | ROTOR SHAFT FOR A TURBOMACHINE - A rotor shaft adapted to rotate about a rotor axis thereof. The rotor shaft includes a rotor cavity configured concentrically to the rotor axis inside the rotor shaft. The rotor shaft further includes a plurality of cooling bores extending radially outward from the rotor cavity to feed cooling air into an internal cooling system in a blade. Each cooling bore includes a bore inlet portion and a distal bore outlet portion. The respective bore inlet portion ends in a plateau, projecting above the outer circumference contour of the rotor cavity. Thus, cooling bore inlets are shifted to a low stress area and the lifetime of the rotor is improved. | 02-19-2015 |
20150098835 | 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. | 04-09-2015 |
20150110639 | TURBINE BUCKET INCLUDING COOLING PASSAGE WITH TURN - Turbine frequency tuning, fluid dynamic efficiency, and performance can be improved using a particular profile for a turn of a cooling passage in an airfoil. By blending aspects of baseline and bulb contours into a blended turn with a non-uniform profile, mechanical and/or thermal stress can be reduced in the turn and in an airfoil including the turn, particularly on an outflow side of the turn. Stresses on the airfoil can be reduced using a turn profile that is a blend of a baseline profile and a bulb profile and that can be described by the airfoil core profile. | 04-23-2015 |
20150110640 | TURBINE BUCKET HAVING SERPENTINE CORE - Various embodiments of the invention include turbine buckets and systems employing such buckets. Various particular embodiments include a turbine bucket having: a base; and an airfoil connected with the base at a first end of the airfoil, the airfoil including: a casing having: 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, the casing including an aperture on the leading edge; and a core within the casing, the core having a serpentine shape for supporting the casing and a leading edge passage fluidly connected with the aperture on the leading edge of the casing. | 04-23-2015 |
20150118062 | BUCKET ASSEMBLY FOR USE IN A TURBINE ENGINE - An assembly for use in a turbine engine is provided. The assembly includes a first bucket, a second bucket circumferentially adjacent to the first bucket, a shroud extending between the first and second buckets, and an aerodynamic shell substantially encircling the shroud such that a cavity is formed between the aerodynamic shell and the shroud. | 04-30-2015 |
20150132147 | TURBINE AIRFOIL WITH LATERALLY EXTENDING SNUBBER HAVING INTERNAL COOLING SYSTEM - A turbine airfoil usable in a turbine engine and having at least one snubber with a snubber cooling system positioned therein and in communication with an airfoil cooling system is disclosed. The snubber may extend from the outer housing of the airfoil toward an adjacent turbine airfoil positioned within a row of airfoils. The snubber cooling system may include an inner cooling channel separated from an outer cooling channel by an inner wall. The inner wall may include a plurality of impingement cooling orifices that direct impingement fluid against an outer wall defining the outer cooling channel. In one embodiment, the cooling fluids may be exhausted from the snubber, and in another embodiment, the cooling fluids may be returned to the airfoil cooling system. Flow guides may be positioned in the outer cooling channel, which may reduce cross-flow by the impingement orifices, thereby increasing effectiveness. | 05-14-2015 |
20150345300 | COOLING STRUCTURE FOR STATIONARY BLADE - A cooling structure for a stationary blade is provided. The cooling structure may include a first chamber in an endwall of the stationary blade directing a first cooling fluid from the stationary blade to a first cooling circuit, and a second chamber in the endwall of the stationary blade directing a second cooling fluid from the stationary blade to a second cooling circuit different than the first cooling circuit. The first cooling fluid has a lower temperature than the second cooling fluid. | 12-03-2015 |
20150354369 | GAS TURBINE ENGINE AIRFOIL PLATFORM COOLING - An airfoil for a gas turbine engine includes an airfoil that extends from a platform that has first and second circumferential sides that respectively extend to first and second circumferential edges. The first circumferential side has a tapered surface at a first angle relative to a flow path surface. The second circumferential surface has a cooling hole that extends toward the second lateral edge at a second angle relative to the flow path surface. The tapered surface and the cooling hole are axially aligned with one another. | 12-10-2015 |
20150377395 | CONNECTING PIPING AND STEAM TURBINE SYSTEM - The present invention is provided with: a cylindrical main body section that is provided with at least one annular rib on the outer circumferential surface; and a deforming section that is configured from a bellows that can be deformed in an axis line direction and diameter direction. | 12-31-2015 |
20160047356 | ROTOR BLADE OF A WIND TURBINE - The invention relates to a rotor blade of a wind turbine, having a rotor blade nose, a rotor blade rear edge, a rotor blade root region for the attachment of the rotor blade to a hub of the wind turbine, a rotor blade tip, wherein the rotor blade extends from the rotor blade root region along a longitudinal direction to the rotor blade tip and the rotor blade internally comprises at least a first cavity facing the rotor blade nose and a second cavity facing the rotor blade rear edge, and the first cavity is heated by a first, and the second cavity is heated by a second heating means, in order to heat the rotor blade nose or the rotor blade rear edge respectively. In addition, it is suggested that the rotor blade have a rear edge segment disposed in the region of the rotor blade rear edge up to the root region, wherein the rear edge segment is designed having several parts having at least two segment sections. | 02-18-2016 |
20160069194 | TURBINE BLADES AND METHODS OF FORMING TURBINE BLADES HAVING LIFTED RIB TURBULATOR STRUCTURES - The present disclosure provides various embodiments of cooling circuits, turbine blades with cooling circuits, and methods of forming such turbine blades, having raised rib turbulator structures, which may be used in gas turbine engines. In one exemplary embodiment, a cooling circuit for directing a flow of fluid is disclosed, the cooling circuit includes a cooling circuit wall and a plurality of raised turbulator ribs, each turbulator rib of the plurality of raised turbulator ribs being spaced apart from the cooling circuit wall to allow the fluid to flow between the cooling circuit wall and the plurality of turbulator ribs. | 03-10-2016 |
20160090847 | COOLING SCHEME FOR A TURBINE BLADE OF A GAS TURBINE - A turbine blade of a gas turbine includes a radially extending airfoil with a suction side and pressure side, which extend each in axial direction between a leading edge and a trailing edge of the airfoil. The leading edge is cooled by means of impingement cooling with rows of radially distributed jets of a cooling medium impinging on the inner side of the leading edge. The row of radially distributed jets is generated at an internal web, which divides the hollow interior of the airfoil into first and second cavities, with the second cavity being arranged at the leading edge. An enhanced cooling is achieved by the internal web that includes two rows of radially distributed cooling medium supply holes, through which cooling medium enters the second cavity in form of impinging jets. The cooling medium supply holes are oriented such that the directions of the jets of one row cross the directions of the jets of the other row. | 03-31-2016 |
20160123151 | STEAM TURBINE ROTOR - The invention relates to a steam turbine rotor wherein the inter blade region rotor surface, the feed region rotor surface, the piston region rotor surface and the stress relief groove rotor surface of the rotor are configured and arranged as steam exposed surfaces during normal operation of the steam turbine rotor. The steam turbine rotor has a thermal barrier coating on at least the piston region rotor surface. | 05-05-2016 |
20160146024 | HYBRID BONDED TURBINE ROTORS AND METHODS FOR MANUFACTURING THE SAME - Hybrid bonded turbine rotors and methods for manufacturing the same are provided. A hybrid bonded turbine rotor comprises a turbine disk and a plurality of turbine blades each metallurgically bonded to a corresponding raised blade attachment surface of a plurality of raised blade attachment surfaces of the turbine disk to define a bond plane located at a selected radial position. Turbine disk has a rim portion comprising a live rim of circumferentially continuous material and a plurality of live rim notches in an outer periphery of the turbine disk alternating with the plurality of raised blade attachment surfaces defining the outer periphery. The selected radial position is outboard of the live rim. Each pair of adjacent turbine blades defines a shank cavity therebetween. The shank cavity extends radially outwardly from the live rim and includes a live rim notch disposed below the bond plane and above the live rim. | 05-26-2016 |
20160169052 | ROTATING GAS TURBINE BLADE AND GAS TURBINE WITH SUCH A BLADE | 06-16-2016 |
20160376890 | TURBINE | 12-29-2016 |
20170234136 | TURBINE BLADE AND TURBINE | 08-17-2017 |
20220136393 | CMC VANE ARC SEGMENT WITH CANTILEVERED SPAR - A vane arc segment includes a hollow ceramic airfoil section that has walls that define an internal cavity, a trailing edge, a leading edge, a pressure side, and a suction side. A structural spar piece supports the hollow ceramic airfoil section. The spar piece includes a first spar platform, a hollow leg that has a proximal end at the spar platform and an opposed distal end. The hollow leg extends into the internal cavity of the hollow ceramic airfoil section and has an outer surface with raised bearing pads. Multiple ones of the bearing pads contact the walls of the hollow ceramic airfoil section to support the hollow ceramic airfoil section. A second spar platform is secured with the distal end of the leg. | 05-05-2022 |
20100247327 | RECESSED METERING STANDOFFS FOR AIRFOIL BAFFLE - An internally cooled airfoil comprises an airfoil body, a baffle and a plurality of standoffs. The airfoil body is shaped to form leading and trailing edges, and pressure and suction sides surrounding an internal cooling channel. The baffle is disposed within the internal cooling channel and comprises a liner body having a perimeter shaped to correspond to the shape of the internal cooling channel and to form a cooling air supply duct. The baffle includes a plurality of cooling holes extending through the liner body to direct cooling air from the supply duct into the internal cooling channel. The standoffs maintain minimum spacing between the liner body and the airfoil body. In one embodiment, the standoffs are recessed into a surface of either the baffle or the airfoil body. In another embodiment, the standoffs are elongated to meter flow between the liner body and the airfoil body. | 09-30-2010 |
20120003103 | TURBINE ROTOR ASSEMBLY - A turbine rotor assembly ( | 01-05-2012 |
20120034100 | RECESSED METERING STANDOFFS FOR AIRFOIL BAFFLE - An internally cooled airfoil comprises an airfoil body, a baffle and a plurality of standoffs. The airfoil body is shaped to form leading and trailing edges, and pressure and suction sides surrounding an internal cooling channel. The baffle is disposed within the internal cooling channel and comprises a liner body having a perimeter shaped to correspond to the shape of the internal cooling channel and to form a cooling air supply duct. The baffle includes a plurality of cooling holes extending through the liner body to direct cooling air from the supply duct into the internal cooling channel. The standoffs maintain minimum spacing between the liner body and the airfoil body. The standoffs are recessed into a surface of either the baffle or the airfoil body such that a height of the standoffs is greater than the spacing. | 02-09-2012 |
20120244009 | INSERTS FOR TURBINE COOLING CIRCUIT - In one embodiment, an insert for a turbine cooling circuit includes: a radial cooling passage for receiving downstream fluid; an axial passage extending from the radial cooling passage within a lower portion of the insert; and a plurality of radial passages extending from the axial passage, each radial passage extending to a bottom of a partially circumferential dovetail slot of the insert. In another embodiment, an insert for a turbine cooling includes: a plurality of radial passages, each radial passage extending from a bottom of a partially circumferential dovetail slot of the insert; an axial passage extending from the plurality of radial passages within a lower portion of the insert; and an exhaust passage extending from the axial passage. | 09-27-2012 |
20130058793 | System for Heat Dissipation from an Internal Actuator in a Rotor Blade - A rotor blade for an aircraft includes an airfoil skin, an active element, and an actuator configured to operate the active element. A heat pipe is configured to promote heat transfer from the actuator to the airfoil skin. The heat pipe has a slope gradient such that a centrifugal force generated during rotation of the rotor blade promotes travel of a condensed working fluid within the heat pipe to move from a condenser end of the heat pipe toward an evaporator end of the heat pipe. | 03-07-2013 |
20140212297 | GAS TURBINE ENGINE SERPENTINE COOLING PASSAGE WITH CHEVRONS - A gas turbine engine component includes a structure having a cooling passage providing upstream and downstream portions separated from one another by an inner wall and fluidly connected by a bend. First and second trip strips are respectively arranged in the upstream and downstream portions. The first trip strips are arranged at a first spacing from one another. The second trip strips are arranged at a second spacing from one another. A turbulence promoter is arranged in the bend and at a third spacing from the first trip strips that is different than the first spacing. The turbulence promoter is arranged at a fourth spacing from the second trip strips that is different than the second spacing. | 07-31-2014 |
20160023275 | ADDITIVE MANUFACTURING BAFFLES, COVERS, AND DIES - A method includes (a) depositing a layer of a powder material on a work stage, the layer having a thickness, (b) solidifying a portion of the layer based upon data that defines an insert with a body that is shaped to fit into a cavity in a gas turbine engine component, and (c) lowering the work stage by the thickness. Steps (a)-(c) can then be repeated until the insert is complete. The insert can then be removed from the work stage. An insert formed by the above process is also disclosed. | 01-28-2016 |