| Entries |
| Document | Title | Date |
|---|
| 20080199303 | Gas Turbine Engine Cooling System and Method | 08-21-2008 |
| 20080219833 | Inducer for a Fan Blade of a Tip Turbine Engine - A fan-turbine rotor assembly for a tip turbine engine includes an inducer with an inducer inlet section and an inducer passage section in communication with a core airflow passage within a fan blade. Each inducer inlet section is canted toward a rotational direction of the fan-turbine rotor assembly such that the inducer inlet section operates as an air scoop during rotation of the fan-turbine rotor assembly. Both axial and centrifugal compression of the airflow occurs within the inducer passage section to effectively pump the airflow through the inducer section and into the core airflow passage. | 09-11-2008 |
| 20080226441 | Method for impingement air cooling for gas turbines - In impingement air cooling of gas turbine components, cooling air velocity packs of a certain amplitude and a given frequency are applied to impingement air openings, with intervallic annular swirl structures being formed which penetrate a cross-flow and hit a component to be cooled with high intensity, thus providing for efficient cooling. In order to obtain annular swirl structures with optimum cooling effect, the Strouhal number, which is determined by a ratio of amplitude, frequency of the velocity packs and size of impingement air cooling openings, ranges between 0.2 and 2.0, and preferably between 0.8 and 1.2. | 09-18-2008 |
| 20080247862 | COMPRESSOR INLET DUCT - An inlet for a compressor having a housing defining a generally C-shaped inlet opening and a rotor cavity configured to contain first and second rotors is provided. The inlet includes an inlet duct defining an inlet opening and a generally C-shaped outlet opening. The inlet duct has an inner wall defining a cavity operable to communicate airflow between the inlet opening and the generally C-shaped outlet opening. The generally C-shaped outlet opening of the inlet duct is substantially similar to the shape of the generally C-shaped inlet opening of the housing. The inner wall includes a floor portion and a roof portion. At least a portion of the floor portion is contoured to impart a velocity component to the airflow complementary to the tangential velocity of each of the first and second rotors during rotation of the first and second rotors. | 10-09-2008 |
| 20080267768 | HIGH-PRESSURE TURBINE OF A TURBOMACHINE - High-pressure turbine of a turbomachine A high-pressure turbine ( | 10-30-2008 |
| 20080273963 | Impingement skin core cooling for gas turbine engine blade - Turbine components, and in particular turbine blades, are provided with impingement cooling channels. Air is delivered along central channels, and the central channels deliver the air through crossover holes to core channels adjacent both a pressure wall and a suction wall. The air passing through the crossover holes impacts against a wall of the core channels. | 11-06-2008 |
| 20080286090 | Turbine Component - A plurality of film cooling holes | 11-20-2008 |
| 20080304956 | COAL NOZZLE TIP SHROUD - An outer shroud for a solid fuel nozzle tip includes: an top shell portion and a bottom shell portion, each portion fabricated from a preform produced from a single sheet of flat stock and each shell portion including a forward area and a backward area and outlet sidewalls, wherein a right outlet sidewall and a left outlet sidewall are each separated from the forward area by a rounded corner; and a left inlet sidewall and a right inlet sidewall coupled to the top shell portion and the bottom shell portion | 12-11-2008 |
| 20080317585 | RECIPROCAL COOLED TURBINE NOZZLE - A turbine nozzle includes first and second vanes joined to outer and inner bands. The vanes include outboard sides defining outboard flow passages containing axial splitlines, and opposite inboard sides defining an inboard flow passage without axial splitline. The two vanes include different cooling circuits for differently cooling the inboard and outboard vane sides. | 12-25-2008 |
| 20090003987 | Airfoil with improved cooling slot arrangement - The present invention relates to airfoils, and in particular turbine blades and vanes, having cooling slots that are angled from a line of reference to effect metering of cooling air through the cooling slots thereof. This metered cooling airflow also creates a more stable film cooling layer about the surface of the airfoil. | 01-01-2009 |
| 20090003988 | Vane assembly with metal trailing edge segment - Embodiments of the invention relate to a vane assembly formed by a forward airfoil segment and an aft airfoil segment. The aft segment is made of metal and can define the trailing edge of the vane assembly. The forward segment can be made of ceramic, CMC or metal. The forward and aft segments cannot be directly joined to each other because of differences in their rates of thermal expansion and contraction. The forward and aft segments can be positioned substantially proximate to each other so as to form a gap therebetween. In one embodiment, the gap can be substantially sealed by providing a coupling insert or leaf springs in the gap. A separate metal aft segment can take advantage of the beneficial thermal properties of the metal to improve cooling efficiency at the trailing edge without limiting the rest of the vane to being made out of metal. | 01-01-2009 |
| 20090003989 | Blade with tangential jet generation on the profile - A blade of a fluid-flow machine has at least one cavity | 01-01-2009 |
| 20090016871 | Systems and Methods Involving Variable Vanes - Systems and methods involving vanes are provided. In this regard, a representative method for modifying the throat area between vanes of a gas turbine engine includes: directing a gas flow path of the gas turbine engine between a first vane and a second vane, wherein each of the first vane and the second vane has an outer surface and an interior; and emitting pressurized air from outlet ports communicating between the outer surface and the interior of the first vane, wherein the emitted pressurized air from the first vane modifies a throat area between the first vane and the second vane. | 01-15-2009 |
| 20090028692 | Systems and Methods for Providing Vane Platform Cooling - Systems and methods for cooling vane platforms are provided. In this regard, a representative method for cooling a vane platform includes: providing a cooling channel on a platform from which a vane airfoil extends, the cooling channel being defined by a cooling surface and a channel cover, the channel wall being spaced from the cooling surface and located such that the cooling surface is positioned between a gas flow path of the vane and the channel cover; and directing a flow of cooling air through the cooling channel such that heat is extracted from the cooling surface of the platform by the flow of cooling air. | 01-29-2009 |
| 20090060712 | Turbine airfoil cooling system with rotor impingement cooling - A turbine airfoil cooling system of a turbine engine having a hollow, disc post body positioned between adjacent roots of turbine airfoils and aligned with the roots to cool inner aspects of the turbine engine. The hollow, disc post body may be configured to pass cooling fluids through impingement orifices in the hollow, disc post body to impinge on inner surfaces of platforms of the turbine airfoils. The cooling fluids may then be directed to the internal cooling systems of the turbine airfoils rather than being discharged as film cooling fluids through the platforms of the turbine airfoils. | 03-05-2009 |
| 20090060713 | DIRECTION-SWITCHABLE PNEUMATIC CYLINDER - A direction-switchable pneumatic cylinder includes: a cylinder body with two intakes, several exhaustion ports and a rotary shaft; and a predetermined number of movable wheels and fixed wheels arranged in the cylinder body and interlaced with each other. Each of the movable wheels and fixed wheels is formed with several vents concentrically arranged into an inner circle and an outer circle. The rotary shaft is fitted through the movable wheels and fixed wheels. The fixed wheels are not rotatable, while the movable wheels are synchronously rotatably with the rotary shaft. The outer circles of vents of the fixed wheels and the movable wheels are aligned with one intake, while the inner circles of vents of the fixed wheels and the movable wheels are aligned with the other intake. When high-pressure gas is guided into the pneumatic cylinder from one intake, the airflow will flow through the outer circles of vents to drive the movable wheels and the rotary shaft in one direction. When high-pressure gas is guided into the pneumatic cylinder from the other intake, the airflow will flow through the inner circles of vents to drive the movable wheels and the rotary shaft in another direction. | 03-05-2009 |
| 20090060714 | Multi-part cast turbine engine component having an internal cooling channel and method of forming a multi-part cast turbine engine component - A multi-part cast component for a turbine engine includes a first component section having a main body portion including at least one cooling flow passage section, and a second component section having a main body including at least one cooling flow passage section. The first and second component sections are joined along a parting line to form a turbine engine component with the at least one cooling flow passage section of the first component section aligning with the at least one cooling flow passage of the second component section to form a cooling flow channel. | 03-05-2009 |
| 20090060715 | Cooled component - A component, such as a turbine blade of a gas turbine engine, has an internal cooling system which includes a passage ( | 03-05-2009 |
| 20090067986 | VORTEX SPOILER FOR DELIVERY OF COOLING AIRFLOW IN A TURBINE ENGINE - A vortex spoiler ( | 03-12-2009 |
| 20090067987 | Airfoil replacement repair - A method of repairing a vane segment for a gas turbine engine includes removing an engine-run cooling baffle from the vane segment, forming a non-engine-run manufacturing detail that includes an inner platform, an outer platform, and an airfoil, attaching the engine-run cooling baffle to the non-engine-run manufacturing detail, and marking the non-engine-run manufacturing detail with a serial number associated with the vane segment from which the engine-run cooling baffle was removed. | 03-12-2009 |
| 20090074562 | NOZZLE GUIDE VANES - A turbine nozzle guide vane | 03-19-2009 |
| 20090074563 | SEAL FOR GAS TURBINE ENGINE COMPONENT - A gas turbine engine component includes a pressurized fluid source, an airfoil, and a seal member for selectively providing sealing at an end of the airfoil. The seal member includes a stowed position for non-sealing and a deployed position for sealing. The seal member is operatively connected with a pressurized fluid source for moving the seal member between the stowed position and the deployed position. | 03-19-2009 |
| 20090081024 | Turbine blade - An aerofoil for a gas turbine engine, the aerofoil comprises a leading edge and a trailing edge, pressure and suction surfaces and defines therebetween an internal passage for the flow of cooling fluid therethrough. A particle deflector means is disposed within the passage to deflect particles within a cooling fluid flow away from a region of the aerofoil susceptible to particle build up and subsequent blockage, such as a cooling passage for a shroud of a blade. | 03-26-2009 |
| 20090081025 | SEGMENTED COOLING AIR CAVITY FOR TURBINE COMPONENT - A component for a gas turbine engine has an airfoil with internal cooling channels for delivering air from a radially outer end of the airfoil toward a radially inner end of the airfoil. The cooling channels are separated from adjacent cooling channels by sets of at least two disconnected wall segments. | 03-26-2009 |
| 20090092478 | SYSTEM AND METHOD FOR IMPROVING FLOW IN PUMPING SYSTEMS - A technique is provided for improving the efficiency of a centrifugal pump. The centrifugal pump comprises diffusers that optimize the area schedule through the diffuser to diffuse the total fluid velocity and recover dynamic head while minimizing flow separation. Each diffuser comprises an improved transition from the diffuser blade into the diffuser discharge duct to remove abrupt changes in area and to reduce fluid separation. The impellers also can be constructed with impeller transitions able to reduce fluid separation and improve the efficiency of the pump. | 04-09-2009 |
| 20090104018 | COOLED BLADE FOR A TURBOMACHINE - The present invention relates to a cooled blade forming an upstream guide vane element for a turbomachine, wherein the airfoil comprises a longitudinal cavity with a first opening at one end and a second opening at the other end, a tubular sleeve being housed in the cavity with a first end in the first opening and a second end in the second opening, first spacers on the side of the first end and second spacers on the side of the second end of the sleeve making a space between the outer face of the sleeve and the wall of the cavity, the blade being arranged so that the sleeve is inserted into the cavity through the first opening. | 04-23-2009 |
| 20090116953 | Turbine airfoil with platform cooling - Convective cooling of gas turbine engine airfoil platforms is enhanced by grooving the interface of the platforms with corresponding platform-to-platform seals, thereby accelerating cooling airflow over the platform surfaces. | 05-07-2009 |
| 20090123267 | INLET FILM COOLING OF TURBINE END WALL OF A GAS TURBINE ENGINE - The disclosure provides a gas turbine engine having a plurality of nozzle vanes extending between an inner shroud wall and an outer shroud wall of the engine. Each of the vanes includes a leading edge and at least one cooling protrusion extending upstream from a center of the leading edge. A cooling system is provided that is operable to inject cooling air upstream from the vanes. The disclosure also provides a method of cooling a gas turbine engine. The method includes the steps of supplying cooling air to an engine nozzle upstream of a gas directing vane, and adjusting the flow of the cooling air with a cooling protrusion extending forward from a leading edge center of the vane. | 05-14-2009 |
| 20090123268 | Z-NOTCH SHAPE FOR A TURBINE BLADE - In one embodiment, a turbine bucket includes: a tip shroud with a front edge and a following edge, the front edge and the following edge including a Z-Notch profile according to the Cartesian coordinate values of X, Y and Z set forth in Table I; wherein the coordinate values are dimensional values representing a distance from an origin of an internal coordinate system for the bucket; and wherein when the X and Y values are connected by smooth continuing arcs, the Z-Notch profile is defined. A turbine is provided. | 05-14-2009 |
| 20090129915 | Turbine Airfoil Cooling System with Recessed Trailing Edge Cooling Slot - A cooling system for a turbine airfoil of a turbine engine having a trailing edge cooling slot positioned within the generally elongated, hollow airfoil and extending from the trailing edge chordwise into the generally elongated, hollow airfoil toward the leading edge such that a secondary trailing edge is offset upstream from the trailing edge. As such, the trailing edge cooling slot reduces stress formation at the trailing edge of the turbine airfoil. | 05-21-2009 |
| 20090129916 | TURBINE APPARATUS - A gas turbine engine comprising a rotor and a stator which define first, second and third cavities; the rotor and stator define a seal therebetween and which is located for sealing between the second and third cavities, the rotor comprises an aperture through which a gas flow passes from the first cavity to the second cavity characterized in that the seal comprises a deflector that extends axially over at least a portion of the aperture to deflect at least a part of the gas towards the rotor. | 05-21-2009 |
| 20090142181 | Gas Turbine Engine Systems Involving Mechanically Alterable Vane Throat Areas - Gas turbine engine systems involving mechanically alterable vane throat areas are provided. In this regard, a representative vane for a gas turbine engine includes: a leading edge; a trailing edge; a suction side surface extending between the leading edge and the trailing edge; a cavity having an aperture located in the suction side surface; and a barrel located within the cavity and being moveable therein such that movement of the barrel alters an extent to which the barrel protrudes through the aperture. | 06-04-2009 |
| 20090148269 | Gas Turbine Engines and Related Systems Involving Air-Cooled Vanes - Gas turbine engines and related systems involving air-cooled vanes are provided. In this regard, a representative vane for a gas turbine engine includes: an airfoil having a leading edge, a pressure surface, a trailing edge and a suction surface; and a cooling air channel; the suction surface being formed by an exterior surface of a first wall portion and an exterior surface of a second wall portion, the first wall portion spanning a length of the suction surface between the second wall portion and the trailing edge; the cooling air channel being defined, at least in part, by an interior surface of the first wall portion, the first wall portion exhibiting a thickness that is thinner than a thickness exhibited by the second wall portion. | 06-11-2009 |
| 20090155050 | DIVERGENT TURBINE NOZZLE - A turbine nozzle includes a row of vanes extending radially in span between inner and outer bands. The vanes include opposite pressure and suction sidewalls and opposite leading and trailing edges. Each vane includes an inner pattern of inner cooling holes and an outer pattern of outer cooling holes distributed along the leading edge. The inner and outer holes diverge toward the corresponding inner and outer bands to preferentially discharge cooling air. | 06-18-2009 |
| 20090155051 | DUPLEX TURBINE SHROUD - A gas turbine engine shroud includes a row of different first and second shroud segments alternating circumferentially therearound. The first segments have a first pattern of first cooling holes extending therethrough. The second segments have a second pattern of second cooling holes extending therethrough. The corresponding patterns have different collective flowrate capabilities. | 06-18-2009 |
| 20090155052 | MOUNTING TUBES FOR PRESSURIZING AN INTERNAL ENCLOSURE IN A TURBOMACHINE - A turbomachine comprising high- and low-pressure compressor shafts guided in bearings isolated from an internal enclosure by a sealing end plate, and radial pressurization tubes connecting the enclosure to an air passage passing through the intermediate casing, the ends of these tubes being engaged in sealed manner in radial ducts of the intermediate casing and in radial chimneys of the sealing end plate, the chimneys being of a length is sufficient to enable the ends of the tubes to be moved in translation therein between a service position and a mounting position for the tubes. | 06-18-2009 |
| 20090162189 | Systems and Methods Involving Variable Throat Area Vanes - Systems and methods involving variable throat area vanes are provided. In this regard, a representative gas turbine engine includes: a vane extending into a gas flow path and having: an interior operative to receive pressurized air; a pressure surface portion; and a first port communicating between the interior and pressure surface portion, the first port being operative to receive the pressurized air from the interior and emit the pressurized air, wherein the emitted pressurized air displaces the gas flow path such that a throat area defined, at least in part, by the vane is modified. | 06-25-2009 |
| 20090162190 | Centrifugal Impeller With Internal Heating - An internal heating arrangement for a centrifugal impeller for a gas turbine engine is provided having at least one heating passage extending through into the rotor for directing air bled from the rotor exit along the backface and forwardly through the impeller. | 06-25-2009 |
| 20090169359 | HEAT EXCHANGER ARRANGEMENT FOR TURBINE ENGINE - A turbine engine cooling arrangement includes a core passage for receiving a core flow for combustion, a first airflow source including a first passage adjacent the core passage for conveying a first airflow, and a second airflow source including a second passage adjacent the first passage for conveying a second airflow. A heat exchanger is thermally connected with the first passage and the second passage for transferring heat between the first airflow and the second airflow. | 07-02-2009 |
| 20090169360 | Turbine Nozzle Segment - A turbine nozzle segment includes a band having a flowpath side and a non-flowpath side and an enclosure disposed on the non-flowpath side of the band. A plenum may be defined between the band and the enclosure and a discourager may extend from the enclosure. | 07-02-2009 |
| 20090169361 | COOLED TURBINE NOZZLE SEGMENT - A turbine nozzle segment may have a band having a flange extending radially from a non-flowpath side and an aft end. A plurality of airfoils may extend radially from a flowpath side of the band and may have trailing edges. A plurality of cooling holes may be disposed in the flange and directed at the aft end between the trailing edges. | 07-02-2009 |
| 20090180860 | PROTECTION DEVICE FOR A TURBINE STATOR - Protection device ( | 07-16-2009 |
| 20090180861 | COOLING ARRANGEMENT FOR TURBINE COMPONENTS - A turbine component includes an aft cooling circuit that extends between a turbine midsection and a turbine trailing end. The aft cooling circuit includes a trailing end section proximate the trailing end, a first interior section proximate the turbine midsection, and a first intermediate section fluidly connected between the trailing end section and the first interior section. A forward cooling circuit of the turbine component extends between the turbine midsection and a turbine leading end. The forward cooling circuit includes a leading end section proximate the leading end, a second interior section proximate the turbine midsection, and a plurality of second intermediate sections fluidly connected between the leading end section and the second interior section. The leading end section, the second intermediate section, the first intermediate section, and the trailing end section each include a plurality of coolant discharge openings for facilitating cooling of the turbine component. | 07-16-2009 |
| 20090185896 | GAS TURBINE AND GAS TURBINE COOLING METHOD - A gas turbine includes a nozzle vane and a sealing unit engaged with the nozzle vane inside a turbine supplied with combustion gases produced by mixing and burning air for combustion and fuel. The nozzle vane and the sealing unit are disposed in a channel of the downward flowing combustion gases on the outlet side of a gas path. A plurality of engagement portions between the sealing unit and the nozzle vane are provided successively from the upstream side toward the downstream side in a direction of flow of the combustion gases, and a downstream one of the plurality of engagement portions has a contact interface formed in a direction across a turbine rotary shaft. A reduction in the thermal efficiency of the gas turbine can be suppressed. | 07-23-2009 |
| 20090196736 | APPARATUS AND RELATED METHODS FOR TURBINE COOLING - An apparatus and a method for cooling and/or sealing a gas turbine by selectively boosting the pressure of air extracted at a lower extraction stage is provided. The pressure of the extracted air is boosted by an external compressor before it becomes available for cooling and/or sealing the turbine components. A bypass line includes a higher extraction stage providing air for cooling the turbine. | 08-06-2009 |
| 20090196737 | Cooling airflow modulation - A gas turbine engine airfoil ( | 08-06-2009 |
| 20090196738 | GAS TURBINE AND GAS TURBINE COOLING METHOD - A gas turbine includes a nozzle vane and a sealing unit engaged with the nozzle vane inside a turbine supplied with combustion gases produced by mixing and burning air for combustion and fuel. The nozzle vane and the sealing unit are disposed in a channel of the downward flowing combustion gases on the outlet side of a gas path. A plurality of engagement portions between the sealing unit and the nozzle vane are provided successively from the upstream side toward the downstream side in a direction of flow of the combustion gases, and a downstream one of the plurality of engagement portions has a contact interface formed in a direction across a turbine rotary shaft. A reduction in the thermal efficiency of the gas turbine can be suppressed. | 08-06-2009 |
| 20090202338 | BLADE FOR A FLOW MACHINE - In a blade intended to be exposed to a gas flow at high speed during operation of a flow machine comprising the blade, the blade includes a front end designed to face towards the incoming gas flow and a rear end. The front end is provided with a concave area that is such that during operation a stagnation point for the incoming gas flow arises at a distance in front of an outer blade surface that defines the concave area and such that the outer blade surface is thereby at least partially protected from the incoming gas flow. | 08-13-2009 |
| 20090202339 | PLATFORM COOLING STRUCTURE FOR GAS TURBINE MOVING BLADE - A platform cooling structure for a gas turbine moving blade is provided which is capable of improving cooling performance of a platform and of improving reliability of a moving blade in such a manner that a portion in the vicinity of a side edge of the platform which is away from moving blade cooling passageways and is easily influenced by thermal stress caused by high-temperature combustion gas, that is, an upper surface of the side edge is effectively cooled by guiding high-pressure cooling air, flowing to the moving blade cooling passageways, to a discharge opening formed in a surface of the platform in the vicinity of the side edge of the platform without particularly attaching an additional member such as a cover plate to the platform. A moving blade cooling passageway | 08-13-2009 |
| 20090208323 | METHODS AND APPARATUS FOR COOLING ROTARY COMPONENTS WITHIN A STEAM TURBINE - A method for cooling a rotating component within a steam turbine is provided. The method includes channeling a cooling fluid through an outer plenum defined in a stationary component of the steam turbine and channeling the cooling fluid from the outer plenum through a passageway defined in an airfoil of the stationary component. The cooling fluid is discharged from the airfoil passageway through an inner plenum of the stationary component to facilitate cooling an adjacent rotating component. | 08-20-2009 |
| 20090208324 | Casing structure for stabilizing flow in a fluid-flow machine | 08-20-2009 |
| 20090214328 | BLADES FOR GAS TURBINE ENGINES - A blade for a gas turbine engine comprises an aerofoil having a root portion, a tip portion located radially outwardly of the root portion, and leading and trailing edges extending between the root portion and the tip portion. A shroud extends transversely from the tip portion of the aerofoil and the aerofoil defines interior cooling passages which extend between the root portion and the tip portion. The aerofoil includes a wall member adjacent the trailing edge and a support structure extending from the wall member to the shroud to support the shroud. The support structure permits a flow of cooling air from a cooling passage to the trailing edge at a region proximate the tip portion of the aerofoil. Optionally, the aerofoil also includes a flow disrupting arrangement. | 08-27-2009 |
| 20090220331 | TURBINE NOZZLE WITH INTEGRAL IMPINGEMENT BLANKET - A turbine nozzle segment includes: (a) an arcuate outer band segment; (b) a hollow, airfoil-shaped turbine vane extending radially inward from the outer band segment; (c) a manifold cover secured to the outer band such that the manifold cover and the outer band segment cooperatively define an impingement cavity; and (d) an impingement blanket disposed in the impingement cavity, the impingement blanket having at least one impingement hole formed therethrough which is arranged to direct cooling air at the outer band segment. A method is provided for impingement cooling the outer band segment. | 09-03-2009 |
| 20090220332 | AXIAL FLOW FLUID APPARATUS AND BLADE - An axial flow fluid apparatus axially provided with a plurality of blade rows having a plurality of blades arranged around a shaft is provided. A fluid passage for jetting a fluid to a downstream velocity defect region resulting from a blade is formed in at least one of blades constituting a blade row installed on the upstream side of the plurality of blade rows so as to lead from a positive pressure surface to a negative pressure surface or a trailing edge. | 09-03-2009 |
| 20090245999 | HYBRID IMPINGEMENT COOLED AIRFOIL - A turbine nozzle for a gas turbine engine includes: (a) spaced-apart arcuate inner and outer bands; (b) a hollow, airfoil-shaped turbine vane extending between the inner and outer bands, the interior of the vane defining at least a forward cavity and a mid-cavity positioned aft of the forward cavity; (c) a hollow impingement insert received inside the mid-cavity, the impingement insert having walls which are pierced with at least one impingement cooling hole; (d) a passage in the turbine vane at a radially outer end of the forward cavity adapted to be coupled to a source of cooling air; and (e) a passage in the inner band in fluid communication with a radially inner end of the forward cavity and a radially inner end of the impingement insert. | 10-01-2009 |
| 20090252596 | Fluid flow machine with fluid injector assembly - A fluid flow machine has a flow duct | 10-08-2009 |
| 20090269184 | Gas Turbine Engine Systems Involving Turbine Blade Platforms with Cooling Holes - Gas turbine engine systems involving turbine blade platforms with mateface cooling holes are provided. In this regard, a representative turbine blade for a gas turbine engine includes: an airfoil having a leading edge, a trailing edge, a pressure side and a suction side; and a blade platform on which the airfoil is disposed, the blade platform having a pressure side mateface located adjacent to the pressure side of the airfoil and a suction side mateface located adjacent to the suction side of the airfoil, the blade platform having a cooling hole operative to direct a flow of cooling air toward an adjacent blade platform. | 10-29-2009 |
| 20090274549 | WALL COOLING ARRANGEMENT - A wall cooling arrangement comprising on one side of a wall a multiplicity of cooling fluid inlet apertures and on the opposite of the wall a multiplicity of cooling fluid exit apertures, and in the body of the wall linking said inlet and exit apertures a network of multiply branched cooling passages. Flow of cooling fluid through a network is controlled by a throat positioned either at or close to the inlet to the passage network or at a location part way through the network, in which case there may be a plurality of inlet apertures feeding through a single throat to a plurality of outlet apertures. | 11-05-2009 |
| 20090274550 | GAS TURBINE COMPONENTS AND METHOD FOR MACHINING GAS TURBINE COMPONENTS - The present technology relates to the problem that during diverse machining steps of application to the production or reconditioning of internally cooled gas turbine blades, an undesired effect may be had on sections of the gas turbine blades and proposes, as an improvement, to inject the cavity of the gas turbine blades before the machining steps with a plastic material which can be removed without trace, such as polystyrene, which can be subsequently removed again, in particular by heat. | 11-05-2009 |
| 20090285670 | APPARATUS AND METHOD FOR DOUBLE FLOW TURBINE FIRST STAGE COOLING - A method of cooling a double flow steam turbine includes supplying steam flow to each nozzle of the sections of the turbine; reversing a portion of each steam flow to provide a reverse steam flow from an aft side to a forward side of each section. Each reverse steam flow is directed to an annular space between a rotor and a tub. The method further includes removing the reverse steam flows through a pipe, the pipe having a first end at the annular space at a first pressure and a second end at a second pressure that is lower than the first pressure. A double flow steam turbine, includes a pair of nozzles, each nozzle being provided at a section of the turbine; a rotor supporting buckets of the sections; a tub supporting the pair of nozzles; and a pipe extending from an annular space between the tub and the rotor. The pipe has a first end at the annular space and second end. A pressure at the first end of the pipe is greater than a pressure at the second end. | 11-19-2009 |
| 20090297335 | ASYMMETRIC FLOW EXTRACTION SYSTEM - A system for asymmetric flow extraction is described and claimed, the system comprising a flow path, a bleed slot in the flow path, a bleed cavity for receiving at least a portion of the fluid extracted from the flow path and a bleed passage in flow communication with the bleed slot and the bleed cavity wherein the bleed passage has at least one deflector having a shape such that the width of the bleed passage cross section varies in a direction normal to the direction of fluid flow in the bleed passage. In another embodiment, the deflector has an aerodynamic surface having a shape such that the flow passage between the aerodynamic surface and a surface located away from it has a cross sectional shape that is non-axisymmetric. In another embodiment, the bleed passage comprises an assembly of a plurality deflectors, arranged circumferentially. | 12-03-2009 |
| 20090311090 | WINDWARD COOLED TURBINE NOZZLE - A turbine nozzle includes a hollow vane mounted between inner and outer bands. The inner band includes a mounting flange between forward and aft lips. An aft pocket is found in the inner band between the flange and aft lip. And, an impingement bore extends through the flange into the windward half of the pocket and is directed aft toward the opposite leeward half of the pocket for co-rotation with purge flow during operation. | 12-17-2009 |
| 20090311091 | IMPELLER AND CENTRIFUGAL PUMP INCLUDING THE SAME - An example impeller includes: an impeller body in which an internal channel is formed, the internal channel extending inside the impeller body in a direction of a rotation axis spirally about the rotation axis to connect an inlet and an outlet; and at least one centrifugal vane provided in the impeller body. The internal channel including the inlet and the outlet has a predetermined passage diameter. An external channel is formed so as to continue to the outlet and go around the circumferential surface of the impeller body, the external channel being defined by the centrifugal vane and being recessed inward in the radial direction from the circumferential surface of the impeller body. At least a part in a flow direction of the external channel has a channel width in the direction of the rotation axis smaller than the width of the outlet. | 12-17-2009 |
| 20090317234 | CROSSFLOW TURBINE AIRFOIL - A turbine airfoil includes pressure and suction sidewalls extending axially in chord between opposite leading and trailing edges. The sidewalls are spaced transversely apart to define flow channels extending longitudinally and separated chordally by partitions bridging the sidewalls. A perforate partition includes a row of crossover holes extending obliquely therethrough. | 12-24-2009 |
| 20090324385 | Airfoil for a gas turbine - An airfoil is provided for a gas turbine comprising an outer structure comprising a first wall, an inner structure comprising a second wall spaced relative to the first wall such that a cooling gap is defined between at least portions of the first and second walls, and seal structure provided within the cooling gap between the first and second walls for separating the cooling gap into first and second cooling fluid impingement gaps. An inner surface of the second wall may define an inner cavity. The inner structure may further comprise a separating member for separating the inner cavity of the inner structure into a cooling fluid supply cavity and a cooling fluid collector cavity. The second wall may comprise at least one first impingement passage, at least one second impingement passage, and at least one bleed passage. | 12-31-2009 |
| 20090324386 | GAS TURBINE - A gas turbine includes a shaft directional passage provided to a rotating member that rotates along a central axis of a rotor, which is a rotating axis of the rotating member, about the central axis, or a rotating axis of the rotating member, and in which cooling air flows along a direction of the rotating axis of the rotor, a plurality of radial directional passages provided in a circumferential direction of the rotating member, and compressing the cooling air by being provided outwardly from the center of the rotor, in which one end of each of the radial directional passages is communicated with the shaft directional passage and the other end is communicated with exterior of the rotating member. | 12-31-2009 |
| 20090324387 | Aft frame with oval-shaped cooling slots and related method - An aft frame adapted to interface between a combustor transition piece and a first stage turbine nozzle includes: a closed-periphery frame comprised of a top, bottom and pair of side walls. A plurality of cooling holes or apertures having elliptical or oval cross-sectional shapes are provided in one or more of the top, bottom and pair of side walls, extending axially through the closed-periphery frame. The cooling holes have major and minor axes arranged such that the major axes are substantially parallel with the top and bottom walls. | 12-31-2009 |
| 20100008758 | LEADING EDGE COOLING WITH MICROCIRCUIT ANTI-CORIOLIS DEVICE - A turbine engine component, such as a high pressure turbine blade, has an airfoil portion having a pressure side, a suction side, and a leading edge. A cooling system is provided within the leading edge. The cooling system includes at least one peripheral leading edge cooling channel for creating anti-Coriolis forces in the leading edge of the airfoil portion. | 01-14-2010 |
| 20100008759 | METHODS AND APPARATUSES FOR PROVIDING FILM COOLING TO TURBINE COMPONENTS - Methods and apparatuses for film cooling of one or more turbine components are provided. A cooling gas flow passage provides a cooling gas to an turbine component with the hot gas path of a turbine. The cooling gas flow passage includes at least one feed aperture operable to receive cooling gas from at least one cooling gas compartment associated with a turbine component. The cooling gas flow passage also includes at least one slot with a converging portion having a first opening and a second opening, where the first opening receives the cooling gas from the feed aperture, and the second opening provides the cooling gas to at least a portion of an outer surface of the turbine component. | 01-14-2010 |
| 20100008760 | GAS TURBINE ENGINE ASSEMBLIES WITH RECIRCULATED HOT GAS INGESTION - A gas turbine engine assembly includes a housing including an annular duct wall that at least partially defines a mainstream hot gas flow path configured to receive mainstream hot gas flow. The assembly further includes a stator assembly including a stator vane that extends into the mainstream hot gas flow path and a turbine rotor assembly downstream of the stator assembly that includes a turbine disk and a turbine blade extending from the turbine disk into the mainstream hot gas flow path. The stator assembly and turbine assembly define a turbine disk cavity, and the turbine disk cavity includes a recirculation cavity configured to recirculate gas ingested from the mainstream hot gas flow path back into the mainstream hot gas flow path. | 01-14-2010 |
| 20100008761 | COOLABLE AIRFOIL TRAILING EDGE PASSAGE - An airfoil suitable for use in a gas turbine engine having at least one feed passage at least in part defined along a feed axis which is at least perpendicular to a cavity axis to reduce dirt ingestion into a trailing edge passage. | 01-14-2010 |
| 20100014958 | TURBINE ENGINE ROTOR DISC WITH COOLING PASSAGE - Disclosed is a gas turbine engine rotor disc with a plurality of cooling passages having an essentially radial orientation relative to an axis of rotation of the rotor disc, each cooling passage having an inlet and an outlet and being included relative to a rotor disc surface and a cut-out arranged at the passage at an outlet end of the passage. Each cooling passage terminating in a slot is arranged in the periphery of the rotor disc. Each slot is sized and configured to receive a glade root. | 01-21-2010 |
| 20100028131 | Component for a Turbine Engine - A component for use in a turbine engine including a first member and a second member associated with the first member. The second member includes a plurality of connecting elements extending therefrom. The connecting elements include securing portions at ends thereof that are received in corresponding cavities formed in the first member to attach the second member to the first member. The connecting elements are constructed to space apart a first surface of the second member from a first surface of the first member such that at least one cooling passage is formed between adjacent connecting elements and the first surface of the second member and the first surface of the first member. | 02-04-2010 |
| 20100034638 | Impingement cooling arrangement - An impingement cooling arrangement comprises a projection extending partially across a coolant passage upstream of a jet aperture. An end surface of the projection increases the available surface area for heat exchange with a cross flow whilst a coolant air flow jetted from the jet aperture can transgress a proportion of the air flow passing between the end surface and a junction surface incorporating the jet aperture. A spacing gap B between the end surface and the junction surface avoids localised distortions to the cross flow whilst the projection provides that the coolant air flow projected from the jet aperture mostly passes through a lower turbulence wake downstream of the projection for greater impingement upon a target surface for heat transfer and cooling efficiency. Typically the impingement cooling arrangement is incorporated within turbine blades or vanes of a jet engine. | 02-11-2010 |
| 20100034639 | Air directing assembly and method of assembling the same - A gas turbine engine assembly is provided. The gas turbine engine assembly includes a high-pressure compressor including a first rotor and a second compressor rotor disposed downstream from the first rotor, a high-pressure turbine coupled downstream to the compressor by a first shaft, and an air directing assembly coupled between the first and second compressor rotors for selectively channeling airflow discharged from the first compressor rotor through the first shaft. A method of assembling a gas turbine engine is also provided. | 02-11-2010 |
| 20100034640 | NESTED CORE GAS TURBINE ENGINE - A fan for creating lift or thrust having a fan hub and fan blades depending from the fan hub. The fan blades have slots formed therein with openings facing substantially aft, relative to a rotation direction of the fan blades, wherein air blowing from the fan hub into the fan blades and out of the openings of the slots contribute to the aerodynamic performance of the fan blades to enhance the aerodynamic performance of the fan blades. | 02-11-2010 |
| 20100040455 | METHOD AND DEVICE FOR JOINING METAL ELEMENTS - A method for connecting metallic components, in particular components of a gas turbine, including: corresponding connecting surfaces of the components being connected by means of inductive HF pressure welding, and that during or after a sufficiently great heating of the connecting surfaces, the first component is moved by a definite path toward the second component, and is pressed against it and held there. | 02-18-2010 |
| 20100047056 | Duplex Turbine Nozzle - A duplex turbine nozzle includes a row of different first and second vanes alternating circumferentially between radially outer and inner bands in vane doublets having axial splitlines therebetween. The vanes have opposite pressure and suction sides spaced apart in each doublet to define an inboard flow passage therebetween, and corresponding outboard flow passages between doublets. The vanes have different patterns of film cooling holes with larger cooling flow density along the outboard passages than along the inboard passages. | 02-25-2010 |
| 20100047057 | Aerofoil - An aerofoil comprising a pressure-side wall, a suction-side wall and an intermediate wall extending from a free end of the pressure-side wall at an acute angle relative thereto towards the suction-side wall. A cooling fluid passageway extends through a region where the intermediate wall meets the pressure-side wall at an apex. The fluid passageway has an opening, at least in part, in the face of the pressure-side wall. | 02-25-2010 |
| 20100068032 | Turbine Airfoil Cooling System with Diffusion Film Cooling Hole - 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 a first section extending from an inner surface of the outer wall into the outer wall, a second section extending the first section toward an outer wall, and a third section extending from the second section and terminating at an outer surface of the outer wall. The diffusion film cooling hole may provide a metering capability together with diffusion sections that provide a larger film cooling hole breakout and footprint, which create better film coverage and yield better cooling of the turbine airfoil. The diffusion film cooling hole may provide a smooth transition, which allows the film cooling flow to diffuse better in the second and third sections of the diffusion film cooling hole. | 03-18-2010 |
| 20100068033 | Turbine Airfoil Cooling System with Curved Diffusion Film Cooling Hole - 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 a first sidewall having a first radius of curvature about an axis generally orthogonal to a centerline of cooling fluid flow through the diffusion film cooling hole and a second sidewall having a second radius of curvature about an axis generally orthogonal to the centerline of cooling fluid flow through the at least one diffusion film cooling hole. The radii of curvature of the first and second sidewalls are different such that the diffusion film cooling hole includes an ever increasing cross-sectional area moving from an inlet to an outlet, thereby diffusing and reducing the velocity of cooling fluids flowing there through. | 03-18-2010 |
| 20100068034 | CMC Vane Assembly Apparatus and Method - A metal vane core or strut ( | 03-18-2010 |
| 20100074726 | GAS TURBINE AIRFOIL - A gas turbine airfoil ( | 03-25-2010 |
| 20100080687 | Multiple Piece Turbine Engine Airfoil with a Structural Spar - A multiple piece turbine airfoil having an outer shell with an airfoil tip that is attached to a root with an internal structural spar is disclosed. The root may be formed from first and second sections that include an internal cavity configured to receive and secure the one or more components forming the generally elongated airfoil. The internal structural spar may be attached to an airfoil tip and place the generally elongated airfoil in compression. The configuration enables each component to be formed from different materials to reduce the cost of the materials and to optimize the choice of material for each component. | 04-01-2010 |
| 20100080688 | HOT GAS COMPONENT OF A TURBOMACHINE INCLUDING AN EMBEDDED CHANNEL - A component, especially a hot gas component of a turbomachine, has at least one passage ( | 04-01-2010 |
| 20100086394 | HYDRAULIC MACHINE - A hydraulic machine having a hydro turbine runner which has a crown at a center and a band along an outer periphery, and is formed around the axis of rotation, long blades which are arranged along the circumferential direction of the axis of rotation, and whose center-side ends are supported by the crown, and periphery-side ends are supported by the band, and short blades which are arranged between the long blades, and whose center-side ends are supported by the crown, periphery-side ends are supported by the band, and rear edges are curved in a rotation direction of the hydro turbine runner in turbine operation, on a plane of projection perpendicular to the axis of rotation. | 04-08-2010 |
| 20100098526 | AIRFOIL WITH COOLING PASSAGE PROVIDING VARIABLE HEAT TRANSFER RATE - A turbine engine airfoil includes an airfoil structure having a side with an exterior surface. The structure includes a cooling passage extending a length within the structure and providing a convection surface facing the side. The convection surface is twisted along the length, which varies a heat transfer rate between the exterior surface and the convection surface along the length. In one example, the cooling passage is provided by a refractory metal core that is used during the airfoil casting process. The core includes multiple legs joined by a connecting portion. At least one of the legs is twisted along its length. The legs are deformed toward one another opposite the connecting portion to provide a desired core shape that corresponds to the shape of the cooling passage. Accordingly, the cooling passage provides desired cooling of the airfoil. | 04-22-2010 |
| 20100098527 | FLUID FLOW MACHINE WITH PERIPHERAL ENERGIZATION NEAR THE SUCTION SIDE - A fluid flow machine has a main flow path (“MFP”) | 04-22-2010 |
| 20100104419 | TURBINE AIRFOIL WITH NEAR WALL INFLOW CHAMBERS - 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 for receiving cooling fluids from a cooling fluid supply source, passing those fluids through the chambers in the outer wall, and exhausting those fluids into central cooling fluids collection chambers. The outer wall may include a plurality of outer wall cooling chambers that may be configured to pass cooling fluids in a counter flow direction. The outer wall cooling chambers may include a plurality of ribs including a plurality of impingement orifices for increasing the cooling efficiency of the cooling system. | 04-29-2010 |
| 20100111670 | SHROUD HANGER WITH DIFFUSED COOLING PASSAGE - A shroud hanger for a gas turbine engine has an arcuate body with opposed inner and outer faces and opposed forward and aft ends, the channel having at least one cooling passage therein which includes: (a) a generally axially-aligned channel extending through the body, the channel having one end open to an exterior of the body; and (b) a generally radially-aligned diffuser extending through the inner face and intersecting the channel. | 05-06-2010 |
| 20100111671 | METHODS AND APPARATUS INVOLVING SHROUD COOLING - A turbine cooling component comprising a circumferential leading edge, a circumferential trailing edge, a pair of spaced and opposed side panels connected to the leading and trailing edges, an arcuate base connected to the trailing and leading edges having a fore portion, a midsection portion, an aft portion, opposed side portions, an outer surface partially defining a cavity operative to receive pressurized air, and an arcuate inner surface in contact with a gas flow path of a turbine engine, a first side cooling air passage in the base extending along the first side portion from the fore portion to the aft portion, and a fore cooling air passage in the fore portion of the base communicative with the side cooling air passage and the cavity, operative to receive the pressurized air from the cavity. | 05-06-2010 |
| 20100111672 | Hesting Power Turbine Device - A power turbine employs a plurality of turbine blades which allows the turbine to regulate pressure and ventilate for a greater ability to generate power without back pressure of the power source. It also relieves much of the stress on the uni-body construction. The vent ports #30 work at various times to transfer pressure naturally from one compartment to another, greatly increasing the balance of pressure throughout the process within the turbine. | 05-06-2010 |
| 20100119357 | Gas Turbine - A gas turbine including a rotor shaft, a plurality of rotor blades that extend generally radially outwardly from the rotor shaft, each rotor blade including a shroud radially outward of an aerofoil, and a plurality of guide vanes located adjacent to the plurality of rotor blades, the plurality of guide vanes also extending generally radially outwardly is provided. The guide vanes operate to direct gas flowing through the turbine onto the rotor blades. A guide vane accommodates a flow of cooling fluid to an aperture in the guide vane that is located in a region that is adjacent both the radially outer end of the guide vane and the trailing edge of the guide vane. The flow of cooling fluid emanating from the aperture travels to impinge upon the shroud thereby cooling the shroud. The aperture is located in the high or low pressure side of the guide vane. | 05-13-2010 |
| 20100124483 | Apparatus and method for cooling a turbine airfoil arrangement in a gas turbine engine - A turbine airfoil arrangement for a gas turbine engine includes an airfoil having an inlet and an exit, the inlet configured to receive a cooling gas flow operable to cool at least part of an other airfoil; and a passage disposed in the airfoil and fluidly coupled to the inlet and the exit, the exit being configured to pass at least some of the cooling gas flow to the other airfoil. | 05-20-2010 |
| 20100124484 | Aerofoil and method for making an aerofoil - Within aerofoils, and in particular nozzle guide vane aerofoils in gas turbine engines problems can occur with regard to coolant flows from respective inlets at opposite ends of a cavity within the aerofoil. The cavity generally defines a hollow core and unless care is taken coolant flow can pass directly across the internal cavity. Previously baffle plates were inserted within the cavity to prevent such direct jetting across the cavity. Such baffle plates are subject to additional costs as well as potential unreliability problems. Baffles formed integrally with a wall within the aerofoil allow more reliability with regard to positioning as well as consistency of performance. The baffles can be perpendicular, upward or downwardly orientated or have a compound angle. | 05-20-2010 |
| 20100124485 | Aerofoil cooling arrangement - Within aerofoils ( | 05-20-2010 |
| 20100129194 | 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. | 05-27-2010 |
| 20100129195 | 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 plurality of metallic casting cores. Each metallic casting core has opposite first and second faces and a respective portion along the trailing edge of the airfoil. At least two of the metallic cores have sections offset between the pressure side and the suction side. | 05-27-2010 |
| 20100129196 | COOLED GAS TURBINE VANE ASSEMBLY - A gas turbine vane to improve vane performance by addressing known failure mechanisms. A cooling circuit to the trailing edge of a vane airfoil is fed from the outer diameter platform, which prevents failure due to an oxidized and eroded airfoil trailing edge. The gas turbine includes an outer diameter platform, a hollow airfoil and an inner diameter platform with a plurality of cooling tubes extending radially through the airfoil. The cooling tubes are open at the outer diameter end and closed with covers at the inner diameter end. The inner diameter platform is also cooled and includes a meterplate for a portion of the cooling passageway and includes an undercut to improve thermal deflections of the inner diameter platform. | 05-27-2010 |
| 20100129197 | METHOD AND SYSTEM FOR COOLING ENGINE COMPONENTS - A method and system for a rotatable member of a turbine engine are provided. The rotatable member includes a substantially cylindrical shaft rotatable about a longitudinal axis, and a hub coupled to the cylindrical shaft through a conical shaft portion wherein the conical shaft portion includes a plurality of circumferentially-spaced air passages and wherein at least one of the plurality of air passages includes a non-circular cross section. | 05-27-2010 |
| 20100129198 | HYDRAULIC MACHINE INCLUDING MEANS FOR INJECTING A FLOW DRAWN FROM A MAIN FLOW - The invention relates to a hydraulic machine through which a main flow (E) of water passes, including at least one turbine blade profile ( | 05-27-2010 |
| 20100129199 | Platform Cooling of Turbine Vane - A turbine vane is provided which includes a radial outer platform, a radial inner platform and an airfoil extending between the outer platform and the inner platform. Each platform has a gas washed surface facing towards the respective other platform, a non gas washed surface facing away from the respective other platform and a peripheral surface extending from the gas washed surface to the non gas washed surface. The peripheral surface includes an upstream section that is designed to be directed towards the gas flow washing the gas washed surface. Cooling fluid channels each include an opening in the peripheral surface or in the gas washed surface and are located in at least a section of the outer platform and/or in at least a section of the inner platform. The respective section directly adjoins the upstream section of the peripheral surface of the respective platform. | 05-27-2010 |
| 20100135772 | TURBINE AIRFOIL COOLING SYSTEM WITH PLATFORM COOLING CHANNELS WITH DIFFUSION SLOTS - A cooling system for a turbine airfoil of a turbine engine having suction side platform cooling channels and pressure side platform cooling channels for cooling hot spots in a platform attached to a turbine blade. The cooling system may include one or more pressure side platform cooling chambers having a diffusion slot for cooling downstream platforms on the suction side of the turbine blade. The diffusion slots reduce the velocity of the cooling fluids released from the platform to increase the capacity of the film cooling of downstream platforms. | 06-03-2010 |
| 20100158669 | MICROCIRCUITS FOR SMALL ENGINES - A turbine engine component for use in a small engine application has an airfoil portion having a root portion, a tip portion, a suction side wall, and a pressure side wall. The suction side wall and the pressure side wall have the same thickness. Still further, the turbine engine component has a platform with an internal cooling circuit. | 06-24-2010 |
| 20100178157 | HEAT EXCHANGE ELEMENT, MANUFACTURING METHOD THEREOF, AND HEAT EXCHANGE VENTILATOR - A heat exchange element according to the present invention has a stacked-layer structure in which sheet-like partition members and spacing members are stacked alternately, while the spacing members are joined with the partition members so as to form air flow passages together with the partition members. A plurality of adhesive layers included in the stacked-layer structure i.e., the plurality of adhesive layers that cause the partition members to be each joined with a corresponding one of the spacing members include one or more colored adhesive layers. It is therefore easy to manufacture a heat exchange element having a desired color arrangement at a low cost, regardless of the type of the heat exchange element. | 07-15-2010 |
| 20100183427 | TURBINE BLADE WITH MICRO CHANNEL COOLING SYSTEM - A cooling system for a turbine airfoil of a turbine engine has a multi-pass serpentine flow circuit providing a flow path from a forward cooling flow entry at the root and exhausting towards the trailing edge through a series of chord wise micro channels extending from the rearward pass of the multi-pass serpentine circuit to pressure side bleed slots, each having a forward pressure side lip and opening onto the pressure side adjacent the trailing edge. The micro channels can be formed by a series of spaced fins stacked span wise and extending between the outer wall on the pressure side and the outer wall on the suction side and extending chord wise from the rearward pass to the trailing edge. At least two trip strips can extend from sides of the fins into the micro channels and be staggered relative to trip strips extending into the micro channel from an adjacent fin, whereby turbulent flow levels in the micro channels are increased. | 07-22-2010 |
| 20100183428 | MODULAR SERPENTINE COOLING SYSTEMS FOR TURBINE ENGINE COMPONENTS - A cooling system for use in a turbine engine component exposed to high temperatures during engine operation. The system includes a serpentine flow passage and an exhaust region. The serpentine flow passage includes a coolant supply inlet. The passage can be configured so that neighboring portions of the passage have coolant flowing in the same direction or, alternatively, in opposite directions. A number of flow disrupting structures, such as microfins and trip strips, can be located along the flow passage. The exhaust region can discharge coolant from the system at reduced exit momentum. The exiting flow can provide film cooling to the component. The cooling system can be provided in a small modular form, which can increase cooling design flexibility and can allow cooling designs tailored to the unique cooling requirements of the individual component. As a result, the modules can result in high levels of cooling effectiveness. | 07-22-2010 |
| 20100183429 | TURBINE BLADE WITH MULTIPLE TRAILING EDGE COOLING SLOTS - A cooling system for a turbine airfoil of a turbine engine has a multiple suction side cooling slots extending from a front edge on the suction side to the center of the trailing edge or even to the pressure side of the center line and a pressure side cooling slot curving to a pressure side outlet forward of the trailing edge and having a front pressure side lip that is aligned with or forward of the front edge of the suction side cooling slots. The suction side cooling slots receive cooling flow from the pressure side cooling slots through a boundary layer bleed valve, which is also aligned with or rearward of the pressure side lip. The cooling system may also combine double impingement cooling with these features. The cooling system minimizes shear mixing, reduces hot spots and can reduce the trailing edge thickness, resulting in more efficient stage performance and extended operational life. | 07-22-2010 |
| 20100209229 | Airfoil inserts, flow-directing elements and assemblies thereof - Disclosed are examples of flow-directing elements, airfoil inserts, and assemblies thereof. A flow-directing element has an inner buttress with an airfoil extending outwardly therefrom. The airfoil includes a cavity that extends within the airfoil to an exit port disposed in the inner buttress. A shelf disposed about the buttress defines the exit port, and the shelf includes a discourager extending into the cavity. An airfoil insert has a tubular body, with an outlet at one end. A plate affixed to the body at the outlet partially blocks the outlet, and includes a tab extending away from the body and defining a portion of an outlet periphery. Upon assembly of the flow directing element and the insert, the tab interacts with the discourager to direct a coolant to the exit port while restricting leakage of the coolant back into the cavity, between the airfoil insert and the flow-directing element. | 08-19-2010 |
| 20100221098 | Peripheral Microcircuit Serpentine Cooling for Turbine Airfoils - A turbine component has an airfoil portion with at least one central core element, a pressure side wall, and a suction side wall. The airfoil portion also has a serpentine cooling passageway in at least one of the walls. In a preferred embodiment, the airfoil portion has a serpentine cooling passageway in both of the pressure and suction side walls. A refractory metal core for forming the serpentine cooling passageway(s) is also described. | 09-02-2010 |
| 20100226755 | Turbine Vane for a Gas Turbine Engine Having Serpentine Cooling Channels Within the Outer Wall - A turbine vane for a gas turbine engine having an outer wall containing a plurality of serpentine cooling channels. The serpentine cooling channels may be configured to receive cooling fluids from internal cooling fluids supply channels. The serpentine cooling channels may be positioned in the pressure side and suction side outer walls and configured such that a first pass is positioned radially outward from an internal chamber a greater distance than a second pass. As such, cooling fluids are first passed proximate to an outer surface where the fluids are heated and then passed proximate to an inner surface, thereby establishing a smaller thermal gradient than typically found in conventional turbine blade outer walls. | 09-09-2010 |
| 20100232930 | Gas turbine engine - A gas turbine engine for highly efficient fuel consumption includes as elements a rotor, a stator and a spindle. The rotor includes a system of passageways, a combustion chamber, and a plurality of exhaust ports. The system of passageways comprises a plurality of axial passageways and radial passageways configured to receive a precombustion air-fuel mixture and transport the precombustion air-fuel mixture to the combustion chamber. The exhaust ports provide fluid communication of postcombustion gas between the combustion chamber and the exterior of the rotor. | 09-16-2010 |
| 20100232931 | HEAT DISSIPATION FAN - A heat dissipation fan includes a base with a plurality of holes defined therein, a stator mounted on the base and being placed around the holes of the base, and an impeller rotatably attached to the base. The impeller includes a hub and a plurality of blades, the hub includes a top wall with a plurality of holes defined therein and an annular wall depending from the top wall. The blades are arranged around the annular wall of the hub. An axial air passage is defined in the stator. The holes of the base communicate with the holes of the top wall via the air passage. | 09-16-2010 |
| 20100239412 | Film-Cooling Augmentation Device and Turbine Airfoil Incorporating the Same - A turbine airfoil is disclosed. The airfoil includes one of a turbine shroud, liner, vane or blade, including an airfoil sidewall having a film-cooling hole that extends between an airfoil cooling circuit and an airfoil surface. The airfoil also includes an insert disposed in the film-cooling channel having a body. The body has a proximal end configured for disposition proximate the airfoil surface and a distal end. The body is also configured to define a passageway that extends between the distal end and proximal end upon disposition in the film-cooling hole. | 09-23-2010 |
| 20100247290 | TURBINE BLADE AND GAS TURBINE - A turbine blade and a gas turbine are provided in which the velocity of cooling fluid at the inlet of a pin fin region is improved so that the cooling performance at the trailing edge of the turbine blade can be improved. It includes an airfoil; a supply channel extending through the interior of the airfoil in the span direction, through which cooling fluid flows; a pin fin channel extending from the supply channel along the center line of the airfoil toward the trailing edge of the airfoil and opening at the trailing edge to the exterior of the airfoil; a plurality of gap pin fins projecting from a pair of opposing inner walls that constitute the pin fin channel at a region at the supply channel side of the pin fin channel and forming a gap therebetween extending in the span direction; pin fins connecting the pair of opposing inner walls at a region at the trailing edge side of the pin fin channel; and an insertion portion disposed in the gap to decrease the area of the channel of the cooling fluid at the region at the supply channel side of the pin fin channel. | 09-30-2010 |
| 20100254801 | COOLED AEROFOIL FOR A GAS TURBINE ENGINE - A cooled aerofoil for a gas turbine engine has an aerofoil section with pressure and suction surfaces extending between inboard and outboard ends thereof. The aerofoil section includes first and second internal passages for carrying cooling air. The aerofoil section further includes a plurality of holes in the external surface of the aerofoil section which receive cooling air from the internal passages. The external holes are arranged such that cooling air exiting a first portion of the external holes participates in a cooling film extending from the leading edge of the aerofoil section over said pressure surface and cooling air exiting from a second portion of the external holes participates in a cooling film extending from the leading edge over said suction surface. The first portion of external holes receives cooling air from the first internal passage, and the second portion of external holes receives cooling air from the second internal passage. The first and second internal passages are supplied with cooling air from respective and separate passage entrances. Each entrance is located at either the inboard end or the outboard end of the aerofoil section. | 10-07-2010 |
| 20100254802 | ROTOR ARRANGEMENT - With highly loaded rotors and stators problems can occur with secondary flows sweeping low momentum fluid across the blades reducing efficiency. By provision of collector slots to collect the secondary air and direct that air to a return slot in a rotor hub it is possible to provide impetus to the collected secondary flow to an outlet slot such that there is dispersal of the secondary flow and therefore reduce the effects upon the overall performance of a gas turbine engine incorporating the arrangement. | 10-07-2010 |
| 20100266385 | Separation resistant aerodynamic article - An airfoil disclosed herein comprises a pressure surface | 10-21-2010 |
| 20100266386 | FLANGE COOLED TURBINE NOZZLE - A turbine nozzle includes outer and inner bands bounding nozzle vanes. The outer band includes an aft flange. An impingement baffle bridges the outer band and aft flange at the root thereof to provide impingement cooling. | 10-21-2010 |
| 20100266387 | TURBINE ENGINE ROTATING CAVITY ANTI-VORTEX CASCADE - A gas turbine engine rotor drum includes spaced apart discs providing a cavity between the discs. The discs are configured to rotate in a rotational direction about an axis. An annular support is mounted on at least one of the discs and within the cavity. A cascade of relatively short anti-vortex members is mounted circumferentially on the annular support. The anti-vortex members are tubular in shape and provide a radially extending passage. The anti-vortex members include an outer end having a circumferential side with an opening in fluid communication with the radial passage. | 10-21-2010 |
| 20100278631 | TURBINE ENGINE HAVING COOLING PIN - In one embodiment, a turbine system may include a turbine casing, a shroud block coupled to the turbine casing, a fluid passage in the shroud block; and a pin configured to interface with the fluid passage. The pin may include a hollow shaft; a rod inserted into the hollow shaft; and a valve disposed on a distal end of the rod, wherein the valve is configured to open and close the fluid passage when the rod is actuated remotely through the hollow shaft. | 11-04-2010 |
| 20100303610 | COOLED GAS TURBINE STATOR ASSEMBLY - A stator assembly for a gas turbine engine is provided having an annular body, an inner gas path platform, a plurality of fairings, and at least one nozzle. The annular body has an outer gas path platform and a circumferentially extending annular cavity disposed radially outside of the outer gas path platform. The fairings extend radially between the inner gas path platform and the outer gas path platform. Each fairing includes a gas passage extending from the annular cavity through the inner gas path platform. The at least one nozzle has an inlet orifice disposed outside of the annular cavity and an exit orifice disposed within the annular cavity. The exit orifice is oriented within the annular cavity such that cooling air exiting the nozzle travels in a substantially circumferential direction within the annular cavity. | 12-02-2010 |
| 20100316486 | COOLED COMPONENT FOR A GAS TURBINE ENGINE - There is disclosed a cooled component for a gas turbine engine, the component preferably taking the form of a shrouded turbine blade, and having a segment region defining a segment of an annulus for the passage of hot gases therethrough. The segment region has a pair of opposed side faces configured to lie substantially adjacent respective corresponding side faces of the segments of similar operationally and circumferentially adjacent components when a series of such components are mounted in an engine such that their respective segments define an annulus. The component of the present invention is characterised by the provision of an elongate cooling slot in at least one of said side faces, said cooling slot being arranged in fluid communication with at least one flow passage within said segment region for the supply of cooling fluid to said slot, the slot being substantially closed at its upstream end and open at its downstream end so as to define an outlet for said cooling fluid at the operationally downstream region of said side face. | 12-16-2010 |
| 20100329846 | TURBINE ENGINE COMPONENTS - A turbine engine component includes a wall, a main opening, and two clusters of two or more auxiliary openings. The wall includes cool and hot air sides. The main opening extends between the cool air side and the hot air side and has an inlet and an outlet. The inlet is formed on the cool air side, and the outlet is formed on the hot air side. The first cluster of two or more auxiliary openings extends from the main opening to the hot air side. The second cluster of two or more auxiliary openings extends from the main opening to the hot air side. The main opening may be cylindrical or conical with a converging passage extending from the cool air side to the hot air side. The converging main opening may enhance flow through the auxiliary openings especially at high blowing ratios. | 12-30-2010 |
| 20100329847 | STATIONARY BLADE AND STEAM TURBINE - A stationary blade and a steam turbine capable of reducing self-excited vibrations with a simple configuration are provided. A stationary blade has a cavity, extending in a blade-width direction, formed therein and slits communicating between the cavity and the outside. A wave-shaped plate spring that is in sliding contact with at least one of a pressure-side member and a suction-side member is provided between the pressure-side member, which is a portion on the pressure side of the cavity, and the suction-side member, which is a portion on the suction side of the cavity. When the stationary blade is elastically deformed, the wave-shaped plate spring causes friction between itself and at least one of the pressure-side member and the suction-side member. This friction attenuates relative positional displacement between the pressure-side member and the suction-side member. Thus, self-excited vibrations occurring at the stationary blade can be reduced. | 12-30-2010 |
| 20100329848 | SHROUDLESS BLADE - A shroudless blade for use in a compressor or fan of an axial flow gas turbine engine includes a treatment to the tips of the blades to improve the surge margin of the compressor. A series of cross bleed holes are formed at the tip of the blade, which may be a stator or rotor component, extending between the pressure and suction sides of the blades. Notches are provided at the tip of the blades to provide initiation sites for any fatigue cracks such that the cracks propagate from the base of these slots or notches and radially inwards into the body of the blades. The present invention provides an arrangement, which by provision of the cross bleed holes the propagation of the cracks is arrested. A preferred placement and orientation for the holes for best mechanical and aerodynamic performance is described. | 12-30-2010 |
| 20100329849 | TURBINE ROTOR - A turbine rotor which is easy to manufacture and has a high tolerable temperature is provided. A highly efficient steam turbine power plant is also provided. The turbine rotor is configured from a rotor shaft, an inner rotor disc constructed integrally with the rotor shaft, and an outer rotor disc which is welded to the inner rotor disc via a weld metal part and has a structure for fixing a turbine blade. The outer rotor disc preferably has a cooling hole which extends in an axial direction to penetrate the outer rotor disc over the thickness of the outer rotor disc. | 12-30-2010 |
| 20110038708 | TURBINE ENDWALL COOLING ARRANGEMENT - An airfoil is provided and includes an airfoil body having a pressure surface extendable between radial ends and a fluid path in an airfoil interior defined therein. The pressure surface is formed to further define a passage by which coolant is deliverable from the fluid path in the airfoil interior, in a perimetric direction from the pressure surface for the purpose of cooling a portion on the surface of the radial end. | 02-17-2011 |
| 20110038709 | Turbine Vane for a Gas Turbine Engine Having Serpentine Cooling Channels - A turbine vane for a gas turbine engine having an internal cooling system formed from at least one serpentine cooling channel with enhanced cooling elements. The serpentine cooling channel may include a first turn manifold with purge air discharge orifices inline with a first pass of the serpentine cooling channel. Cooling fluids may be used to cooling the leading edge of the vane and passed through the purge air discharge orifices to purge the rim cavity proximate to the endwall. The first turn manifold may also include a plurality of trip strips. The trips strips may be positioned on the suction and pressure sidewalls and may be offset from trip strips on the opposing sidewall. The cooling system may also include an aft purge rim orifice. | 02-17-2011 |
| 20110038710 | Application of Dense Vertically Cracked and Porous Thermal Barrier Coating to a Gas Turbine Component - A configuration for coating a turbine component such as a blade or vane with various forms of thermal barrier coating to provide enhanced temperature capability and increased strain tolerance is disclosed. A gas path surface of the platform, airfoil and airfoil fillet region are first coated with a bond coating. A dense vertically cracked (DVC) thermal barrier coating is then applied to at least the gas path surface of the platform and can be applied to the fillet region. A porous thermal barrier coating is then applied to at least the airfoil. The porous thermal barrier coating can also be applied over the DVC thermal barrier coating if desired. | 02-17-2011 |
| 20110044795 | TURBINE VANE PLATFORM LEADING EDGE COOLING HOLES - A vane for use in a gas turbine engine has a platform connected to an airfoil. There is a cooling passage for supplying cooling air to the platform. A cooling chamber supplies cooling air to a plurality of cooling slots at the platform. The cooling slots have a non-uniform cross section. | 02-24-2011 |
| 20110044796 | FLUIDFOIL TIP VORTEX DISRUPTION - A fluidfoil tip vortex disruption arrangement. There is a fluid inlet on a pressure side of the fluidfoil; and a fluid outlet in a low pressure region at or near the tip of the fluidfoil. Fluid exiting the fluid outlet inserts instability into a tip vortex to disrupt or destroy the vortex. | 02-24-2011 |
| 20110044797 | ELECTRICAL CONDUCTOR PATHS - A component, for example an aerofoil vane or other gas path structure in a gas turbine machine such as a gas turbine engine, in which an insulated electrical conductor path ( | 02-24-2011 |
| 20110044798 | TURBINE NOZZLE FOR A TURBOMACHINE - A turbine nozzle for a turbomachine, the nozzle including two coaxial platforms interconnected by radial vanes, an inner platform being connected to an annular partition that is festooned or crenellated and on which there is fastened an annular support carrying elements made of abradable material. The support is capable of sliding circumferentially over the partition between a mounting-and-dismounting position and a position for locking the support on the partition. | 02-24-2011 |
| 20110052372 | TURBINE DISC AND RETAINING NUT ARRANGEMENT - A turbine rotor for a gas turbine engine comprises a disc having a hub defining a central bore for receiving an engine shaft. A nut retains the disc on the shaft. The disc retaining nut has at least one cooling passage defined therein and disposed for directing a flow of cooling air passing through the bore of the disc. | 03-03-2011 |
| 20110052373 | HIGH-TURNING DIFFUSER STRUT WITH FLOW CROSS-OVER SLOTS - A turning strut for use in a diffuser of a turbine engine has a leading edge with first and second opposing surfaces depending therefrom that terminate at a trailing edge. Slots extend through the turning strut and reduce in volume from the first surface to the second surface. During turn down operation of the turbine engine, exhaust flow impacts the leading edge at a deviated swirl angle. This results in exhaust flow at the first surface being at a higher pressure than at the second surface, which causes exhaust flow to be induced through the slots. The reduction in slot volume causes exhaust flow through the slots to accelerate. This exhaust flow from the slots is combined with exhaust flow at the second surface. Thusly, momentum of exhaust flow at the second surface is increased to maintain the second laminar boundary layer at the second surface. | 03-03-2011 |
| 20110064564 | Pumps or Generators with Flow-Through Impellers - An axial flow pump includes a perforated impeller through which fluid being pumped flows. The pump can carry a stator which can interact with magnetic elements rotatably carried with the impeller to produce rotation thereof. | 03-17-2011 |
| 20110070069 | STEAM TURBINE HAVING ROTOR WITH CAVITIES - A steam turbine and a rotor are disclosed having steam extending internally along at least part of the rotor. The rotor includes an interface and a steam passage system formed in the rotor, the passage system including a first inlet flow passage to the interface, the first inlet flow passage configured to receive steam from a first region of an outer surface of the rotor, a first outlet flow passage from the interface, the first outlet flow passage configured to pass steam to a second region of the rotor, a second inlet flow passage to the interface, the second inlet flow passage configured to receive steam from a third region of the outer surface of the rotor, a second outlet flow passage from the interface, the second outlet flow passage configured to pass steam to a fourth region of the rotor. | 03-24-2011 |
| 20110076133 | TURBOMACHINE COMPRESSOR WITH AN AIR INJECTION SYSTEM - A turbomachine or high pressure compressor including a stator casing housing a plurality of compression stages that are spaced apart in an axial direction along the central axis of the turbomachine, each compression stage including a row of rotor blades followed by a row of stator vanes. The compressor further includes an air injection system including at least one air injection passage through the casing and including an outlet segment that opens out in an inclined manner upstream from and directed towards a row of rotor blades into a set-back zone of the inside face of the casing. | 03-31-2011 |
| 20110081228 | INTERTURBINE VANE WITH MULTIPLE AIR CHAMBERS - A gas turbine engine has a mid turbine frame disposed between turbine rotor assemblies. The mid turbine frame includes hollow airfoils radially extending through an annular gas path duct. The airfoils each include a double-walled leading edge structure to define a front chamber separated from a rear chamber defined in the remaining space within the airfoil. | 04-07-2011 |
| 20110097191 | METHOD AND STRUCTURE FOR COOLING AIRFOIL SURFACES USING ASYMMETRIC CHEVRON FILM HOLES - A film-cooled turbine structure is configured with one or more asymmetric chevron film cooling holes for improving film cooling for a variety of airfoil surfaces or airfoil regions, particularly in regions and applications where the surface fluid streamline curvature is significant. | 04-28-2011 |
| 20110103932 | STATOR BLADE FOR A GAS TURBINE AND GAS TURBINE HAVING SAME - A stator blade for a gas turbine with sequential combustion, has a blade airfoil which extends in the radial direction between a blade tip and a shroud, with cooling passages extending inside the blade airfoil, through which a cooling medium can flow for cooling the blade and can then discharge from the stator blade into the hot gas flow flowing through the turbine. The blade airfoil has a sharply curved shape in space in the radial direction, and three cooling passages, which extend in the radial direction, arranged inside the blade airfoil in series in the hot gas flow direction and are interconnected by deflection regions, which are arranged at ends of the blade airfoil, so that the cooling medium flows through the cooling passages one after the other, with change of direction. The cooling passages follow the curvature of the blade airfoil in space in the radial direction. | 05-05-2011 |
| 20110110761 | GAS TURBINE HAVING AN IMPROVED COOLING ARCHITECTURE - A thermal machine includes a hot gas channel; a shell bounding the hot gas channel; a cooling shirt surrounding the shell; and a cooling channel disposed between the shell and the cooling shirt and configured to convection cool the hot gas channel with a cooling medium, wherein the cooling shirt includes at least one local divergence in the guidance of the cooling medium so as to compensate for non-uniformities in at least one of a thermal load on the shell and a flow of the cooling medium in the cooling channel. | 05-12-2011 |
| 20110123310 | TURBINE AIRFOIL PLATFORM COOLING CORE - A gas turbine engine component has a platform and an airfoil extending from the platform. The platform has a pressure side and a suction side. A cooling passage is formed within the platform, and extends along a pressure side of the platform. Air leaves the passage through an air outlet on a suction side of the platform. | 05-26-2011 |
| 20110123311 | SERPENTINE CORED AIRFOIL WITH BODY MICROCIRCUITS - A gas turbine engine component has an airfoil that extends from a leading edge to a trailing edge, and a suction side and has a pressure side. There are cooling passages extending from a root of the airfoil toward a tip of the airfoil. The cooling passages include a straight passage extending from the root toward the tip and adjacent the leading edge. A serpentine passage has at least three connected paths and is spaced from the straight passage toward the trailing edge. A cooling circuit is provided between the pressure wall and each of the three serpentine paths, and the straight path. A cooling circuit is provided between the suction wall and the straight passage. There is no cooling between at least a downstream one of the at least three paths of the serpentine passage and the suction wall. | 05-26-2011 |
| 20110123312 | GAS TURBINE ENGINE COMPONENTS WITH IMPROVED FILM COOLING - An engine component includes a body; and a plurality of cooling holes formed in the body. At least one of the cooling holes has cross-sectional shape with a first concave portion and a first convex portion. | 05-26-2011 |
| 20110135446 | Castings, Casting Cores, and Methods - If a refractory metal core (RMC) is punched, the punching asymmetry may be reflected in an asymmetry of the cast article features cast by the punched features. The punched features may have a shear zone and a fracture zone. The shear zone of the RMC will cast a relatively narrow portion of the post near one end; whereas the fracture zone will cast a relatively broader portion near the other end. The broader portion will also have a relatively shallow transition to the adjacent face of the slot-like passageway cast by the RMC. Where there is a stress asymmetry in the cast article in-use, the punching direction may be chosen so that the relatively broad portions of the post fall along the relatively higher stress face of the passageway. | 06-09-2011 |
| 20110135447 | SYSTEM FOR REDUCING THE LEVEL OF EROSION AFFECTING A COMPONENT - A system for removing moisture from a steam/water mixture engaging a stationary component of a steam turbine. The system includes an airfoil located within a flow path of a steam turbine. The airfoil is configured for removing moisture from a steam/water mixture traveling in the flow path. To this end, the airfoil includes a cavity in flow communication with the steam path through at least one inlet and outlet opening, near the leading and trailing edge of the airfoil, respectively. Moisture and steam are extracted from the surface through the inlet openings, the steam and water are separated in the cavity, the separated water flows towards the bottom end, and the dry steam flows through the outlet opening and returns to the steam path. The dry steam blowing out of the trailing edge reduces the size of secondary droplets, and thereby prevents erosion. | 06-09-2011 |
| 20110142597 | TURBINE BLADE STRUCTURE - Provided is a turbine blade structure that is capable of suppressing quality variations of cast products during the manufacturing of turbine blades. A turbine blade structure wherein the space inside an air foil is divided into a plurality of cavities, partitioned by rib members provided substantially perpendicular to the center line connecting a leading edge and a trailing edge, is provided with partition members that partition the inside of the cavities located in the central portion of the blade, excluding the blade leading-edge side and the blade trailing-edge side, into blade pressure side cavities and blade suction side cavities substantially along the center line, wherein blade leading-edge end portions and blade trailing-edge end portions of the partition members are inserted from one shroud surface side to the other shroud surface side along engagement grooves formed on the rib members. | 06-16-2011 |
| 20110164960 | HEAT TRANSFER ENHANCEMENT IN INTERNAL CAVITIES OF TURBINE ENGINE AIRFOILS - An airfoil includes a leading edge, a trailing edge, a suction side and a pressure side; a plurality of internal cooling cavities extending radially within the airfoil, one of the plurality of internal cavities extending along the trailing edge. The trailing edge is provided with a plurality of coolant exit apertures extending therealong. A plurality of vortex generators is formed on an internal surface of at least one of the pressure and suction sides of the airfoil. The vortex generators are arranged in radially spaced relationship in one of the plurality of internal cooling cavities, extending substantially parallel to and in proximity to the plurality of coolant exit apertures. | 07-07-2011 |
| 20110171005 | STEAM TURBINE - According to an embodiment, at least one first outer ring has an annular outer ring cavity to which external cooling steam is supplied. A radial direction cooling hole connecting with the outer ring cavity is formed in the stator blades connected to the first outer ring. An annular inner ring cavity connecting with the radial direction cooling hole is formed in a first inner ring constituting one diaphragm together with the first outer ring. Cooling steam blowing holes connecting an annular wheel space and the inner ring cavity are formed. The annular wheel space is formed between the first inner ring and a rotor wheel adjacent to the first inner ring. | 07-14-2011 |
| 20110188993 | RING SECTOR OF TURBOMACHINE TURBINE - A turbine ring sector comprising:
| 08-04-2011 |
| 20110188994 | METHOD AND SYSTEM FOR DETERMINING GAS TURBINE TIP CLEARANCE - A system for sensing at least one physical characteristic associated with an engine including a turbine having a plurality of blades turning inside a casing, the system including: a pressure sensor coupled substantially adjacent to the casing and including at least one output; a port in the turbine casing for communicating a pressure indicative of a clearance between the blades and casing to the pressure sensor; a cooling cavity substantially surrounding the pressure sensor; and, an inlet for receiving fluid from the engine and feeding the fluid to the cooling cavity to cool the pressure sensor; wherein, the pressure sensor output is indicative of the clearance between the blades and casing. | 08-04-2011 |
| 20110217158 | COOLED TURBINE RIM SEAL - A rim seal assembly includes an annular seal element circumscribing an engine centerline or axis and mounted on an annular platform including a radially inwardly extending annular platform flange disposed between and connected to forward and aft flanges at distal ends of forward and aft annular elements respectively. Annular forward and aft outer rim cavities radially disposed between the forward and aft annular elements and the platform are axially separated by the platform flange. Cooling slots extend radially across axially facing forward and aft surfaces of the forward and aft flanges. The platform flange and the forward and aft flanges are bolted together. The annular seal element may include seal teeth in sealing relationship with an annular seal land such as in a labyrinth seal. The rim seal assembly may be incorporated in a low pressure turbine use a compressor as a source of cooling air. | 09-08-2011 |
| 20110217159 | PREFERENTIAL COOLING OF GAS TURBINE NOZZLES - Turbine nozzle assemblies include a plurality of circumferentially spaced first components and second components, which are designed to provide different amounts of cooling. The second components, which are generally aligned with an opening of transition pieces, are designed to provide more cooling than the first components, which are generally aligned with interfaces between the transition pieces. | 09-08-2011 |
| 20110223004 | APPARATUS FOR COOLING A PLATFORM OF A TURBINE COMPONENT - The present subject matter discloses a turbine component including a platform and an airfoil extending radially upward from the platform. A plurality of curved cooling passages may be defined in the platform. Each of the curved cooling passages may have at least one end disposed at an exterior surface of the platform. Additionally, each of the cooling passages may be configured to direct a cooling medium through the platform. | 09-15-2011 |
| 20110223005 | Airfoil Having Built-Up Surface with Embedded Cooling Passage - A component in a gas turbine engine includes an airfoil extending radially outwardly from a platform associated with the airfoil. The airfoil includes opposed pressure and suction sidewalls, which converge at a first location defined at a leading edge of the airfoil and at a second location defined at a trailing edge of the airfoil opposed from the leading edge. The component includes a built-up surface adjacent to the leading edge at an intersection between the pressure sidewall and the platform, and at least one cooling passage at least partially within the built-up surface at the intersection between the pressure sidewall and the platform. The at least one cooling passage is in fluid communication with a main cooling channel within the airfoil and has an outlet at the platform for providing cooling fluid directly from the main cooling channel to the platform. | 09-15-2011 |
| 20110229305 | COVER PLATE FOR TURBINE VANE ASSEMBLY - Embodiments of a cover plate and material blank for forming the cover plate may include features and characteristics to accommodate tolerance issues in a turbine vane assembly. In one embodiment, the cover plate is formed from a material blank configured to provide the cover plate with a flexible flange area that can be secured to the turbine vane assembly. The flange area may have a range of motion that may be responsive to an installation force, which effectively modifies the configuration of the cover plate to such degree as to seal the cover plate to the turbine vane assembly. | 09-22-2011 |
| 20110229306 | ROTOR BLADE TIP CLEARANCE CONTROL - A gas turbine engine has a row of circumferentially spaced rotor blades and a plurality of seal segments circumscribing the rotor blade tips and attached to a radially inward side of a casing of the engine. The seal segments are spaced from the casing by a spacing cavity. A flow of relatively hot cooling air is routed to the spacing cavity to cool the seal segments. The engine has an external cooling arrangement for impinging relatively cold cooling air on a radially outward side of the casing. The engine has a wall containing a plurality of through-holes which is attached to a radially inward side of the casing adjacent the seal segments. The wall is spaced from the casing to define a heating control chamber between the wall and the casing. The engine has one or more closable ducts which allow air to be exhausted from the heating control chamber. | 09-22-2011 |
| 20110243711 | INTERIOR COOLING CHANNELS - Cooling channels through the interior of a machine component that include: a first set of cooling channels, the first set of cooling channels including a plurality of parallel channels that reside in a first plane; a second set of cooling channels, the second set of cooling channels including a plurality of parallel channels that reside in a second plane. Along a longitudinal axis, the cooling channels of the first and second set of cooling channels may include an alternating diverging-converging configuration, the alternating diverging-converging configuration creating a series of broader chamber sections connected by a series of narrower throat sections. The first set of cooling channels and the second set of cooling channels may be configured such that, when viewed from the side, a crisscrossing pattern with a plurality of intersections is formed. The first plane resides in spaced relation to the second plane, with the first plane being offset from the second plane such that a plurality of the chamber sections of the first set of cooling channels connect to a plurality of the chamber sections of the second set of cooling channels. | 10-06-2011 |
| 20110250053 | FLUID TURBINES - Shrouded fluid turbines of various configurations are disclosed. The shrouded fluid turbines include an impeller, a turbine shroud surrounding the impeller, and an ejector shroud around the turbine shroud. The ejector shroud may completely surround the turbine shroud. The turbine shroud may have a plurality of mixing lobes that form a crenellated trailing edge. Alternatively, the turbine shroud may have a plurality of open slots. Means for directing fluid flow into the plurality of open slots may include an ejector shroud that seals with the turbine shroud downstream of the open slots. A plurality of fluid ducts may also connect individually to each open slot. An external stator may be connected to an exterior surface of the ejector shroud. | 10-13-2011 |
| 20110255956 | Gas turbine having cooling insert - A gas turbine including a plurality of rotor blades assembled into rotor blade rows and arranged on a turbine shaft and including a plurality of guide vanes assembled into guide van rows and mounted on a turbine housing by means of a guide van carrier is provided. The guide vane carrier includes a plurality of cooling air holes, and has a particularly high efficiency, while maintaining maximum operating reliability. Therefore, a cooling insert is introduced into a cooling air hole. | 10-20-2011 |
| 20110262265 | INSTALLATION HAVING A THERMAL TRANSFER ARRANGEMENT - A gas turbine engine | 10-27-2011 |
| 20110286834 | GUIDE VANE FOR A GAS TURBINE - A guide vane is provided for a gas turbine and has an airfoil extending in the radial direction between an inner platform and an outer platform. The airfoil extends transversely to the direction of the hot gas flow between a leading edge and a trailing edge and has a pressure side and a suction side. A cooling slot running parallel to the trailing edge is provided on the pressure side in front of the trailing edge, a cooling medium can exit through the cooling slot from the guide vane over the entire length of the guide vane and can cool the trailing edge of the guide vane. In such a guide vane, the service life is extended by a thermal stress reducing element provided on the inner platform below the trailing edge and the cooling slot. | 11-24-2011 |
| 20110311349 | ROTOR ELEMENT WITH A FLUID PASSAGE AND PASSAGE-BLOCKING MEMBER AND TURBINE ENGINE INCLUDING THE ROTOR ELEMENT - A rotor element including an annular surface portion about the rotor rotational axis, a fluid passage being formed through the surface portion, a passage-blocking mechanism including a blocking element that is deformable depending on the rotor rotating speed and arranged so as to adjust the fluid flow depending on the rotor rotating speed, and an annular collar with a free edge engaging with the blocking element so as to form the blocking mechanism. The free edge of the annular collar defines, together with the blocking portion of the blocking element, a diaphragm that blocks the fluid passage. | 12-22-2011 |
| 20120003077 | ANNULAR MULTI-ROTOR DOUBLE-WALLED TURBINE - An annular single or multi-rotor double-walled turbine. The turbine includes an outer shroud, an inner shroud, and a plurality of driveshafts. The turbine also includes a plurality of rotors coaxially attached to the plurality of driveshafts at spaced intervals. Each of the plurality of rotors comprises a plurality of turbine blades extending between the inner and outer shrouds. Each of the plurality of turbine blades comprises a face. The inner shroud and the outer shroud form a continuous channel for directing a fluid entering the turbine towards the faces of the turbine blades and for directing fluid discharged from a first of the plurality of rotors to the remaining rotors. The channel greatly improves efficiency of power extraction from all augmented and non-augmented fluid streams. | 01-05-2012 |
| 20120014780 | FAN DOWNSTREAM GUIDE VANES OF A TURBOFAN ENGINE - Fan downstream guide vane profiles have an optimized form of skeleton line angle distribution in an area situated between an upper and a lower limitation as well as a specific thickness distribution superimposed on the respective skeleton line angle distribution. Such guide vanes are characterized by lower pressure losses and a larger working range than the known downstream guide vanes, thereby reducing fuel consumption of the engine and increasing the operating stability thereof. | 01-19-2012 |
| 20120034068 | VENTILATION INLET - A ventilation inlet comprising a ventilation pipe to receive flow from a first flow zone and to deliver the flow to a second flow zone; a divider arranged to divide a portion of the ventilation pipe into a static pressure zone and a total pressure zone; and a deflector arranged to direct flow from the total pressure zone at least partially across the static pressure zone to restrict delivery of the flow from the static pressure zone to the second flow zone dependent on the pressure of the flow in the first flow zone. | 02-09-2012 |
| 20120057960 | RING SEGMENT WITH FORKED COOLING PASSAGES - A ring segment is provided for a gas turbine engine includes a panel and a cooling system. Cooling fluid is provided to an outer side of the panel and an inner side of the panel defines at least a portion of a hot gas flow path through the engine. The cooling system is located within that panel and receives cooling fluid from the outer side of the panel for cooling the panel. The cooling system includes a plurality of cooling fluid passages that receive cooling fluid from the outer side of the panel. The cooling fluid passages each have a generally axially extending portion that includes at least one fork. The fork(s) divide each cooling fluid passage into at least two downstream portions that each receives cooling fluid from the respective axially extending portion. | 03-08-2012 |
| 20120057961 | TURBINE STAGE SHROUD SEGMENT - A shroud segment for a turbine stage of a gas turbine engine forms an endwall for the working gas annulus of the stage. The segment also provides a close clearance to the tips of a row of turbine blades which sweep across the segment. In use, a mainstream flow of the working gas passes through the passages formed between adjacent turbine blades. The segment has a plurality of cooling holes and respective air feed passages for the cooling holes. The cooling holes are distributed over that part of the gas-washed surface of the segment which is swept by the blade tips. The cooling holes deliver, in use, cooling air which spreads over the gas-washed surface. The feed passages are configured such that the delivered air has swirl directions which are co-directionally aligned with the swirl directions of the mainstream flow at the segment. | 03-08-2012 |
| 20120121381 | TURBINE TRANSITION COMPONENT FORMED FROM AN AIR-COOLED MULTI-LAYER OUTER PANEL FOR USE IN A GAS TURBINE ENGINE - A cooling system for a transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine is disclosed. The transition duct may have a multi-panel outer wall formed from an inner panel having an inner surface that defines at least a portion of a hot gas path plenum and an intermediate panel positioned radially outward from the inner panel such that at least one cooling chamber is formed between the inner and intermediate panels. The transition duct may also include an outer panel. The inner, intermediate and outer panels may include one or more metering holes for passing cooling fluids between cooling chambers for cooling the panels. The intermediate and outer panels may be secured with an attachment system coupling the panels to the inner panel such that the intermediate and outer panels may move in-plane. | 05-17-2012 |
| 20120128467 | INTEGRATED VARIABLE GEOMETRY FLOW RESTRICTOR AND HEAT EXCHANGER - One or more heat exchangers mounted in a duct have heat transfer cooling passages therein and a variable geometry flow restrictor is integral with each of the heat exchangers. An annular slide valve axially translatable within the duct is operable to open and close or vary a variable area between the heat exchangers and one of inner and outer casings bounding the duct. The heat exchangers may be being circumferentially distributed around an annular duct and include radial or circumferentially curved heat transfer tubes or vanes. | 05-24-2012 |
| 20120134777 | ENGINE CASE WITH WASH SYSTEM - A gas turbine engine includes a structure defining a circumferential passage in fluid communication with an internal passage in at least one strut radially extending into the engine, circumferential passage also in fluid communication with a plurality or nozzles or jets to provide a wash manifold integrated with the engine casing structure. One or more nozzles are provided in the manifold for directing a washing fluid injected into the duct. | 05-31-2012 |
| 20120134778 | AXIAL FLOW GAS TURBINE - An axial flow gas turbine ( | 05-31-2012 |
| 20120134779 | GAS TURBINE OF THE AXIAL FLOW TYPE - In an axial flow gas turbine ( | 05-31-2012 |
| 20120134780 | AXIAL FLOW GAS TURBINE - In an axial flow gas turbine efficient cooling and a long life-time can be achieved by providing the outer blade platforms ( | 05-31-2012 |
| 20120134781 | AXIAL FLOW GAS TURBINE - In an axial flow gas turbine ( | 05-31-2012 |
| 20120148383 | GAS TURBINE VANE WITH COOLING CHANNEL END TURN STRUCTURE - A vane structure for a gas turbine engine. The vane structure includes a radially outer platform and a radially inner platform, and an airfoil having an outer wall extending radially between the outer platform and the inner platform. A cooling passage is defined within the outer wall and has a plurality of radially extending channels. An outer end turn structure is located at the outer platform to conduct cooling fluid in a chordal direction between at least two of the channels. The outer end turn structure includes an enlarged portion wherein the enlarged portion is defined by an enlarged dimension, in a direction transverse to the chordal direction, between the at least two channels. | 06-14-2012 |
| 20120177478 | IMPINGEMENT PLATE FOR TURBOMACHINE COMPONENTS AND COMPONENTS EQUIPPED THEREWITH - An impingement plate adapted to reduce thermally-induced strains and stresses that may damage the plate or its attachment to a second component. The plate includes an interior region having cooling holes, a peripheral wall surrounding the interior region and projecting out of the plane of the interior region, a peripheral flange surrounding the peripheral wall and lying in a plane spaced apart from the plane of the interior region, and one or more through-thickness rib. One such rib may be disposed in the interior region, project away from and out of the plane of the interior region, and linearly extend across the interior region. Alternatively or in addition, one such rib may be disposed between the peripheral wall and flange and project out of the plane of the flange. | 07-12-2012 |
| 20120177479 | INNER SHROUD COOLING ARRANGEMENT IN A GAS TURBINE ENGINE - A component in a gas turbine engine includes an airfoil and a shroud. The shroud has an outer surface supporting an end of the airfoil and defines a portion of an annular gas path. The shroud includes axial edges extending between upstream and downstream edges thereof. Each of the axial edges includes a seal slot that receives a seal member extending between the shroud and an adjacent shroud. A cooling air channel extends between the upstream and downstream edges of the shroud. A cooling air supply passage extends from a cooling air chamber at an inner surface of the shroud to the cooling air channel. At least one cooling air exit passage extends from the cooling air channel to one of the axial edges. The cooling air channel is located radially between the outer surface of the shroud and the seal slot. | 07-12-2012 |
| 20120177480 | ROTOR WITH COOLING PASSAGE - A gas turbine engine is disclosed having a cooling passage that rotates with a turbine and is capable of providing cooling flow to the turbine. In one embodiment the cooling passage can receive cooling flow from an interior of a shaft of the gas turbine engine and increase the pressure of the cooling flow before delivering it to a location near a blade of the turbine. In one form the cooling passage can have an inducer section. In one form the cooling passage can have internal vanes useful in increasing the pressure of the cooling flow. | 07-12-2012 |
| 20120183389 | SEAL SYSTEM FOR COOLING FLUID FLOW THROUGH A ROTOR ASSEMBLY IN A GAS TURBINE ENGINE - A sealing system for a rotor assembly in a gas turbine engine is disclosed. The sealing system may include a seal formed from a side block and an upper seal that seals a gap between a radially outward extending first rotor supply channel in a rotor assembly terminating at an inlet of an axially extending second rotor supply channel that is in fluid communication with an internal blade cooling system of a turbine blade. The seal may include components that enhance the flow of cooling fluids over conventional configurations. In another embodiment, the sealing system may include an integrated sealing block configured to seal a gap between adjacent turbine blades at an intersection between the first and second rotor supply channels. The integrated sealing block may be formed from a radially inward extending leg and central body. | 07-19-2012 |
| 20120195737 | Gas turbine engine - A gas turbine engine including a segment of an annular guide vane assembly is provided. When the engine is used, the segment directs hot combustion gases onto rotor blades of the engine. The segment includes a platform disposed at a side of the segment radially inward/outward with respect to the axis of rotation of the engine. The platform has a trailing edge portion downstream with respect to the flow of hot combustion gases through the segment, the trailing edge portion includes a rail that extends radially inwardly/outwardly from the trailing edge portion. The engine also includes a support and cooling arrangement for supporting the segment and directing a cooling fluid to cool the segment. The arrangement is located radially inward/outward of the platform, and includes a flange part that extends radially outwardly/inwardly from the arrangement. The arrangement further includes a leaf seal and a retaining pin. | 08-02-2012 |
| 20120201652 | CROSS-OVER PURGE FLOW SYSTEM FOR A TURBOMACHINE WHEEL MEMBER - A wheel member includes a body having a first surface that extends to a second surface through an intermediate portion. The body includes an outer diametric surface and a central bore. A first plurality of purge circuits are formed in the body. The first plurality of purge circuits extend from a first end to a second end through the body. The first plurality of purge circuits are arranged to direct a first purge flow in a first direction. A second plurality of purge circuits are formed in the body and fluidly isolated from the first plurality of purge circuits. The second plurality of purge circuits extend from a first end portion to a second end portion through the body and are arranged to direct a second purge flow in a second direction, that is distinct from the first direction, to establish a cross-over purge flow system. | 08-09-2012 |
| 20120201653 | GAS TURBINE ENGINE AND COOLED FLOWPATH COMPONENT THEREFOR - One embodiment of the present invention is a unique gas turbine engine. Another embodiment of the present invention is a unique cooled gas turbine engine flowpath component. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and cooled gas turbine engine flowpath components. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. | 08-09-2012 |
| 20120213626 | EXPLOSION-WELDED GAS TURBINE SHROUD AND A PROCESS OF FORMING AN EXPLOSION-WELDED GAS TURBINE - An explosion-welded turbine shroud and a process of forming an explosion-welded gas turbine shroud are disclosed. The explosion-welded gas turbine shroud includes a first alloy explosion welded to a second alloy. In the explosion-welded gas turbine shroud, the first alloy forms at least a portion of a hot gas path or an expansion region of the gas turbine shroud includes the first alloy. The process includes explosion welding a first alloy to a second alloy to form the gas turbine shroud. | 08-23-2012 |
| 20120219401 | ENDWALL COMPONENT FOR A TURBINE STAGE OF A GAS TURBINE ENGINE - A component of a turbine stage of a gas turbine engine is provided, the component forming an endwall for the working gas annulus of the stage. The component has one or more internal plena behind the endwall which, in use, contain a flow of cooling air. The component further has a plurality of exhaust holes in the endwall. The holes connect the plena to a gas-washed surface of the endwall such that the cooling air effuses through the holes to form a cooling film over the gas-washed surface. Each exhaust hole has a flow cross-sectional area which is greater at an intermediate position between the entrance of the hole from the respective plenum and the exit of the hole to the gas-washed surface than it is at said exit. | 08-30-2012 |
| 20120219402 | VANE - A vane is provided for directing hot gases in a gas turbine engine. The vane includes a hollow aerofoil portion, which in use spans the working gas annulus of the engine. The vane further includes an impingement tube which forms a covering over the interior surface of the aerofoil portion and which has jet-forming apertures formed therein for the production of impingement cooling jets. The impingement tube includes two tube portions which are separately insertable into position into the aerofoil portion to form the covering. The o impingement tube further includes an expansion member which, when the tube portions are in position in the aerofoil portion, is locatable in the aerofoil portion to urge each tube portion outwardly and thereby holds the tube portions in position against the aerofoil portion. | 08-30-2012 |
| 20120237333 | COOLED GAS TURBINE ENGINE COMPONENT - A gas turbine engine component is disclosed having a cooling fluid passageway that provides relatively cool fluid to a surface of the gas turbine engine component. The cooling fluid passageway can be shaped in cross section to reduce a stress present in the gas turbine engine component. One form of the shape is non-circular. The gas turbine engine component can be formed such that an overhanging material otherwise formed by the intersection of a cooling fluid passageway and a surface of the gas turbine engine component is absent. The gas turbine engine component can also have a depression formed near the surface of the gas turbine engine component such that the cooling fluid passageway exits into an upstream portion and a downstream portion of the depression. | 09-20-2012 |
| 20120251295 | GAS TURBINE ENGINE COMPONENT - A component of a gas turbine engine is provided. The component includes an external wall which, in use, is exposed on one surface thereof to working gas flowing through the engine. The component further includes effusion cooling holes formed in the external wall. In use, cooling air blows through the cooling holes to form a cooling film on the surface of the external wall exposed to the working gas. The component further includes an air inlet arrangement which receives the cooling air for distribution to the cooling holes. The component further includes a plurality of metering feeds and a plurality of supply plena. The metering feeds meter the cooling air from the air inlet arrangement to respective of the supply plena, which in turn supply the metered cooling air to respective portions of the cooling holes. | 10-04-2012 |
| 20120263575 | LOW PRESSURE COOLING SEAL SYSTEM FOR A GAS TURBINE ENGINE - A low pressure cooling system for a turbine engine for directing cooling fluids at low pressure, such as at ambient pressure, through at least one cooling fluid supply channel and into a cooling fluid mixing chamber positioned immediately downstream from a row of turbine blades extending radially outward from a rotor assembly to prevent ingestion of hot gases into internal aspects of the rotor assembly. The low pressure cooling system may also include at least one bleed channel that may extend through the rotor assembly and exhaust cooling fluids into the cooling fluid mixing chamber to seal a gap between rotational turbine blades and a downstream, stationary turbine component. Use of ambient pressure cooling fluids by the low pressure cooling system results in tremendous efficiencies by eliminating the need for pressurized cooling fluids for sealing this gap. | 10-18-2012 |
| 20120263576 | TURBINE SHROUD SEGMENT COOLING SYSTEM AND METHOD - The present embodiments are generally directed toward systems and methods for cooling one or more shroud segments of a gas turbine engine. For example, in a first embodiment, a shroud segment is provided that is configured to at least partially surround a turbine blade of a turbine engine. The shroud segment includes a body and a microchannel disposed in the body. The microchannel is configured to flow a cooling fluid through the body. | 10-18-2012 |
| 20120301275 | INTEGRATED CERAMIC MATRIX COMPOSITE ROTOR MODULE FOR A GAS TURBINE ENGINE - A rotor module for a gas turbine engine includes a multiple of CMC airfoil rows which extend from a common CMC drum. | 11-29-2012 |
| 20120308360 | OVERLAP SEAL FOR TURBINE NOZZLE ASSEMBLY - Systems for thermally regulating portions of a turbine are disclosed. In one embodiment, a turbine nozzle assembly includes: an outer diaphragm ring; a vane physically connected to the outer diaphragm ring; and an inner diaphragm ring physically connected to the vane, the inner diaphragm ring including a first axial tooth configured to interact and substantially form a seal with a second axial tooth disposed on a bucket shank. | 12-06-2012 |
| 20120321441 | VENTILATED COMPRESSOR ROTOR FOR A TURBINE ENGINE AND A TURBINE ENGINE INCORPORATING SAME - A turbine engine includes a plurality of compressor rotors that include ventilation slots to vent the spaces between adjacent compressor rotors. Each compressor rotor is formed from a flat disk of material having first and second circular faces. A circular ridge of material protrudes outward from the one of the circular faces of the disc adjacent an outer edge of the disc. The ventilation slots are formed in the circular ridge of material. Each ventilation slot is a depression in the circular ridge of material, the depression having a longitudinal axis that extends substantially in a radial direction of the disc. | 12-20-2012 |
| 20120328413 | SYSTEM AND METHOD FOR SUPPORTING A NOZZLE ASSEMBLY - A system for supporting a nozzle assembly includes a first member connected to a stationary component and a second member extending from the first member radially through at least a portion of the nozzle assembly. A distal end of the second member is radially displaced from the first member and configured to contact the nozzle assembly. A method for supporting a nozzle assembly includes connecting a first member to a stationary component and extending a second member from the first member radially through at least a portion of the nozzle assembly. The method further includes contacting a distal end of the second member to the nozzle assembly, wherein the distal end is radially displaced from the first member. | 12-27-2012 |
| 20130004294 | DUCTILE ALLOYS FOR SEALING MODULAR COMPONENT INTERFACES - A vane assembly ( | 01-03-2013 |
| 20130004295 | TURBINE VANE - A stator for a turbine includes an arrangement of vanes including at least a first vane and a second vane circumferentially neighbouring the first vane. Each of the first vane and the second vane include: an airfoil; a channel system configured to cool the respective vane with cooling gas; and an inner diameter platform disposed at an inner end of the airfoil, the inner diameter platform including an inner diameter platform cavity and a circumferentially arranged side wall which delimits the inner diameter platform cavity, the inner diameter platform cavity being connected with the channel system so as to feed the cooling gas to the inner diameter platform. At least one sealing plate is disposed between the circumferentially arranged side walls of the first vane and the second vane so as to form an intermediate cavity that is fluidically separated from the inner diameter platform cavities. | 01-03-2013 |
| 20130004296 | SEGMENTED CERAMIC MATRIX COMPOSITE TURBINE AIRFOIL COMPONENT - A segmented component for use with a gas turbine engine comprises a radially extending gas path portion. The gas path portion is for interacting with gas flow from the gas turbine engine. The gas path portion comprises a forward portion forming a leading edge of a stationary vane, an aft portion forming a trailing edge of the stationary vane, and a plurality of middle portions forming a pressure side and a suction side of the stationary vane. The component is divided into axially aligned segments comprising a forward segment, an aft segment, and a plurality of middle segments disposed between the forward segment and the aft segment. The middle segments comprise radially elongate ceramic matrix composite material plates. | 01-03-2013 |
| 20130011238 | COOLED RING SEGMENT - A ring segment for a gas turbine engine includes a panel and a first mating edge cooling system. Cooling fluid is provided to an outer side of the panel and an inner side of the panel defines at least a portion of a hot gas flow path. A cooling system receives a portion of the cooling fluid provided to the outer side and includes at least one impingement chamber. Each impingement chamber includes at least one metering supply passage and at least one metering discharge passage. The metering supply passage(s) extends from the outer side of the panel to the impingement chamber. Cooling fluid impinges on a surface of the panel defining the impingement chamber as it flows therein through the metering supply passage(s). The metering discharge passage(s) extends from the impingement chamber to a first or second mating edge of the panel. | 01-10-2013 |
| 20130017064 | GAS TURBINE AIRFOIL WITH SHAPED TRAILING EDGE COOLANT EJECTION HOLES - A turbine blade or vane includes at least one internal radial channel for the circulation of cooling medium bordered on a pressure side by a pressure side wall and on a suction side by a suction side wall joined at a upstream side at a leading edge and at and downstream side at the trailing edge. At least one exit hole extends through at least one of the pressure side wall or the suction side wall for blowing out of cooling medium from the internal radial channel to a medium surrounding the blade or vane. At least one trailing edge exit hole along the trailing edge has a surfacial exit opening disposed at the pressure side of the trailing edge. | 01-17-2013 |
| 20130022450 | PUMP IMPELLER AND SUBMERSIBLE PUMP HAVING SUCH PUMP IMPELLER - A non-clogging type pump impeller ( | 01-24-2013 |
| 20130045083 | TURBINE ROTOR DISK INLET ORIFICE FOR A TURBINE ENGINE - A turbine rotor body having at least one inlet orifice in fluid communication with a pre-swirl system such that the inlet orifice receives cooling fluids from the pre-swirl system is disclosed. The inlet orifice may be configured to reduce the relative velocity loss associated with cooling fluids entering the inlet orifice in the rotor, thereby availing the cooling system to the efficiencies inherent in pre-swirling the cooling fluids to a velocity that is greater than a rotational velocity of the turbine rotor body. As such, the system is capable of taking advantage of the additional temperature and work benefits associated with using the pre-swirled cooling fluids having a rotational speed greater than the turbine rotor body. | 02-21-2013 |
| 20130051979 | TURBINE SHROUD SEGMENT WITH INTEGRATED IMPINGEMENT PLATE - A turbine shroud segment is metal injection molded (MIM) about an insert having a cooling air cavity covered by an impingement plate. The insert is held in position in an injection mold and then the MIM material is injected in the mold to form the body of the shroud segment about the insert. | 02-28-2013 |
| 20130051980 | HIGH-PRESSURE TURBINE NOZZLE FOR A TURBOJET - A high temperature turbine nozzle with automatic regulation of flow of cooling air passing therethrough. Each of vanes includes a first sleeve drilled with holes and a second sleeve that is engaged in the first sleeve, drilled with corresponding holes, and made of a material possessing a coefficient of expansion different from that of the first sleeve. | 02-28-2013 |
| 20130108413 | SECONDARY FLOW ARRANGEMENT FOR SLOTTED ROTOR | 05-02-2013 |
| 20130136579 | EXHAUST-GAS TURBOCHARGER - An exhaust-gas turbocharger ( | 05-30-2013 |
| 20130149106 | STEAM TURBINE, BLADE, AND METHOD - A stator blade ring comprising a plurality of stator blade modules defining an annular chamber is provided. The plurality of stator blade modules comprises an elongated blade portion comprising a first and a second blade shell portion, a longitudinal passageway, and at least one opening extending through at least one of the first and the second blade shell portion to the longitudinal passageway, an inner portion brazed to a first longitudinal end of the elongated blade portion, wherein the inner portion comprises a through hole forming a portion of the annular chamber, and an inner passageway extending from the through hole to the longitudinal passageway, and an outer portion brazed to a second longitudinal end of the elongated blade portion and engaged to a steam turbine, the outer portion comprising an outer passageway open to a surface of the steam turbine and the longitudinal passageway. | 06-13-2013 |
| 20130156549 | USE OF MULTI-FACETED IMPINGEMENT OPENINGS FOR INCREASING HEAT TRANSFER CHARACTERISTICS ON GAS TURBINE COMPONENTS - An improved nozzle vane for a gas turbine engine, comprising a vane wall having inner and outer wall surfaces, the wall surfaces being spaced from one another to define a plurality of fluid passageways for a cooling medium; discreet cavities formed by interior wall members disposed between the inner and outer wall surfaces and within the fluid passageway for the cooling medium; a plurality of impingement cooling sleeves disposed in the discreet cavities defined by the inner and outer wall surfaces and by interior wall members; and a plurality of non-round, e.g., serrated, openings in each of the impingement cooling sleeves, with the openings being sufficient in size and number to accommodate the flow of a cooling media. | 06-20-2013 |
| 20130164116 | HIGH PRESSURE TURBINE VANE COOLING HOLE DISTRIBUTION - A turbine vane for a gas turbine engine with an airfoil portion including a perimeter wall having first, second, and third sets of cooling holes defined therethrough, including the holes numbered HA-1 to HA-13, HB-1 to HB-13 and PA-1 to PA-9, 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, Z Cartesian coordinate values set forth in Table 3. | 06-27-2013 |
| 20130170954 | High Pressure Compressor - A high-pressure compressor ( | 07-04-2013 |
| 20130177396 | Impingement Cooling System for Use with Contoured Surfaces - The present application provides an impingement cooling system for use with a contoured surface. The impingement cooling system may include an impingement plenum and an impingement plate with a linear shape facing the contoured surface. The impingement surface may include a number of projected area thereon with a number of impingement holes having varying sizes and varying spacings. | 07-11-2013 |
| 20130177397 | SLOTTED TURBINE AIRFOIL - A slotted turbine static nozzle airfoil. In one embodiment, the turbine static nozzle airfoil includes a concave pressure wall having a slot extending therethrough; a convex suction wall adjoined with the concave pressure wall at respective end joints; and a pocket fluidly connected with the slot and located between the convex suction wall and the concave pressure wall, wherein at least one of the convex suction wall or the concave pressure wall includes a thinned segment proximate one of the respective end joints, the thinned segment configured to extend the pocket toward a trailing edge of the turbine static nozzle airfoil. | 07-11-2013 |
| 20130183139 | ENERGY CONVERTER - 1. The invention relates to an energy converter having a supply channel for a medium and a turbine wheel ( | 07-18-2013 |
| 20130189079 | ROTOR WITH INLET PERIMETERS - A device for use in a molten metal pump helps alleviate jams between rotating rotor blades and a stationary pump base. The device includes inlet perimeters that partially define one or more openings, and one or more rotor blades, wherein each rotor blade has a portion that directs molten metal into a pump chamber, and a portion that directs molten metal outwards. Each rotor blade may also include a recess that makes an opening larger to enable more molten metal to pass through the openings. | 07-25-2013 |
| 20130202408 | GAS TURBINE ENGINE WITH IMPROVED COOLING BETWEEN TURBINE ROTOR DISK ELEMENTS - A gas turbine engine is provided comprising a forward rotor disk and blade assembly capable of rotating; an aft rotor disk and blade assembly capable of rotating; and a row of vanes positioned between the forward rotor disk and blade assembly and the aft rotor disk and blade assembly. The vane row and the forward rotor disk and blade assembly may define a forward cavity. The vane row may comprise at least one stator vane comprising: a main body and an inner shroud structure comprising a cover. The cover may include a first inner cavity receiving cooling air. The cover may further include at least one cooling flow passage. Cooling air flowing from the cooling flow passage has a tangential velocity component. | 08-08-2013 |
| 20130202409 | TURBINE VANE HOLLOW INNER RAIL - A guide vane device for a turbine has an inner platform with a through hole forming a fluid channel for a cooling fluid, wherein the inner platform extends in a circumferential direction around a shaft of the turbine. The guide vane device further includes a hollow aerofoil with a cooling opening for exchanging the cooling fluid passing the through hole into or from the hollow aerofoil, wherein the hollow aerofoil is fixed to a first surface of the inner platform, and a rail with a recess with a cooling fluid passage forming a passage for the cooling fluid to the through hole, wherein the rail is fixed to a second surface of the inner platform and the rail extends along the second surface in the circumferential direction around the shaft. The cooling fluid passage has in the circumferential direction at least the dimension of the through hole. | 08-08-2013 |
| 20130209227 | GAS TURBINE ENGINE COMPONENT WITH DIFFUSIVE COOLING HOLE - A component for a gas turbine engine includes a gas path wall having a first surface, a second surface exposed to hot gas flow, and a cooling hole extending through the gas path wall. The cooling hole includes an inlet formed in the first surface, an outlet formed in the second surface, cooling hole surfaces that define the cooling hole between the inlet and the outlet, and a longitudinal ridge formed along at least one of the cooling hole surfaces. The longitudinal ridge separates the cooling hole into first and second lobes. The cooling hole diverges through the gas path wall, such that cross-sectional area of the cooling hole increases continuously from the inlet through the cooling hole to the outlet. | 08-15-2013 |
| 20130209228 | GAS TURBINE ENGINE COMPONENT WITH CUSPED 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 that extends from an inlet in the first surface of the wall to a transition, a diffusing section that extends from the transition to an outlet in the second surface of the wall, and a cusp on the transition. | 08-15-2013 |
| 20130209229 | GAS TURBINE ENGINE COMPONENT WITH CONVERGING/DIVERGING COOLING PASSAGE - A component for a gas turbine engine includes a gas path wall having a first surface and a second surface and a cooling hole extending through the gas path wall from the first surface to the second surface. The cooling hole includes an inlet portion having an inlet at the first surface, an outlet portion having an outlet at the second surface, and a transition defined between the inlet and the outlet. The inlet portion converges in a first direction from the inlet to the transition and diverges in a second direction from the inlet to the transition. The outlet portion diverges at least in one of the first and second directions from the transition to the outlet. | 08-15-2013 |
| 20130209230 | COOLED VANE OF A TURBINE AND CORRESPONDING TURBINE - A vane is provided for use in a fluid flow of a turbine engine. The vane includes a thin-walled radially extending aerodynamic vane body having axially spaced leading and trailing edges, and a radially outer platform. The wall of the vane body includes an outer shell and an inner shell and defines an interior cavity therein for flowing a cooling medium. A radially extending load strut is arranged at the inner shell of the wall of the leading edge of the vane body. | 08-15-2013 |
| 20130209231 | NOZZLE GUIDE VANE WITH COOLED PLATFORM FOR A GAS TURBINE - A platform for supporting a nozzle guide vane for a gas turbine is provided. The platform has a gas passage surface arranged to be in contact with a streaming operation gas, and a cooling channel for guiding a cooling fluid within the cooling channel formed in an inside of the platform. A cooling portion of an inner surface of the cooling channel is in thermal contact with the gas passage surface. The platform is an integrally formed part representing a segment in a circumferential direction of the gas turbine. The cooling channel has a first cooling channel portion and a second cooling channel portion arranged downstream of the first cooling channel portion with respect to a streaming direction of the operation gas. The first cooling channel portion and the second cooling channel portion are interconnected. | 08-15-2013 |
| 20130216355 | WATER PUMP IN VEHICLE - The present invention relates to a water pump in a vehicle. | 08-22-2013 |
| 20130223986 | GAS TURBINE ENGINE WITH FAN-TIED INDUCER SECTION AND MULTIPLE LOW PRESSURE TURBINE SECTIONS - A gas turbine engine includes a first shaft defining an axis of rotation and a second shaft rotatable about the axis of rotation and spaced radially outwardly relative to the first shaft. A speed change mechanism is driven by the second shaft. A fan includes a fan rotor driven by the speed change mechanism such that the fan and the first shaft rotate at a slower speed than the second shaft. At least one inducer stage is positioned aft of the fan and is coupled for rotation with the fan rotor. | 08-29-2013 |
| 20130223987 | Turbine Nozzle Insert - A turbine nozzle insert of a gas turbine engine is disclosed. The insert may comprise an elongated hollow body portion, a flange portion formed at a first end of the elongated body portion, and a contact portion formed at a second end of the elongated body portion opposite the first end. | 08-29-2013 |
| 20130251508 | DUAL-USE OF COOLING AIR FOR TURBINE VANE AND METHOD - A turbine vane of a gas turbine engine is provided with a hollow core in the leading edge of the outer platform thereof. The core is interconnected with the leading edge core of the airfoil whereby to create a cooling air stream having a dual purpose and cooling both the leading edge of the outer platform and of the airfoil and thereby reducing cooling air consumption. The cooling air enters the core of the outer platform through an inlet port and exits through cooling holes provided in the leading edge of the airfoil. | 09-26-2013 |
| 20130251509 | System and Method for Cooling Gas Turbine Components - The present application provides a cooling system for a gas turbine. The cooling system may include a source of CO | 09-26-2013 |
| 20130259645 | 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 in downstream serial cooling flow relationship, a curved inlet, a constant area and constant width metering section, and a spanwise diverging section leading into the trailing edge cooling slot, and a spanwise height substantially greater than a hole width through the cooling hole. A pressure sidewall surface of the pressure sidewall may be planar through the metering and diverging sections. The width may be constant through the metering and diverging sections. A raised floor may include a flat up ramp in the diverging section, a flat down ramp in the slot. | 10-03-2013 |
| 20130280040 | COOLING ASSEMBLY FOR A GAS TURBINE SYSTEM - A cooling assembly for a gas turbine system includes a turbine nozzle having at least one channel comprising a channel inlet configured to receive a cooling flow from a cooling source, wherein the at least one channel directs the cooling flow through the turbine nozzle in a radial direction at a first pressure to a channel outlet. Also included is an exit cavity for fluidly connecting the channel outlet to a region of a turbine component, wherein the region of the turbine component is at a second pressure, wherein the first pressure is greater than the second pressure. | 10-24-2013 |
| 20130294889 | GAS TURBINE ENGINE COMPONENTS WITH FILM COOLING HOLES HAVING CYLINDRICAL TO MULTI-LOBE CONFIGURATIONS - An engine component includes a body; and a plurality of cooling holes formed in the body, at least one of the cooling holes having a multi-lobed shape with at least a first lobe, a second lobe, and a third lobe. | 11-07-2013 |
| 20130302141 | Convective Shielding Cooling Hole Pattern - The cooling system for a turbine may include a plurality of platform cooling openings positioned in a platform of the turbine airfoil. In particular, the first set of cooling openings may create a first cooling path and a second set of cooling openings may be placed in the path of the first cooling path where the first cooling flow will cool the second set of cooling openings. | 11-14-2013 |
| 20130302142 | ELECTRIC FLUID PUMP - An electric water pump includes a housing, a propelling mechanism disposed in the housing, and a driving mechanism for driving a driving shaft of the propelling mechanism. The driving shaft has a fluid passage in fluid communication with an impeller chamber and a cooling chamber in the housing. The propelling mechanism further includes a spiral rod disposed in the fluid passage and co-rotatable with the driving shaft, and an impeller disposed in the impeller chamber. During operation of the propelling mechanism, the impeller drives flow of a fluid, and the spiral rod co-rotates with the driving shaft to change the rate of the fluid flowing through the driving shaft. | 11-14-2013 |
| 20130323019 | APPARATUS FOR MINIMIZING SOLID PARTICLE EROSION IN STEAM TURBINES - Solid particle erosion in a steam turbine is minimized by diverting through angled slots formed in appendages of outer rings of the diaphragms, a portion of the steam from the steam flow path thereby bypassing downstream rotating components. The slot through the first stage appendage lies in communication with a passage through a downstream outer ring of a following stage such that the diverted solid particle containing steam may be extracted from the steam flow path and passed to the feed water heater of the turbine. The slot in the second stage appendage diverts steam from between the first and second stages and about the second stage. Solid particle erosion in various regions, i.e., the trailing edge of the stator vanes, along the surfaces of the buckets and in the regions of the cover and its connection with the buckets as well as the sealing devices is thereby minimized. | 12-05-2013 |
| 20130336767 | COOLING FOR A TURBINE AIRFOIL TRAILING EDGE - An assembly for a gas turbine engine includes a first platform and an airfoil extending from the first platform. The airfoil includes a first fillet, pressure side biased discharge openings, and a first center cooling discharge opening. A pressure side wall of the airfoil and the first platform form an acute angle at the trailing edge. The first fillet is formed around a perimeter of the airfoil where the airfoil extends from the first platform. The pressure side biased cooling discharge openings are along the trailing edge outside of the first fillet. Each pressure side biased cooling discharge opening extends from the trailing edge along the pressure side wall. The first center cooling discharge opening extends along the trailing edge into the first fillet and is centrally located between the pressure side wall and the suction side wall. | 12-19-2013 |
| 20130343872 | COOLED COMPONENT FOR THE TURBINE OF A GAS TURBINE ENGINE - A component for the turbine of a gas turbine engine is provided. The component two facing walls interconnected by one or more generally elongate divider members to partially define side-by-side, generally elongate, cooling fluid passage portions which form a multi-pass cooling passage within the component. The passage portions are connected in series fluid flow relationship by respective bends formed by joined ends of neighbouring of the passage portions. The component further includes one or more core tie linking passages formed in the divider members. One or more differential pressure reducing arrangements are formed in the multi-pass cooling passage adjacent respective of the core tie linking passages. | 12-26-2013 |
| 20130343873 | TURBINE ENGINE VARIABLE AREA VANE - A turbine engine stator vane is provided that rotates about an axis, and includes an airfoil, a flange and a shaft. The airfoil extends axially between a first airfoil end and a second airfoil end. The airfoil includes a concave side surface, a convex side surface and a cavity. The concave and the convex side surfaces extend between an airfoil leading edge and an airfoil trailing edge. The cavity extends axially into the airfoil from a cavity inlet in an end surface at the second airfoil end. The flange is connected to the second airfoil end. The flange extends circumferentially around at least a portion of the cavity inlet, and radially away from the concave and the convex side surfaces to a distal flange edge. The shaft extends along the axis, and is connected to the second airfoil end. | 12-26-2013 |
| 20140017066 | DEVICE FOR SEPARATING WATER DROPLETS FROM A GAS OR VAPOR FLOW - A moisture separator for a steam turbine power plant for separating moisture from a flow of vapor or gas that includes a bundle of vanes having a corrugated portion and a trailing edge. The trailing edge includes means to collect moisture on the surfaces of the trailing edge. These means can include a clip-like attachment, or a U-shape of the trailing edge itself. The collection means on the trailing edge contribute to the overall moisture separation efficiency of the moisture separator and a decrease of risk of damage to the steam turbine driven by the steam flow. | 01-16-2014 |
| 20140023483 | AIRFOIL ASSEMBLY INCLUDING VORTEX REDUCING AT AN AIRFOIL LEADING EDGE - An airfoil assembly including an endwall and an airfoil extending from the into a gas flow path. The endwall includes upstream and downstream edges, and is defined on a platform structure having a front surface extending radially in a direction of a thickness of the platform structure. At least one fluid injection passage extends through the platform structure in a direction from the upstream edge toward the downstream edge of the endwall. The fluid injection passage has an outlet opening defined at the endwall and an inlet opening in fluid communication with a pressurized fluid source. The fluid injection passage extends at a shallow angle relative to a plane of the endwall wherein the fluid injection passage defines a passage axis passing through the front surface and the endwall for effecting energization of a boundary layer between the outlet opening and the airfoil leading edge. | 01-23-2014 |
| 20140030064 | TURBINE ENGINE COMBUSTOR AND STATOR VANE ASSEMBLY - A turbine engine assembly includes a combustor and a stator vane arrangement having a plurality of stator vanes. The combustor includes a combustor wall that extends axially from a combustor bulkhead to a distal combustor wall end, which is located adjacent to the stator vane arrangement. The combustor wall includes a support shell with a plurality of impingement apertures, and a heat shield with a plurality of effusion apertures. The combustor wall end includes a plurality of circumferentially extending film cooled regions. At least one of the film cooled regions is circumferentially aligned with one of the stator vanes and includes a cooling aperture. | 01-30-2014 |
| 20140030065 | Steam Turbine, and Steam Turbine Stationary Blade - The present invention is a steam turbine comprising a turbine stage having a stationary blade and a moving blade provided on the downstream side of the stationary blade in a working fluid flow direction, wherein the stationary blade is formed in a hollow blade shape by deformation processing a metal plate, and wherein a slit to guide liquid droplets deposited on a blade wall surface to the inside of the blade is formed in the blade wall by overlaying an airfoil suction-side metal plate and an airfoil pressure-side metal plate with a gap therebetween in a blade tail part of the stationary blade. | 01-30-2014 |
| 20140037429 | TURBINE VANE - A plurality of film cooling holes are formed, so as to communicate with a front cooling passage, in a vane surface on the front-edge side of a stator vane body of a turbine stator vane. The hole cross-section of each of the film cooling holes has a rectangular long-hole shape extending in a direction parallel to the cross-section along the span direction and having a rounded corner. The hole-center line of each of the film cooling holes is inclined with respect to the thickness direction in the cross-section along the span direction. The exit-side portion of the hole wall surface of each of the film cooling holes is inclined with respect to the thickness direction by a greater degree than that of the hole-center line. | 02-06-2014 |
| 20140056690 | GAS TURBINE - A gas turbine includes a turbine blade, a turbine vane, a ring segment circumferentially surrounding the turbine blade, an outer shroud circumferentially surrounding the turbine vane, and a combustion gas flow-path provided in the ring segment and the outer shroud. The outer shroud is positioned on an upstream side of the ring segment in a gas flow direction of the combustion gas. Seal gas, of which temperature is lower than that of the combustion gas, is fed between the ring segment- and the outer shroud into the combustion gas flow-path. The outer shroud has a guide surface that is provided on an inner circumference thereof on a downstream side of the gas flow direction. The guide surface is formed such that a flow passage area of the combustion gas flow-path is gradually increased. | 02-27-2014 |
| 20140056691 | IMPULSE TURBINE FOR USE IN BI-DIRECTIONAL FLOWS - A turbine arrangement for a bi-directional reversing flow is provided. The turbine arrangement may include a rotor rotatably mounted to rotate about an axis of the turbine arrangement, and the rotor may have a plurality of rotor blades disposed circumferentially thereabout. A first set of guide vanes may be circumferentially disposed about the axis for directing the bi-directional reversing flow to and from the rotor blades via a first flow passaged defined by a first duct. A second set of guide vanes may be axially spaced from the first set of guide vanes and circumferentially disposed about the axis for directing the bi-directional reversing flow to and from the rotor blades via a second flow passage defined by a second duct. The guide vanes may be disposed at a greater radius than the rotor blades, such that the guide vanes are radially offset from the rotor blades. | 02-27-2014 |
| 20140079536 | FAN MODULE - A fan module includes a module housing, a first set of stationary blades, a second set of stationary blades, a motor, and a dynamic blade combination. The module housing has a channel having an inlet. The first and second sets of stationary blades are disposed at the inner wall of the module housing. The motor is disposed in the channel and includes a rotor. The dynamic blade combination includes a hub fixed to a periphery of the rotor, and a first set of dynamic blades and a second set of dynamic blades that surround and are disposed to a periphery of the hub. The first set of dynamic blades is located between the inlet and the first set of stationary blades. The second set of dynamic blades is located between the first and second set of stationary blades. | 03-20-2014 |
| 20140099189 | GAS TURBINE ENGINE COMPONENTS WITH LATERAL AND FORWARD SWEEP FILM COOLING HOLES - An engine component includes a body having an internal surface and an external surface, the internal surface at least partially defining an internal cooling circuit. The engine component further includes a plurality of cooling holes formed in the body and extending between the internal cooling circuit and the external surface of the body. The plurality of cooling holes includes a first cooling hole with forward diffusion and lateral diffusion. | 04-10-2014 |
| 20140105725 | FISH MOUTH SEAL CARRIER - A fish mouth seal carrier for a guide vane arrangement of a gas turbine comprises a first half-shell element and a second half-shell element bonded to it, which together form a box profile with two axial arms and two radial arms. A sealing element is arranged on one of the axial arms of the box profile. At least one of the two half-shell elements has an integrally formed axial flange for forming, with a guide vane platform, a fish mouth seal accommodating an axial flange of an adjoining moving blade radially between the guide vane platform and the axial flange of the seal carrier. | 04-17-2014 |
| 20140105726 | TURBINE AIRFOIL VANE WITH AN IMPINGEMENT INSERT HAVING A PLURALITY OF IMPINGEMENT NOZZLES - A turbine vane includes a generally elongated hollow airfoil and a cooling system. The cooling system is positioned within the airfoil and includes a cooling chamber and an impingement insert positioned in the cooling chamber. The impingement insert and an inner surface of an outer wall of the airfoil define a cooling channel therebetween. The impingement insert includes a plurality of impingement nozzles extending toward the inner surface of the outer wall and a plurality of impingement orifices. At least one of the impingement orifices is arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of the at least one impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent impingement orifice. | 04-17-2014 |
| 20140112758 | High Temperature Components With Thermal Barrier Coatings for Gas Turbine - The most principal feature of the present invention is as follows: Namely, in the gas-turbine-use high-temperature component including the thermal barrier coating and a cooling structure, micro passages are provided inside an alloy bond-coat layer and a thermal-barrier ceramic top-coat layer of the thermal barrier coating, the micro passages being in communication from the substrate side to the surface side. Moreover, a partial amount of coolant of a coolant for cooling the high-temperature component is caused to flow out to the outside of the high-temperature component via these micro passages. The employment of the structure like this makes it possible to expect the implementation of a high-temperature component's heat-resistant-temperature enhancement effect based on the transpiration cooling effect. | 04-24-2014 |
| 20140119888 | IMPINGEMENT COOLING OF TURBINE BLADES OR VANES - A turbine assembly includes a basically hollow aerofoil. A wall segment may be arranged at a side of the aerofoil. An insertion aperture in the wall segment provides access to the aerofoil and an impingement tube may be inserted via the insertion aperture into the aerofoil to be located within the aerofoil and extend at least in a span wise direction of the aerofoil. A protrusion section of the impingement tube may extend in a direction basically perpendicular to the span wise direction over an edge of the insertion aperture. The protrusion section may be overlapped by at least a part of the wall segment. Adjacent to the protrusion section, an overlap section of the impingement tube is arranged to abut the edge of the insertion aperture. The protrusion section and the overlap section may be formed integrally with each other in one piece. | 05-01-2014 |
| 20140212270 | GAS TURBINE ENGINE COMPONENT HAVING SUCTION SIDE CUTBACK OPENING - An airfoil for a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a pressure side wall and a suction side wall spaced apart from the pressure side wall and each extending between a leading edge portion and a trailing edge portion. A plurality of cutback openings are spaced along a radial axis of the suction side wall. | 07-31-2014 |
| 20140219778 | GAS TURBINE ENGINE AIRFOIL WITH VANE PLATFORM COOLING PASSAGE - A stator vane for a gas turbine engine includes an airfoil extending in a radial direction and supported by a platform having a gas flowpath surface. A cooling passage is arranged in the platform and includes a circumferential passage that is fluidly connected to an inlet passage extending through and edge of the platform, and film cooling holes extending from the gas flowpath surface to the circumferential passage, radial extending passage through the edge of the platform. A void is interconnected to at least one of the radially extending passage and the inlet passage. | 08-07-2014 |
| 20140241859 | COMPRESSOR CIRCUIT FOR A PNEUMATIC CONTROL DEVICE OF A MOTOR VEHICLE - The invention relates to a compressor circuit for a pneumatic control device of a motor vehicle, including: a compressor, and a pressure regulator, the pressure regulator being arranged between an input side of the compressor and the pneumatic control device, wherein the pressure regulator is designed to measure an upstream pressure of air from the pneumatic control device present on its side facing the pneumatic control device and to supply this air at an input pressure to the input side of the compressor, the pressure regulator being designed to control the input pressure on the basis of said measurement. | 08-28-2014 |
| 20140248129 | LPC FLOWPATH SHAPE WITH GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION - A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flowpath. A shaft provides a rotational axis. A hub is operatively supported by the shaft. A rotor is connected to the hub and supports a compressor section. The compressor section is arranged in a core flow path axially between the inlet case flow path and the intermediate case flow path. The core flowpath has an inner diameter and an outer diameter. At least a portion of inner diameter has an increasing slope angle relative to the rotational axis. A bearing is mounted to the hub and supports the shaft relative to one of the intermediate case and the inlet case. | 09-04-2014 |
| 20140255158 | COATING PROCESS FOR GAS TURBINE ENGINE COMPONENT WITH COOLING HOLES - A method of coating a component having a multiple of cooling holes includes removing at least a portion of a prior coating; directing a gas through at least one of the multiple of cooling holes; and applying a coat layer while directing the gas through at least one of the | 09-11-2014 |
| 20140255159 | INTEGRATED STRUT-VANE - An integrated strut and turbine vane nozzle (ISV) has inner and outer annular duct walls defining an annular flow passage therebetween. Circumferentially spaced-apart struts extend radially across the flow passage. Circumferentially spaced-apart vanes also extend radially across the flow passage and define a plurality of inter-vane passages. Each of the struts is integrated to an associated one of the vanes to form therewith an integrated strut-vane airfoil. The inter-vane passages on either side of the integrated strut-vane airfoil may be adjusted for aerodynamic considerations. The vanes may be made separately from the struts and manufactured such as to cater for potential misalignments between the parts. | 09-11-2014 |
| 20140255160 | Hydro Turbine - A hydro turbine includes at least two annular wheel discs, a center shaft disposed at the centers of the annular wheel discs, spokes disposed between the annular wheel discs and the center shaft, and a plurality of blades annularly disposed at the annular wheel discs. The number of the blades is at least 28 and at most the integer number of an outer circumference measured in centimeters of one annular wheel disc. The sum of arc length of one side of all the blades is 0.85-2 times the outer circumference. A blade angle formed between two line segments connecting the middle point to the two endpoints of an arc of each blade is in a range of 100°-170°. A blade installation angle formed between the chord line of the arc and a radius line of each annular wheel disc is in a range of 15°-75°. | 09-11-2014 |
| 20140271129 | Cast-in cooling features especially for turbine airfoils - A method is provided for making a mold for casting advanced turbine airfoils (e.g. gas turbine blade and vane castings) which can include complex internal and external air cooling features to improve efficiency of airfoil cooling during operation in the gas turbine hot gas stream. The method steps involve incorporating at least one fugitive insert in a ceramic material in a manner to form a core and at least a portion of an integral, cooperating mold wall wherein the core defines an internal cooling feature to be imparted to the cast airfoil and the at least portion of the mold wall has an inner surface that defines an external cooling feature to be imparted to the cast airfoil, selectively removing the fugitive insert, and incorporating the core and the at least portion of the integral, cooperating mold wall in a mold for receiving molten metal or alloy cast in the mold. | 09-18-2014 |
| 20140286749 | MULTI-STAGE AXIAL COMPRESSOR WITH COUNTER-ROTATION USING ACCESSORY DRIVE - A multi-stage axial compressor for counter rotation. A first series of rotor blade assemblies are mounted on and rotate with the driveshaft, each rotor blade assembly of the first series comprising a rotating stage of the multi-stage axial compressor. A second series of rotor blade assemblies provide a counter-rotating stage of the multi-stage axial compressor. An accessory drive links the second series of rotor blade assemblies to the driveshaft and causes counter-rotation of the second series of rotor blade assemblies. | 09-25-2014 |
| 20140286750 | VARIABLE TURBINE/COMPRESSOR GEOMETRY - A variable turbine/compressor geometry for an exhaust gas turbocharger of an internal combustion engine includes a blade bearing ring having guide blades rotatably mounted therein and a cover disc which is arranged on the face end of the guide blades. At least one gas guide channel may run through at least one guide blade. The gas guide channel including, adjacent to a profile lug, a least one inlet opening, and on at least one of a face end directed towards the cover disc and a bottom side directed towards the blade bearing ring includes an outlet opening. The gas guide channel may be configured to generate a friction-reducing air cushion during operation on at least one of the face end and the bottom side. | 09-25-2014 |
| 20140294566 | TURBOMACHINE INLET BLEED HEATING ASSEMBLY - In a turbomachine having an inlet, a compressor, and a turbine, a closed loop sends fluid from a stage of the compressor to a heat exchanger in the turbine and to the inlet. The closed loop heats the fluid, cools the turbine, and delivers heated fluid to the inlet. A mixer can be interposed between the heat exchanger and the inlet to mix fluid from the heat exchanger with compressor discharge fluid, delivering the mixed fluid to the inlet. The mixer can control flow received so that desired temperature and/or flow rate can be provided to the inlet. | 10-02-2014 |
| 20140294567 | GUIDE VANE FOR A TURBOMACHINE, GUIDE VANE CASCADE, AND METHOD FOR MANUFACTURING A GUIDE VANE OR A GUIDE VANE CASCADE - A guide vane for a turbomachine axially pivotably coupled to a radially outwardly disposed flow-limiting wall and to a radially inwardly disposed inner ring of the turbomachine; and a trailing edge gap being formed between an upper trailing edge of the guide vane and the flow-limiting wall and/or between a lower trailing edge of the guide vane and the inner ring; the upper trailing edge and/or the lower trailing edge of the guide vane having at least one air outlet opening for an air outflow for forming an air curtain for at least partially sealing the trailing edge gap in the area of the upper trailing edge and/or the lower trailing edge in the area of the lower trailing edge. Also, a guide vane cascade, as well as a method for manufacturing a guide vane or a guide vane cascade. | 10-02-2014 |
| 20140321977 | DURABLE TURBINE VANE - A durable nozzle vane includes a leading edge where hot combustion gas impinges on the vane, a pressure side extending from the leading edge to a trailing edge and a suction side extending from the leading edge to the trailing edge, a blunt region along the suction side between the leading edge and a high curvature region, the blunt region having small curvature in the length between the leading edge and the high curvature region allowing for at least two rows of cooling apertures. | 10-30-2014 |
| 20140321978 | Turbine Nozzle and Shroud - A nozzle and shroud for use in an air cycle machine has a plate and a shroud curving in a first axial direction about a center axis of the shroud relative to the plate. A plurality of vanes extends in a second axial direction away from the plate. The plurality of vanes extends for a height away from the plate and a width defined as the closest distance between two adjacent vanes, with a ratio of the height to the width being between 1.7377 and 2.1612. An air cycle machine and a method of repair are also disclosed. | 10-30-2014 |
| 20140321979 | TURBINE NOZZLE PIECE PARTS WITH HVOC COATINGS - A nozzle for an air cycle machine. The nozzle has a disk section having a central axis. The nozzle also includes a plurality of blades which extend a blade height H from a bladed face of the disk section. The plurality of blades are arranged radially about the disk section. The nozzle has a throat width W defined between each radially adjacent pair of the plurality of turbine blades. The nozzle includes a coating substantially encapsulating the disk section and the plurality of blades, wherein the coating contains more than 91 percent tungsten carbide by volume. | 10-30-2014 |
| 20140321980 | COOLING SYSTEM INCLUDING WAVY COOLING CHAMBER IN A TRAILING EDGE PORTION OF AN AIRFOIL ASSEMBLY - An airfoil in a gas turbine engine includes an outer wall, a cooling fluid cavity, and a cooling system. 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 system receives cooling fluid from the cooling fluid cavity for cooling the trailing edge portion of the outer wall and includes a cooling fluid chamber defined by opposing first and second sidewalls that include respective alternating angled sections that provide the cooling fluid chamber with a zigzag shape. | 10-30-2014 |
| 20140328669 | AIRFOIL WITH COOLING PASSAGES - An airfoil having cooling passages inside is provided, wherein each radial cross section of the airfoil has a shape of specific profile, wherein hot gas flows along the airfoil's surface from a leading edge to trailing edge, the airfoil's surface comprises a pressure-side and suction-side defined by the trailing edge and leading edge, wherein the trailing edge has cooling fluid discharge exits, the pressure-side and suction-side are respectively defined by a wall comprising an inner surface and an outer surface, which inner surface has ribs extending in a rib-direction inclined to the radial direction, wherein along a portion of at least 10% of the profile's lengths the inclined ribs contact each other at respective cross-contact-points, forming a 2-dimensional matrix. At least one additional blocking-rib extends from the pressure-side to the suction-side and extends from one cross-contact-point to another to cause additional turbulence of said cooling fluid flow to be discharged. | 11-06-2014 |
| 20140334914 | COMPONENT FOR A THERMAL MACHINE, IN PARTICULAR A GAS TURBINE - The invention relates to a component for a thermal machine, in particular a gas turbine, which includes a corner and/or edge subjected to a thermally high load. The cooling of the component is improved in a manner such that at least one cooling channel is countersunk into the surface of the component in the immediate vicinity of the corner and/or edge in order to cool the corner and/or edge. | 11-13-2014 |
| 20140348636 | AIRFOIL COOLING CIRCUIT - An airfoil cooling circuit includes an impingement cooling circuit and a serpentine cooling circuit. An airfoil for use in a gas turbine engine having a cooling circuit which includes an impingement cooling circuit and a serpentine cooling circuit | 11-27-2014 |
| 20140377054 | NOZZLE FILM COOLING WITH ALTERNATING COMPOUND ANGLES - A nozzle segment for a nozzle ring of a gas turbine engine is disclosed. The nozzle segment includes a first endwall, a second endwall, and an airfoil extending between the first endwall and the second endwall. The airfoil includes a multiple groups of cooling apertures spaced apart and alternating in directionality such that a first grouping of cooling apertures is angled toward the first endwall, a second grouping of cooling apertures is angled toward the second endwall and spaced apart from the first grouping of cooling apertures, and a third grouping of cooling apertures are angled toward the first endwall and spaced apart from the second grouping of cooling apertures. | 12-25-2014 |
| 20150010385 | GAS TURBINE WITH HIGH-PRESSURE TURBINE COOLING SYSTEM - The present invention relates to a gas turbine with a turbine stator wheel, which is fitted with stator vanes and includes a ring segment-shaped vane root, where the stator vanes are designed hollow and have a vane interior which can be supplied with cooling air, where a ring-shaped sealing element of an inter-stage seal is arranged radially on the inside, relative to an engine axis, on the vane root, where in the vane root at least one outflow duct is provided, characterized in that between the vane root and the sealing element an annular space extending substantially in the axial direction is formed, into which the outflow duct issues and which discharges into the area of the inter-stage seal. | 01-08-2015 |
| 20150016961 | COOLED TURBINE GUIDE VANE OR BLADE FOR A TURBOMACHINE - A turbine airfoil for a turbomachine is provided. The airfoil includes a suction side wall and a pressure side wall bordering an airfoil cavity, which receives a cooling fluid for cooling the airfoil. The suction side wall includes one or more protrusions extending inside the cavity. The number of protrusions on the suction side may be higher than the number of protrusions on the pressure side. The density of protrusions on the suction side may be higher than the density of protrusions on the pressure side and/or the surface of protrusions on the suction side may be larger than the surface of protrusions on the pressure side, so that heat transfer from the suction side to the cooling fluid is higher compared to heat transfer from the pressure side to the cooling fluid during operation of the turbomachine. | 01-15-2015 |
| 20150030432 | TRAILING EDGE COOLING ARRANGEMENT FOR AN AIRFOIL OF A GAS TURBINE ENGINE - An airfoil ( | 01-29-2015 |
| 20150044028 | LOW PRESSURE RATIO FAN ENGINE HAVING A DIMENSIONAL RELATIONSHIP BETWEEN INLET AND FAN SIZE - According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan that has a plurality of fan blades. A diameter of the fan has a dimension D that is based on a dimension of the fan blades. Each fan blade has a leading edge. An inlet portion is situated forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and about 0.45. | 02-12-2015 |
| 20150044029 | AEROFOIL - An aerofoil component of a gas turbine engine has an aerofoil portion which spans, in use, a working gas annulus of the engine. The aerofoil portion has a pressure side outer wall and a suction side outer wall, each extending from the leading edge to the trailing edge of the aerofoil portion. The aerofoil portion further has one or more main passages which extend in the annulus-spanning direction of the aerofoil portion and which receive, in use, a flow of coolant. The aerofoil portion further has one or more suction wall passages which extend in the annulus-spanning direction of the aerofoil portion and which receive, in use, a flow of coolant, each suction wall passage being bounded on opposing first sides by the suction side outer wall and an inner wall of the aerofoil portion, the inner wall separating the suction wall passages from the main passages. | 02-12-2015 |
| 20150063982 | MULTI-STAGE INFLOW TURBINE PUMP FOR PARTICLE COUNTERS - An airborne, gas particle counter, includes an inflow multiple stage turbine pump inducing flow through a particle counter. The turbine pump includes a rotor and stator assemblies and establishes gas flow through a second housing, which defines a view volume where particles are counted by light scattering or obscuration of a beam intersecting the light flow path. Intersecting rotor and stator assemblies limit particles to those suspended in the gas flow. The pump assembly includes integral flow paths limiting the need for external tubing. | 03-05-2015 |
| 20150110601 | COOLANT BRIDGING LINE FOR A GAS TURBINE, WHICH COOLANT BRIDGING LINE CAN BE INSERTED INTO A HOLLOW, COOLED TURBINE BLADE - A coolant bridging line for a gas turbine having inner and outer sides which are separated by a wall, wherein the coolant bridging line extends from a first component of the gas turbine to a second component of the gas turbine is provided herein. The coolant bridging line for a gas turbine provides a further increase in the service life of the components of the gas turbine possible with a degree of efficiency which is nevertheless as high as possible. To this end, the coolant bridging line has means which change the heat transfer between and/or the flow conditions on the inner and outer sides. | 04-23-2015 |
| 20150139781 | REACTION TURBINE - A reaction turbine, according to the present invention, includes first and second rotor plates, which are coupled together to form an integrated rotor, and an inner flow path including a combination of first and second flow paths, which are formed on the surfaces of the first and second rotor plates that face each other, respectively, thereby enabling easier manufacturing into a form desired by a designer by eliminating the limitation of a cross-sectional shape of the inner flow path. In addition, a cross section of each of the first and second flow paths can be formed into a semicircular shape thus yielding a circular shape for the inner flow path, which is formed by combining the first and second flow paths, thereby effectively enhancing the performance of a turbine by minimizing pressure loss of a working fluid that passes through the inner flow path. | 05-21-2015 |
| 20150147158 | COOLED AIRFOIL TRAILING EDGE AND METHOD OF COOLING THE AIRFOIL TRAILING EDGE - An airfoil and method of cooling a airfoil including a leading edge, a trailing edge, a suction side, a pressure side and at least one internal cooling channel configured to convey a cooling fluid, is provided. A plurality of trailing edge bleed slots are in fluid communication with the at least one internal cooling channel, wherein a downstream edge of the pressure side of the airfoil lies upstream of a downstream edge of the suction side to expose the plurality of trailing edge bleed slots proximate to the trailing edge of the airfoil. The at least one internal cooling channel is configured to supply the cooling fluid from a source of cooling fluid towards the plurality of trailing edge bleed slots. A plurality of obstruction features are disposed within the at least one internal cooling channel and at a downstream edge of the remaining pressure side. The one or more obstruction features are configured having a predefined substantially polygon shape, to distribute a flow of the cooling fluid and provide distributed cooling to the plurality of trailing edge bleed slots. | 05-28-2015 |
| 20150292353 | HIGH PRESSURE COMPRESSOR THERMAL SHIELD APPARATUS AND SYSTEM - In various embodiments, a high pressure compressor may comprise a thermal shield. The thermal shield may be installed between a first rotor and a second rotor. The thermal shield may also be installed radially inward of a stator. The stator may be a shrouded stator. Moreover, the thermal shield may be configured to thermally isolate and/or reduce the thermal load associated with windage on a rotor hub. | 10-15-2015 |
| 20150292355 | TURBOMACHINE COOLING SYSTEMS - Embodiments of a turbomachine, having a longitudinal axis and a flowpath are provided. The turbomachine includes an impeller circumferentially disposed around the longitudinal axis, and an impeller shroud that surrounds a portion of the impeller. At least one opening formed through the impeller shroud surface provides fluid communication between the flowpath and a dead-headed plenum. | 10-15-2015 |
| 20150300697 | TURBOMACHINE AND REFRIGERATION CYCLE APPARATUS - A turbomachine for a refrigeration cycle apparatus that uses a refrigerant whose saturated vapor pressure is a negative pressure at ordinary temperature includes a rotation shaft and a bearing for supporting the rotation shaft. The rotation shaft includes a lubricant supply passage for supplying the refrigerant as a lubricant that travels smoothly between the rotation shaft and the bearing. The lubricant supply passage includes a main passage and a sub-passage. The main passage extends from an inlet located at an end of the rotation shaft in an axial direction of the rotation shaft. The sub-passage diverges from the main passage and extends to an outlet located in a side surface of the rotation shaft. | 10-22-2015 |
| 20150322801 | AIRFOIL LEADING EDGE FILM ARRAY - An airfoil according to an exemplary aspect of the present disclosure includes, among other things, a first cooling hole with a first cooling passage arranged at a first angle relative to a chordwise axis and a second cooling hole with a second cooling passage arranged at a second different angle relative to the chordwise axis. A radial projection of the first cooling passage intersects a radial projection of the second cooling passage. | 11-12-2015 |
| 20150322964 | Diffuser - A diffuser includes a body and diffusion blades disposed at a circumferential edge of the body. The body includes an air inlet side and an air outlet side. Each diffusion blade has an inclined guide surface. The guide surface has a starting end at the air inlet side of the body and a terminating end adjacent the air outlet side of the body. The guide surface extends from the starting end to the terminating end along a circumferential path and has a width gradually increasing from the starting ending to the terminating end. The diffuser further includes air outlets at the terminating ends of the guide surfaces. Each air outlet fluidly connects the terminating end of a corresponding one of the guide surfaces and the air outlet side of the body. A blower employing the diffuser is also provided. | 11-12-2015 |
| 20150330238 | DISTRIBUTOR DEVICE FOR COOLING AIR WITHIN AN ENGINE - A distributor device for distributing cooling air within a gas turbine engine, the device including a base mountable adjacent an inlet for air to be distributed; and a deflector supported by the base and in fluid communication with the air inlet, the deflector being configured to direct air from the air inlet in a plurality of directions within the engine; wherein the deflector includes one or more deflecting surfaces curvilinearly configured to direct the air differentially in a plurality of desired directions within the engine. In embodiments the deflecting surface(s) is/are curved and are so configured such that air from the air inlet is directable either: (i) in directions of travel in each of a plurality of desired different, non-parallel directions within the engine; or (ii) in each of a plurality of desired different directions within the engine with different flow rates. | 11-19-2015 |
| 20150345297 | GAS TURBINE BLADE - The invention refers to a gas turbine blade including an airfoil extending in radial direction from a blade root to a blade tip, defining a span ranging from 0% at the blade root to 100% at the blade tip, and extending in axial direction from a leading edge to a trailing edge, which limit a chord with an axial chord length defined by an axial length of a straight line connecting the leading edge and trailing edge of the airfoil depending on the span. The axial chord length increases at least from 80% span to 100% span. | 12-03-2015 |
| 20150345322 | VANE SUPPORT SYSTEMS - A vane support system includes a frame and a vane. The frame has a first end configured to engage to a first platform and a second end configured to engage a second platform, so the frame can structurally support at least one of the first platform and the second platform. The first and second ends define a vane axis therebetween. The vane is mounted to the frame about the vane axis. A gas turbine engine includes a case defining a centerline axis of the engine, an inner housing and a plurality of variable vanes. The inner housing is radially inward of the case with respect to the centerline axis. At least one of the variable vanes structurally supports the case and the inner housing in response to at least one of radial, axial or tangential loads with respect to the centerline axis. | 12-03-2015 |
| 20150345327 | Heated Inlet Guide Vane - A heated inlet guide vane is disclosed. The disclosed guide vane may include a body having an inlet cavity disposed alongside of an outlet cavity wherein the outlet cavity is disposed alongside the leading edge of the vane. The inlet cavity is in communication with a source of heated air. The inlet cavity is in communication with the outlet cavity by way of a plurality of impingement holes spaced along the inner wall disposed between the two cavities. Bleed holes are spaced along the length of the outlet cavity. Fresh, heated air enters the outlet cavity along the entire length of the outlet cavity and quickly exits the outlet cavity with minimal cooling so that a uniform d-icing capability is provided along the entire leading edge of the inlet guide vane. | 12-03-2015 |
| 20150345515 | TURBOCHARGER - A turbocharger includes a compressor housing and a bearing housing. The compressor housing includes a diffuser surface, and the bearing housing includes an opposite surface. An adhesion preventing part is disposed on each of the diffuser surface and the opposite surface. The adhesion preventing part is provided with a surface forming part and a tank part and is configured so that air is ejected through the ejection holes of the surface forming part to a diffuser passage. An air supply passage is formed in the compressor housing and the bearing housing. At least one of the compressor housing and the bearing housing includes a depurant injection port for supplying a depurant having compatibility with deposits to the tank part through the air supply passage. | 12-03-2015 |
| 20150361799 | METHOD FOR CREATING A FILM COOLED ARTICLE FOR A GAS TURBINE ENGINE - A method for finishing a film cooled article includes providing a film cooled article including at least one inner cooling plenum and at least one opening connecting the inner cooling plenum to an exterior surface of the film cooled article, positioning a machining element in contact with the exterior surface of the film cooled article, automatically moving the machining element along the exterior surface while maintaining contact between the machining tool and the surface, identifying an actual position of at least one film opening based on sensory feedback from the machining element using a controller, removing material from the exterior surface at the at least one film opening using the machining element, thereby creating a depression at the at least one film opening. | 12-17-2015 |
| 20150369060 | TAPERED THERMAL BARRIER COATING ON CONVEX AND CONCAVE TRAILING EDGE SURFACES - A turbine engine component has an airfoil portion having a pressure side, a suction side, and a trailing edge. The trailing edge has a center discharge cooling circuit, which center discharge cooling circuit has an exit defined by a concave surface on the pressure side of the airfoil portion and a convex surface on the suction side of the airfoil portion. The airfoil portion has a thermal barrier coating on the pressure side and the suction side. The thermal barrier coating on the convex surface tapers to zero in thickness at a point spaced from the trailing edge so as to leave an uncoated portion on the convex surface. | 12-24-2015 |
| 20150369072 | SPLIT AIRFOIL CLUSTER AND METHOD THEREFOR - A method of fabricating airfoil cluster includes providing an airfoil cluster that has a pair of spaced-apart airfoils that extend from a common platform wall. The airfoil cluster is then divided through the common platform wall to provide separate first and second airfoil segments. At least one cooling hole is then formed in at least one of the first and second airfoil segments. The segments are then metallurgically fused together in a distinct metallurgical joint in the common platform wall. | 12-24-2015 |
| 20150369073 | Flow Control Structures For Turbomachines and Methods of Designing The Same - Flow control devices and structures designed and configured to improve the performance of a turbomachine. Exemplary flow control devices may include various flow guiding channels, ribs, diffuser passage-width reductions, and other treatments and may be located on one or both of a shroud and hub side of a machine to redirect, guide, or otherwise influence portions of a turbomachine flow field to thereby improve the performance of the machine. | 12-24-2015 |
| 20160003047 | RADIAL TURBINE - A radial turbine comprising: a housing ( | 01-07-2016 |
| 20160003071 | GAS TURBINE ENGINE STATOR VANE BAFFLE ARRANGEMENT - A stator vane for a gas turbine engine includes an airfoil that has an exterior wall that provides a cooling cavity. The exterior surface has an interior surface that has multiple pin fins extending therefrom. A baffle is arranged in the cooling cavity and is supported by the pin fins. | 01-07-2016 |
| 20160010464 | COMPONENTS WITH COOLING CHANNELS AND METHODS OF MANUFACTURE | 01-14-2016 |
| 20160010465 | GAS TURBINE ENGINE AIRFOIL LEADING EDGE COOLING | 01-14-2016 |
| 20160017736 | COOLED COMPOSITE SHEETS FOR A GAS TURBINE - A laminated sheet for a gas turbine component, the laminated sheet has a first cover layer, a second cover layer and a first intermediate layer, wherein the first cover layer, the second cover layer and the first intermediate layer are stacked together on top of each other. The first intermediate layer is located between the first cover layer and the second cover layer. The first intermediate layer has at least one first elongated through hole, wherein a cooling fluid is flowable through the first elongated through hole. | 01-21-2016 |
| 20160024970 | COMPRESSOR ASSEMBLY FOR GAS TURBINE - A compressor assembly, and more in general relates to a compressor for a gas turbine providing a solution that teaches to locate within a cavity formed by the outer casing of the compressor and the inner vane carrier a separator element, or membrane, such to divide the cavity into two sub-cavities. This advantageously results in a more flexible design with respect to the positioning of the flange blow-off extractor and to the cavity sizing, as the flange position is not necessarily the boundary for the flow anymore as it would be without the separator element. | 01-28-2016 |
| 20160032730 | OBTUSE ANGLE CHEVRON TRIP STRIP - An airfoil includes a cooling air passage for receiving a cooling air flow. A chevron including a first rib and a second rib extends from a common tip is disposed within the cooling passage for generating a turbulent flow to improve heat transfer. The chevron includes an angle between the first rib and the second rib that is greater than 90 degrees. | 02-04-2016 |
| 20160032748 | GUIDE BLADE FOR A GAS TURBINE - A guide blade for a gas turbine is disclosed. The guide blade includes a blade leaf having a receptacle in which at least one sealing element is arranged, where the sealing element is movable relative to the blade leaf between a sealing setting, in which the sealing element is at least partially moved out of the receptacle, and a storage setting, in which the sealing element is moved back into the receptacle. The guide blade further includes at least one fluid channel by which fluid under pressure can be routed into the receptacle in order to move the sealing element from the storage setting into the sealing setting. An inlet opening of the fluid channel is formed on a pressure-side surface of the blade leaf. A housing as well as a gas turbine having at least one guide blade is also disclosed. | 02-04-2016 |
| 20160047250 | SHOWERHEAD HOLE SCHEME APPARATUS AND SYSTEM - The gas turbine component showerhead cooling hole layouts described herein include minimal lateral cooling hole exit diffusion on the middle showerhead cooling hole rows and interior facing sides of outer rows. In this way, rows of cooling holes may be placed close together. Stated another way, the outer showerhead cooling hole rows substantially only include lateral cooling hole exit diffusion in the direction away from the other rows to again allow the rows to be placed close together. | 02-18-2016 |
| 20160069302 | COMPRESSOR OF EXHAUST TURBOCHARGER - A compressor includes an impeller, a housing, and a connecting shaft. An inlet for introducing gas into a suction side of the impeller, and an annular groove formed in an outer periphery of the inlet are formed in the housing. The groove communicates with the inlet on the intake passage side. The groove is blocked on the impeller side. A gas outlet of an EGR passage is connected with a middle of the groove. An inner diameter of the groove is machined to be larger than an inner diameter of an end part of the gas outlet of the intake passage (an end part on the inlet side). | 03-10-2016 |
| 20160069350 | Stator Disk and Vacuum Pump - A stator disk includes a connection hole for improving exhaust efficiency in a vacuum pump including a Seigbahn type molecular pump portion, and a vacuum pump including the stator disk. The vacuum pump according to an embodiment includes a Seigbahn type molecular pump portion and includes, in a stator disk disposed therein, a connection hole that connects an upper space (an inlet port side region, an upstream side region) with a lower space (an outlet port side region, a downstream side region) in the axial direction of the stator disk. | 03-10-2016 |
| 20160076382 | COMPONENT CORE WITH SHAPED EDGES - A gas turbine engine component according to an exemplary aspect of the present disclosure includes, among other things, at least one cooling passage. The at least one cooling passage includes a first wall and a second wall bounding the at least one cooling passage, the first wall having a plurality of first surface features and the second wall having a plurality of second surface features. The plurality of first surface features and the plurality of second surface features are arranged such that a width of the cooling passage varies along a length of the cooling passage defined by the plurality of first surface features and the plurality of second surface features. The plurality of first surface features has a first profile, and the plurality of second surface features has a second, different profile. A casting core for forming cooling passages in an aircraft component is also disclosed. | 03-17-2016 |
| 20160076384 | GAS TURBINE ENGINE COMPONENT HAVING ENGINEERED VASCULAR STRUCTURE - A component according to an exemplary aspect of the present disclosure includes, among other things, a wall and a hollow vascular engineered lattice structure formed inside of the wall. The hollow vascular engineered lattice structure has an inlet hole and an outlet hole that communicate fluid into and out of the hollow vascular structure. The hollow vascular engineered lattice structure further has at least one resupply inlet hole between the inlet hole and the outlet hole to communicate additional fluid into the hollow vascular engineered lattice structure. | 03-17-2016 |
| 20160084089 | INTERNALLY DAMPED AIRFOILED COMPONENT AND METHOD - An airfoiled component comprises: a root section, an airfoil section, a damper pocket enclosed within a portion of the airfoil section, and a damper. The airfoil section includes a suction sidewall and a pressure sidewall each extending chordwise between a leading edge and a trailing edge, and extending spanwise between the root section and an airfoil tip. The damper includes a fixed end unified with a damper mounting surface, and a free end extending into the damper pocket from the damper mounting surface. | 03-24-2016 |
| 20160090846 | GAS TURBINE ENGINE AIRFOIL TRAILING EDGE SUCTION SIDE COOLING - An airfoil for a gas turbine engine includes an outer airfoil wall that provides an exterior surface and multiple radially extending cooling passages. The exterior surface provides pressure and suctions sides joined by leading and trailing edges. The cooling passages include a supply passage arranged upstream from and in fluid communication with a trailing edge passage. A cooling hole extends through the outer airfoil wall from the supply passage to the exterior surface on the suction side. | 03-31-2016 |
| 20160090852 | GAS TURBINE ENGINE - In a gas turbine engine, an annular intermediate-temperature space is formed by a first partition member extending from an inner platform radially inward and a second partition member having a radially inner end portion fixed to the first partition member and a radially outer end portion facing the inner platform, the inner platform being connected to radially inner ends of nozzle guide vanes of a turbine, and the intermediate-temperature space overlapping with the inner platform over substantially its axial entire length. The low-temperature space, supplied with a low-temperature air from a compressor, is formed on a radially inside of the second partition member. Therefore, a temperature of the intermediate-temperature space is maintained at an intermediate temperature between those of a combustion gas and the air, reducing a temperature difference of the inner platform from an outer platform and the vanes to reduce a thermal stress. | 03-31-2016 |
| 20160090904 | SEALING-COUPLED APPARATUS OF TURBOCHARGER - A sealing-coupled apparatus of a turbocharger may include a vane cover installed in front of a turbine vane set in a turbine housing, the vane cover configured to cover the turbine vane and to define a fluid passage through which a fluid may be fed to an outlet hole extending in the turbine housing, and a sealing member having a tubular shape with a predetermined thickness, the sealing member being engaged with a front end of the vane cover and having an annular groove formed around an outer circumferential surface of the sealing member, with a sealing ring fitted over the groove and spacing the vane cover apart from the turbine housing in a radial direction. | 03-31-2016 |
| 20160102563 | GAS TURBINE ENGINE COMPONENT HAVING TRIP STRIPS - Disclosed is a gas turbine engine including a compressor section and a turbine section. The gas turbine engine includes a gas turbine engine component having a first wall providing an outer surface of the gas turbine engine component and a second wall spaced-apart from the first wall. The first wall is a gas-path wall exposed to a core flow path of the gas turbine engine. The second wall is a non-gas-path wall. A cooling passageway is provided between the second wall and the first wall. The second wall has a trip strip provided thereon. | 04-14-2016 |
| 20160102568 | POWER TURBINE HEAT SHIELD ARCHITECTURE - A power turbine section for a gas turbine engine includes a heat shield assembly mounted to a bearing support, the heat shield assembly forms an outer diameter directed toward an inner vane platform of the power turbine vane array. | 04-14-2016 |
| 20160115803 | PLATFORM COOLING CIRCUIT FOR A GAS TURBINE ENGINE COMPONENT - A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having a first path side and a second path side and a platform cooling circuit disposed on one of the first path side and the second path side of the platform. The platform cooling circuit includes a first core cavity, a cavity in fluid communication with the first core cavity, and a cover plate positioned to cover at least the cavity. | 04-28-2016 |
| 20160146018 | TAPERED COOLING CHANNEL FOR AIRFOIL - The present invention includes systems and methods for providing cooling channels located within walls of a turbine airfoil. These cooling channels include micro-circuits that taper in various directions along the length and width of the airfoil. In addition, these cooling channels have a variety of shapes and areas to facilitate convective heat transfer between the surrounding air and the airfoil. | 05-26-2016 |
| 20160146019 | COOLING CHANNEL FOR AIRFOIL WITH TAPERED POCKET - The present invention includes systems and methods for providing cooling channels located within walls of a turbine airfoil. These cooling channels include micro-circuits that taper in various directions along the length and width of the airfoil. In addition, these cooling channels have a variety of shapes and areas to facilitate convective heat transfer between the surrounding air and the airfoil. | 05-26-2016 |
| 20160153283 | GAS TURBINE ENGINE COMPONENT WITH CONVERGING/DIVERGING COOLING PASSAGE | 06-02-2016 |
| 20160153285 | TURBINE BLADE | 06-02-2016 |
| 20160153299 | TURBINE NOZZLE WITH IMPINGEMENT BAFFLE | 06-02-2016 |
| 20160153467 | TURBOMACHINE BLADE, COMPRISING INTERSECTING PARTITIONS FOR CIRCULATION OF AIR IN THE DIRECTION OF THE TRAILING EDGE | 06-02-2016 |
| 20160160657 | STAGGERED CROSSOVERS FOR AIRFOILS - An airfoil according to an exemplary aspect of the present disclosure includes, among other things, an airfoil section having an external wall and an internal wall. The internal wall defines a first reference plane extending in a spanwise direction and through a thickness of the internal wall. A first cavity and a second cavity are separated by the internal wall. A plurality of crossover passages within the internal wall connects the first cavity to the second cavity. Each of the plurality of crossover passages defines a passage axis. The plurality of crossover passages are distributed in the spanwise direction and arranged such that the passage axis of each of the plurality of cooling passages intersects a surface of the second cavity. The plurality of crossover passages include a first set of crossover passages and a second set of crossover passages positioned on opposite sides of the first reference plane. The passage axis of each of the first set of crossover passages is arranged at a first vertical angle relative to a spanwise axis, and the passage axis of each of the second set of crossover passages is arranged at a second, different vertical angle relative to the spanwise axis. A casting core for an airfoil is also disclosed. | 06-09-2016 |
| 20160160757 | AIR INTAKE ARRANGEMENT - An intake for channeling air flowing past a propeller to an inlet of an aircraft engine driving the propeller through a drive shaft comprising: a static cowling extending about an axis and which flares from an upstream end, an intake slot formed in the flared portion which bounds a passage to the inlet of the aircraft engine, the intake slot opening over an arc that extends less than 360 degrees of the circumference of the cowling and having an axially rearward edge that blends into the cowling through a curve having a vertex, wherein the vertex is radially inside a projected extension of the flared portion from an upstream portion of the flared cowling across the intake slot. | 06-09-2016 |
| 20160167112 | CERAMIC CORE FOR COMPONENT CASTING | 06-16-2016 |
| 20160169000 | HEAT TRANSFER PEDESTALS WITH FLOW GUIDE FEATURES | 06-16-2016 |
| 20160169016 | REVERSIBLE FLOW BLADE OUTER AIR SEAL | 06-16-2016 |
| 20160169051 | WATER REMOVAL DEVICE FOR STEAM TURBINE AND METHOD FOR FORMING SLIT | 06-16-2016 |
| 20160177737 | GAS TURBINE ENGINE COMPONENT WITH FILM COOLING HOLE WITH ACCUMULATOR | 06-23-2016 |
| 20160177741 | AEROFOIL BLADE OR VANE | 06-23-2016 |
| 20160177782 | ENGINE COMPONENT HAVING PLATFORM WITH PASSAGEWAY | 06-23-2016 |
| 20160186576 | LARGE-FOOTPRINT TURBINE COOLING HOLE - A cooling hole for a component includes a meter section and a diffuser section. The diffuser section has a footprint region defined by five sides, a first side of the five sides extending along substantially an entire height of the diffuser section and second and third sides of the five sides meeting in an obtuse angle opposite the first side. A component having the cooling hole and a method of forming the cooling hole are also disclosed. | 06-30-2016 |
| 20160186664 | GAS TURBINE SEALING - A gas turbine engine having a turbine that includes a stator blade and a rotor blade having a seal formed in a trench cavity defined therebetween. The seal may include: a stator overhang extending from the stator blade toward the rotor blade so to include an overhang topside, and, opposite the overhang topside, an overhang underside; a rotor outboard face extending radially inboard from a platform edge, the rotor outboard face opposing at least a portion of the overhang face across the axial gap of the trench cavity; an axial projection extending from the rotor outboard face toward the stator blade so to axially overlap with the stator overhang; and an interior cooling channel extending through the stator overhang to a port formed through the overhang underside. The port may be configured to direct a coolant expelled therefrom toward the axial projection. | 06-30-2016 |
| 20160194965 | PARTIAL TIP FLAG | 07-07-2016 |
| 20160194978 | Turbine stator vane with insert and flexible seal | 07-07-2016 |
| 20160201469 | MATEFACE SURFACES HAVING A GEOMETRY ON TURBOMACHINERY HARDWARE | 07-14-2016 |
| 20160201475 | COOLING FOR GAS TURBINE ENGINE COMPONENTS | 07-14-2016 |
| 20160201476 | AIRFOIL FOR A TURBINE ENGINE | 07-14-2016 |
| 20160201487 | SLIDING BAFFLE INSERTS | 07-14-2016 |
| 20160251969 | GAS TURBINE ENGINE AIRFOIL | 09-01-2016 |
| 20160376896 | GAS TURBINE ENGINE COMPONENT COOLING WITH RESUPPLY OF COOLING PASSAGE - A gas turbine engine component includes a structure having a wall that provides an exterior surface. A first cooling passage is arranged adjacent to and interiorly of the wall. A second cooling passage is arranged in the wall and provides a first fluid flow direction. A resupply channel is arranged in the wall and is fluidly interconnected to the second cooling passage. A resupply hole fluidly interconnects the first cooling passage and the resupply channel. The resupply channel is transverse relative to the second cooling passage to provide a second fluid flow direction that extends from the resupply hole to the second cooling passage. | 12-29-2016 |
| 20160376899 | GUIDE VANE ASSEMBLY ON THE BASIS OF A MODULAR STRUCTURE - The invention relates to a guide vane assembly of a turbomachine based on a modular structure, wherein the guide vane elements include at least an airfoil, an inner platform, an outer platform, wherein the guide vane airfoil and/or platforms have at its one ending provisions for connection of the guide vane elements among each other. The connections of the guide vane elements among each other are configured as a detachable, permanent or semi-permanent fixation with respect to the radial or quasi-radial extension of the airfoil compared to the rotor axis of the turbomachine. The assembling of the airfoil with respect to at least one platform is based on a force-fit and/or a form-fit connection, or on the use of a metallic and/or ceramic fitting surface, or on force closure means with a detachable, permanent or semi-permanent fixation. | 12-29-2016 |
| 20170234138 | Airfoil Trailing Edge Cooling | 08-17-2017 |
| 20170234145 | ACCELERATOR INSERT FOR A GAS TURBINE ENGINE AIRFOIL | 08-17-2017 |
| 20170234150 | Thermal Stress Relief Of A Component | 08-17-2017 |
| 20170234151 | AIR SHREDDER INSERT | 08-17-2017 |
| 20180023400 | BLADE WITH INTERNAL RIB HAVING CORRUGATED SURFACE(S) | 01-25-2018 |
| 20180023403 | TURBINE VANE WITH COUPON HAVING CORRUGATED SURFACE(S) | 01-25-2018 |
| 20180023406 | INTERMEDIATE CASE FOR AN AIRCRAFT TURBOMACHINE MADE FROM A SINGLE CASTING WITH A LUBRICANT DUCT | 01-25-2018 |
| 20190145264 | OGV ELECTROFORMED HEAT EXCHANGERS | 05-16-2019 |
| 20220136397 | FABRICATED CMC NOZZLE ASSEMBLIES FOR GAS TURBINE ENGINES - Nozzle segment assemblies for gas turbine engines include an outer band, inner band, and at least one airfoil body between the outer band and inner band. The airfoil body includes an outer end extending through the outer band and having an outer end reinforced wall portion engaging with the outer band and an inner end extending through the inner band and having an inner end reinforced wall portion engaging with the inner band. A thickness of the outer end and inner end reinforced wall portions are each greater than a thickness of a central wall portion of the airfoil body to reduce stresses in the airfoil body. The outer band, the inner band, and the airfoil bodies are ceramic matrix composite (CMC) materials. Direct mounting of the airfoil bodies to the inner and outer hangers of the nozzle further reduce stress. | 05-05-2022 |
