Patent application number | Description | Published |
20080206042 | METHODS AND SYSTEM FOR RECUPERATED CIRCUMFERENTIAL COOLING OF INTEGRAL TURBINE NOZZLE AND SHROUD ASSEMBLIES - A method for cooling a shroud segment of a gas turbine engine is provided. The method includes providing a turbine shroud assembly including a shroud segment having an inner surface and a leading edge that is substantially perpendicular to the inner surface, and coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge. The method also includes directing cooling air into the gap, circumferentially mixing the cooling air in a plenum defined within the leading edge to substantially uniformly distribute the cooling air throughout the gap, and directing the cooling air in the gap through at least one cooling hole formed between the plenum and the inner surface. | 08-28-2008 |
20080229751 | Cooling system for gas turbine engine having improved core system - A gas turbine engine having a longitudinal centerline axis therethrough, including: a fan section at a forward end of the gas turbine engine including at least a first fan blade row connected to a first drive shaft; a booster compressor positioned downstream of and in at least partial flow communication with the fan section including a plurality of stages, each stage including a stationary compressor blade row and a rotating compressor blade row connected to a drive shaft and interdigitated with the stationary compressor blade row; a core system positioned downstream of the booster compressor, the core system further comprising a combustion system for producing pulses of gas having increased pressure and temperature from a fluid flow provided to an inlet thereof so as to produce a working fluid at an outlet; a low pressure turbine positioned downstream of and in flow communication with the core system, the low pressure turbine being utilized to power the first drive shaft; and, a system for cooling the combustion system, wherein fuel is utilized as a cooling fluid prior to being supplied to the combustion system. The core system may further include an intermediate compressor positioned downstream of and in flow communication with the compressor connected to a second drive shaft; and an intermediate turbine positioned downstream of the combustion system in flow communication with the working fluid. | 09-25-2008 |
20080230378 | Methods and systems for forming tapered cooling holes - A method for forming holes in an object is provided. The method includes providing an electrochemical machining (ECM) electrode including a first section having insulation that circumscribes the first section, and a second section having insulation that extends only partially around the second section. The method also includes inserting the electrode into the object, such that in a single pass the electrode forms a hole that includes a first portion having a first cross-sectional area and a second portion having a second cross-sectional area. | 09-25-2008 |
20080230379 | Methods and systems for forming cooling holes having circular inlets and non-circular outlets - A method for forming a hole in an object is provided. The method includes forming a starter hole in the object, providing an electrochemical machining electrode that includes insulation that extends only partially around the electrode, and inserting the electrode into the starter hole to form a hole in the object that has an inlet defined by a first cross-sectional area and an outlet defined by a second cross-sectional area. | 09-25-2008 |
20080230396 | Methods and systems for forming turbulated cooling holes - A method for forming holes in an object is provided. The method includes forming a starter hole in the object, providing an electrochemical machining electrode having at least one insulated section that substantially circumscribes the electrode and at least one uninsulated section, and inserting the electrode into the starter hole to facilitate forming a hole defined by at least one first section having a first cross-sectional area and at least one second section having a second cross-sectional area. | 09-25-2008 |
20080281562 | METHODS FOR OPTIMIZING PARAMETERS OF GAS TURBINE ENGINE COMPONENTS - Methods for optimizing at least one operating parameter of an engine component using an experimentally measured 3D flow field involving providing a magnetic resonance imaging machine, providing a model of an engine component, placing the model into the magnetic resonance imaging machine with a fluid flow source for applying an external fluid flow, applying the external fluid flow to the model, collecting data related to the external fluid flow about the model, and optimizing at least one operating parameter of the component using the data. | 11-13-2008 |
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 |
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 |
20090155088 | DUST HOLE DOME BLADE - A hollow turbine airfoil includes a tip cap bounding an internal cooling circuit between opposite pressure and suction sidewalls. The tip cap includes an internal dome surrounding a dust hole, and the dome is inclined inwardly toward the airfoil root both transversely between the opposite sidewalls and chordally between opposite leading and trailing edges of the airfoil. | 06-18-2009 |
20090165275 | METHOD FOR REPAIRING A COOLED TURBINE NOZZLE SEGMENT - A method for repairing a turbine nozzle segment having a band and a plurality of airfoils, where the band has a flange is described. The method includes the steps of repairing a damaged area on the turbine nozzle segment and drilling a plurality of cooling holes in the flange. | 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 |
20090324422 | Cascade tip baffle airfoil - A turbine blade includes an airfoil tip with first and second ribs extending along the opposite pressure and suction sides. The ribs extend outwardly from a tip floor and are joined together at opposite leading and trailing edges. A cascade tip baffle transversely bridges the two ribs above the tip floor forward of the maximum width of the tip to partition the tip chordally into corresponding tip pockets on opposite sides of the baffle. | 12-31-2009 |
20100025001 | METHODS FOR FABRICATING GAS TURBINE COMPONENTS USING AN INTEGRATED DISPOSABLE CORE AND SHELL DIE - Methods involving providing an integrated disposable core and shell die of an authentic gas turbine component, inserting at least one through-rod through the integrated disposable core and shell die, casting an integrated core and shell mold inside of the integrated disposable core and shell die, removing the integrated disposable core and shell die to obtain the integrated core and shell casting mold having the at least one through-rod disposed therein, casting an authentic gas turbine component replica using the integrated core and shell casting mold, and removing the integrated core and shell casting mold and the at least one through-rod to obtain the authentic gas turbine component replica. | 02-04-2010 |
20100034647 | Processes for the formation of positive features on shroud components, and related articles - A process for the formation of positive features on the surface of a turbine shroud component is described. The process involves applying a feature-forming material to a selected portion of the component surface with a laser consolidation apparatus, according to a pre-selected shape and size for the positive features. A gas turbine engine, comprising a shroud component which contains positive features formed according to embodiments of this process, represents another embodiment of this invention. Methods for modifying the shape of at least one positive feature on a surface of a shroud component are also described. | 02-11-2010 |
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 |
20100047060 | Plasma Enhanced Compressor - A gas turbine engine is disclosed, comprising a compressor having a circumferential row of blades, a casing surrounding the tips of the blades, located radially apart from the tips of the blades and at least one plasma generator located on the casing. The plasma generator comprises a first electrode and a second electrode separated by a dielectric material. The gas turbine engine further comprises an engine control system which controls the operation of the plasma generator such that the stable operating range of the compressor is increased. | 02-25-2010 |
20100074745 | DUAL STAGE TURBINE SHROUD - A turbine shroud includes a shroud hanger having an arcuate panel from which three inner hooks extend inwardly, and from which two outer hooks extend outwardly therefrom. The two outer hooks effect a statically determinate configuration of the shroud. | 03-25-2010 |
20100111682 | CRENELATED TURBINE NOZZLE - A turbine nozzle includes a row of vanes extending radially between annular outer and inner bands. The outer band includes a pair of radial flanges defining an annular seal groove therebetween. One of the flanges is crenelated to improve nozzle life. | 05-06-2010 |
20100143139 | BANKED PLATFORM TURBINE BLADE - A turbine blade includes an airfoil and integral platform at the root thereof. The platform is contoured in elevation from a bank adjacent the pressure side of the airfoil to a trough commencing behind the airfoil leading edge. | 06-10-2010 |
20100158696 | CURVED PLATFORM TURBINE BLADE - A turbine blade includes an airfoil and integral platform at the root thereof. The platform is contoured in elevation from a ridge to a trough, and is curved axially to complement the next adjacent curved platform. | 06-24-2010 |
20100170224 | PLASMA ENHANCED BOOSTER AND METHOD OF OPERATION - A booster system is disclosed, comprising a first rotor stage having a plurality of first rotor blades spaced circumferentially around a rotor hub with a longitudinal axis and having a first pitch-line radius extending from the longitudinal axis, a last rotor stage located axially aft from the first rotor stage, the last rotor stage comprising a plurality of last rotor blades spaced circumferentially around the longitudinal axis and having a second pitch-line radius extending from the longitudinal axis, and a gooseneck duct located axially aft from the last rotor stage and capable of receiving an airflow, the gooseneck duct comprising an inlet end and an exit end located at a distance axially aft from the inlet end and having at least one plasma actuator mounted in the gooseneck duct. | 07-08-2010 |
20100172747 | PLASMA ENHANCED COMPRESSOR DUCT - A compression system is disclosed, comprising a first compressor having a first flowpath, a second compressor having a second flowpath located axially aft from the first compressor, and a transition duct capable of flowing an airfow from the first compressor to the second compressor, the transition duct having at least one plasma actuator mounted in the transition duct. | 07-08-2010 |
20100221122 | Flared tip turbine blade - A turbine blade includes an airfoil terminating in a tip. The tip includes a first rib conforming with a concave pressure side of the airfoil, and a second rib conforming with a convex suction side of the airfoil. The second rib is flared outwardly from the suction side. | 09-02-2010 |
20100251696 | PLASMA ENHANCED RAPIDLY EXPANDED GAS TURBINE ENGINE TRANSITION DUCT - A plasma enhanced rapidly expanded duct system includes a gas turbine engine inter-turbine transition duct having radially spaced apart conical inner and outer duct walls extending axially between a duct inlet and a duct outlet. A conical plasma generator produces a conical plasma along the outer duct wall. An exemplary embodiment of the conical plasma generator is mounted to the outer duct wall and including radially inner and outer electrodes separated by a dielectric material. The dielectric material is disposed within a conical groove in a radially inwardly facing surface of the outer duct wall. An AC power supply is connected to the electrodes to supply a high voltage AC potential to the electrodes. | 10-07-2010 |
20100282721 | SYSTEM AND METHOD FOR IMPROVED FILM COOLING - A system for producing at least one trench to improve film cooling in a sample is provided. The system includes at least one laser source outputting at least one pulsed laser beam. The pulsed laser beam includes a pulse duration including a range less than about 50 μs, an energy per pulse having a range less than about 0.1 Joule, and a repetition rate with a range greater than about 1000 Hz. The system also includes a control subsystem coupled to the laser source, the control subsystem configured to synchronize a position of the sample with the pulse duration and energy level in order to selectively remove at least one of a thermal barrier coating, a bondcoat and a substrate metal in the sample to form the at least one trench. | 11-11-2010 |
20100284780 | Method of Operating a Compressor - A method of operating a compressor having a row of blades for preventing a compressor stall is disclosed, the method comprising the steps of mounting a plasma generator in a casing or a shroud radially outwardly and apart from the blade tips wherein the plasma generator comprises a radially inner electrode and a radially outer electrode separated by a dielectric material; and supplying an AC potential to the radially inner electrode and the radially outer electrode. | 11-11-2010 |
20100284795 | Plasma Clearance Controlled Compressor - A plasma leakage flow control system for a compressor is disclosed, comprising a circumferential row of compressor blades, an annular casing surrounding the tips of the blades, located radially apart from the tips of the blades and at least one annular plasma generator located on the annular casing. The annular plasma generator comprises an inner electrode and an outer electrode separated by a dielectric material. A gas turbine engine having a plasma leakage flow control system further comprises an engine control system which controls the operation of the annular plasma generator such that the blade tip leakage flow can be changed. | 11-11-2010 |
20110033312 | COMPOUND COOLING FLOW TURBULATOR FOR TURBINE COMPONENT - Multi-scale turbulation features, including first turbulators ( | 02-10-2011 |
20110150653 | Plasma Induced Flow Control of Boundary Layer at Airfoil Endwall - Plasma generators ( | 06-23-2011 |
20110171023 | AIRFOIL INCORPORATING TAPERED COOLING STRUCTURES DEFINING COOLING PASSAGEWAYS - A gas turbine engine ( | 07-14-2011 |
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 |
20110262695 | Discreetly Defined Porous Wall Structure for Transpirational Cooling - A wall structure ( | 10-27-2011 |
20110293434 | METHOD OF CASTING A COMPONENT HAVING INTERIOR PASSAGEWAYS - A method of casting a component ( | 12-01-2011 |
20110302924 | COOLED CONDUIT FOR CONVEYING COMBUSTION GASES - A conduit through which hot combustion gases pass in a gas turbine engine. The conduit includes a wall structure having a central axis and defining an inner volume of the conduit for permitting hot combustion gases to pass through the conduit. The wall structure includes a forward end, an aft end axially spaced from the forward end, the aft end defining a combustion gas outlet for the hot combustion gases passing through the conduit, and a plurality of generally radially outwardly extending protuberances formed in the wall structure. The protuberances each include at least one cooling fluid passage formed therethrough for permitting cooling fluid to enter the inner volume. At least one of the protuberances is shaped so as to cause cooling fluid passing through it to diverge in a circumferential direction as it enters into the inner volume. | 12-15-2011 |
20110305582 | Film Cooled Component Wall in a Turbine Engine - A component wall in a turbine engine. The component wall includes a substrate, a trench, and a plurality of cooling passages. The substrate has a first surface and a second surface opposed from the first surface. The trench is located in the second surface and is defined by a bottom surface between the first and second surfaces, a first sidewall, and a second sidewall spaced from the first sidewall. The first sidewall extends radially outwardly continuously from the bottom surface of the trench to the second surface. The first sidewall includes a plurality of first protuberances extending toward the second sidewall. The cooling passages extend through the substrate from the first surface to the bottom surface of the trench. Outlets of the cooling passages are arranged within the trench such that cooling air exiting the cooling passages is directed toward respective ones of the first protuberances. | 12-15-2011 |
20110305583 | COMPONENT WALL HAVING DIFFUSION SECTIONS FOR COOLING IN A TURBINE ENGINE - A film cooling structure formed in a component wall of a turbine engine and a method of making the film cooling structure. The film cooling structure includes a plurality of individual diffusion sections formed in the wall, each diffusions section including a single cooling passage for directing cooling air toward a protuberance of a wall defining the diffusion section. The film cooling structure may be formed with a masking template including apertures defining shapes of a plurality of to-be-formed diffusion sections in the wall. A masking material can be applied to the wall into the apertures in the masking template so as to block outlets of cooling passages exposed through the apertures. The masking template can be removed and a material may be applied on the outer surface of the wall such that the material defines the diffusion sections once the masking material is removed. | 12-15-2011 |
20120006028 | DAMPING RESONATOR WITH IMPINGEMENT COOLING - A resonance chamber ( | 01-12-2012 |
20120006518 | MESH COOLED CONDUIT FOR CONVEYING COMBUSTION GASES - A conduit through which hot combustion gases pass in a gas turbine engine. The conduit includes a wall structure having an inner surface, an outer surface, a region, an inlet, and an outlet. The inner surface defines an inner volume of the conduit. The region extends between the inner and outer surfaces and includes cooling fluid structure defining a plurality of cooling passageways. The inlet extends inwardly from the outer surface and provides fluid communication between the inlet and the passageways. The outlet extends from the passageways to the inner surface to provide fluid communication between the passageways and the inner volume. At least one first cooling passageway intersects with at least one second cooling passageway such that cooling fluid flowing through the first cooling passageway interacts with cooling fluid flowing through the second cooling passageway. | 01-12-2012 |
20120007318 | SEAL INCLUDING FLEXIBLE SEAL STRIPS - A seal member for effecting a seal preventing fluid flow in an axial direction through an annular space formed between two relatively moving components including a rotatable shaft and a stator structure. The seal member includes a plurality of flexible seal strips. Each seal strip includes a planar plate extending radially through the annular space and having a radially outer end supported to the stator structure and a radially inner end defining a tip portion extending widthwise in the axial direction engaged in sliding contact with a peripheral surface of the rotatable shaft. At least one of the seal strips includes a plurality of perforations extending through the seal strip and located between a leading edge and a trailing edge of the seal strip for effecting an increased flexibility of the seal strip adjacent to the tip portion. | 01-12-2012 |
20120014808 | NEAR-WALL SERPENTINE COOLED TURBINE AIRFOIL - A serpentine coolant flow path ( | 01-19-2012 |
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 |
20120057968 | RING SEGMENT WITH SERPENTINE COOLING PASSAGES - A ring segment for a gas turbine engine includes a panel and a cooling system. The cooling system receives cooling fluid from an outer side of the panel for cooling the panel and includes at least one cooling fluid supply passage, at least one serpentine cooling passage, and at least one cooling fluid discharge passage. The cooling fluid supply passage(s) receive the cooling fluid from the outer side of the panel and deliver the cooling fluid to a first cooling fluid chamber within the panel. The serpentine cooling passage(s) receive the cooling fluid from the first cooling fluid chamber, wherein the cooling fluid provides convective cooling to the panel as it passes through the serpentine cooling passage(s). The cooling fluid discharge passage(s) discharge the cooling fluid from the cooling system. | 03-08-2012 |
20120070302 | TURBINE AIRFOIL VANE WITH AN IMPINGEMENT INSERT HAVING A PLURALITY OF IMPINGEMENT NOZZLES - A turbine airfoil vane usable in a turbine engine and including at least one cooling system with an impingement plate having one or more impingement nozzles is disclosed. The turbine vane impingement nozzles may extend towards an outer wall forming the turbine vane and may reduce the mixing of cooling fluids and impingement jets found in conventional configurations. Instead, the nozzles terminate within close proximity of the outer wall, thereby reducing the effect of cooling fluid cross flow. | 03-22-2012 |
20120070306 | TURBINE COMPONENT COOLING CHANNEL MESH WITH INTERSECTION CHAMBERS | 03-22-2012 |
20120076644 | COOLED COMPONENT WALL IN A TURBINE ENGINE - A component wall in a turbine engine includes a substrate, a diffusion section, and at least one cooling passage. The diffusion section is located in a surface of the substrate and is defined by a first sidewall and a second sidewall. The cooling passage(s) include an outlet portion through which cooling air exits in a direction toward the first sidewall. The outlet portion includes a rear section, a front section, and an inner wall having proximal and distal ends. The rear section is located between the first and second sidewalls. The front section extends between the first sidewall and the distal end of the inner wall. The first sidewall extends into the outlet portion of the cooling passage(s) to the inner wall and extends from the first lateral wall to the second lateral wall so as to block the front section of the outlet portion. | 03-29-2012 |
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 |
20120121408 | TURBINE TRANSITION COMPONENT FORMED FROM A TWO SECTION, 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 |
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 |
20120177503 | COMPONENT COOLING CHANNEL - A cooling channel ( | 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 |
20120201674 | COOLING MODULE DESIGN AND METHOD FOR COOLING COMPONENTS OF A GAS TURBINE SYSTEM - A cooling arrangement in a gas turbine system ( | 08-09-2012 |
20120207591 | COOLING SYSTEM HAVING REDUCED MASS PIN FINS FOR COMPONENTS IN A GAS TURBINE ENGINE - A cooling system having one or more pin fins with reduced mass for a gas turbine engine is disclosed. The cooling system may include one or more first surfaces defining at least a portion of the cooling system. The pin fin may extend from the surface defining the cooling system and may have a noncircular cross-section taken generally parallel to the surface and at least part of an outer surface of the cross-section forms at least a quartercircle. A downstream side of the pin fin may have a cavity to reduce mass, thereby creating a more efficient turbine airfoil. | 08-16-2012 |
20120207614 | INTEGRATED AXIAL AND TANGENTIAL SERPENTINE COOLING CIRCUIT IN A TURBINE AIRFOIL - A continuous serpentine cooling circuit forming a progression of radial passages ( | 08-16-2012 |
20120269647 | COOLED AIRFOIL IN A TURBINE ENGINE - An airfoil in a gas turbine engine includes an outer wall and an inner wall. The outer wall includes a leading edge, a trailing edge opposed from the leading edge in a chordal direction, a pressure side, and a suction side. The inner wall is coupled to the outer wall at a single chordal location and includes portions spaced from the pressure and suction sides of the outer wall so as to form first and second gaps between the inner wall and the respective pressure and suction sides. The inner wall defines a chamber therein and includes openings that provide fluid communication between the respective gaps and the chamber. The gaps receive cooling fluid that provides cooling to the outer wall as it flows through the gaps. The cooling fluid, after traversing at least substantial portions of the gaps, passes into the chamber through the openings in the inner wall. | 10-25-2012 |
20120269648 | SERPENTINE COOLING CIRCUIT WITH T-SHAPED PARTITIONS IN A TURBINE AIRFOIL - A serpentine cooling circuit (AFT) in a turbine airfoil ( | 10-25-2012 |
20120272521 | METHOD OF FABRICATING A NEARWALL NOZZLE IMPINGEMENT COOLED COMPONENT FOR AN INTERNAL COMBUSTION ENGINE - A method of forming an internal combustion engine component having a multi-panel outer wall. The multi-panel outer wall has an inner panel ( | 11-01-2012 |
20120282108 | TURBINE BLADE WITH CHAMFERED SQUEALER TIP AND CONVECTIVE COOLING HOLES - A squealer tip formed from a pressure side rib and a suction side rib extending radially outward from a tip of the turbine blade is disclosed. The pressure and suction side ribs may be positioned along the pressure side and the suction side of the turbine blade, respectively. The pressure and suction side ribs may include chamfered leading edges with film cooling holes having exhaust outlets positioned therein. The film cooling holes may be configured to be diffusion cooling holes with one or more tapered sections for reducing the velocity of cooling fluids and increasing the size of the convective surfaces. | 11-08-2012 |
20130017080 | FLOW DIRECTING MEMBER FOR GAS TURBINE ENGINE - In a gas turbine engine, a flow directing member includes a platform supported on a rotor, a radially facing endwall, at least one axial surface extending radially inwardly from a junction with the endwall, an airfoil extending radially outwardly from the endwall, and a fluid flow directing feature. The fluid flow directing feature includes a groove extending axially into the axial surface and has radially inner and outer groove ends. The outer groove end defines an axially extending notch in the junction between the axial surface and the endwall and forms an opening in the endwall for directing a cooling fluid to the endwall. The groove further includes a first groove wall extending from the inner groove end to the outer groove end, and a second groove wall opposed from the first groove wall and extending from the inner groove end to the outer groove end. | 01-17-2013 |
20130017095 | FLOW DIRECTING MEMBER FOR GAS TURBINE ENGINEAANM Lee; Ching-PangAACI CincinnatiAAST OHAACO USAAGP Lee; Ching-Pang Cincinnati OH USAANM Tham; Kok-MunAACI OrlandoAAST FLAACO USAAGP Tham; Kok-Mun Orlando FL USAANM Vitt; Paul H.AACI Liberty TwpAAST OHAACO USAAGP Vitt; Paul H. Liberty Twp OH USAANM Williamson; Stephen R.AACI CincinnatiAAST OHAACO USAAGP Williamson; Stephen R. Cincinnati OH USAANM Montgomery; Matthew D.AACI JupiterAAST FLAACO USAAGP Montgomery; Matthew D. Jupiter FL USAANM Prakash; ChanderAACI OrlandoAAST FLAACO USAAGP Prakash; Chander Orlando FL USAANM Harris; MelissaAACI OrlandoAAST FLAACO USAAGP Harris; Melissa Orlando FL US - In a gas turbine engine, a flow directing member includes a platform supported on a rotor and includes a radially facing endwall and at least one axially facing axial surface extending radially inwardly from a junction with the endwall. The flow directing member further includes an airfoil extending radially outwardly from the endwall and a fluid flow directing feature. The fluid flow directing feature includes a groove extending axially into the axial surface. The groove has a radially inner groove end and a radially outer groove end, wherein the outer groove end defines an axially extending notch in the junction between the axial surface and the endwall and forms an opening in the endwall for directing a cooling fluid to the endwall. | 01-17-2013 |
20130031914 | TWO STAGE SERIAL IMPINGEMENT COOLING FOR ISOGRID STRUCTURES - A system for cooling a wall ( | 02-07-2013 |
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 |
20130045111 | TURBINE BLADE COOLING SYSTEM WITH BIFURCATED MID-CHORD COOLING CHAMBER - A cooling system for a turbine blade of a turbine engine having a bifurcated mid-chord cooling chamber for reducing the temperature of the blade. The bifurcated mid-chord cooling chamber may be formed from a pressure side serpentine cooling channel and a suction side serpentine cooling channel with cooling fluids passing through the pressure side serpentine cooling channel in a direction from the trailing edge toward the leading edge and in an opposite direction through the suction side serpentine cooling channel. The pressure side and suction side serpentine cooling channels may flow counter to each other, thereby yielding a more uniform temperature distribution than conventional serpentine cooling channels. | 02-21-2013 |
20130058756 | FLOW DISCOURAGER INTEGRATED TURBINE INTER-STAGE U-RING - A gas turbine having rotor discs ( | 03-07-2013 |
20130064681 | TRAILING EDGE COOLING SYSTEM IN A TURBINE AIRFOIL ASSEMBLY - An airfoil in a gas turbine engine includes an outer wall, a cooling fluid cavity, and a plurality of cooling fluid passages. The outer wall has a leading edge, a trailing edge, a pressure side, a suction side, and radially inner and outer ends. The cooling fluid cavity is defined in the outer wall, extends generally radially between the inner and outer ends of the outer wall, and receives cooling fluid for cooling the outer wall. The cooling fluid passages are in fluid communication with the cooling fluid cavity and include zigzagged passages that include alternating angled sections, each section having both a radial component and a chordal component. The cooling fluid passages extend from the cooling fluid cavity toward the trailing edge of the outer wall and receive cooling fluid from the cooling fluid cavity for cooling the outer wall near the trailing edge. | 03-14-2013 |
20130089415 | GAS TURBINE WITH OPTIMIZED AIRFOIL ELEMENT ANGLES - A turbine airfoil assembly for installation in a gas turbine engine. The airfoil assembly includes an endwall and an airfoil extending radially outwardly from the endwall. The airfoil includes pressure and suction sidewalls defining chordally spaced apart leading and trailing edges of the airfoil. An airfoil mean line is defined located centrally between the pressure and suction sidewalls. An angle between the mean line and a line parallel to the engine axis at the leading and trailing edges defines gas flow entry angles, α, and exit angles, β. Airfoil inlet and exit angles are substantially in accordance with pairs of inlet angle values, α, and exit angle values, β, set forth in one of Tables 1, 3, 5 and 7. | 04-11-2013 |
20130142666 | TURBINE BLADE INCORPORATING TRAILING EDGE COOLING DESIGN - A turbine blade ( | 06-06-2013 |
20130149120 | GAS TURBINE ENGINE WITH OUTER CASE AMBIENT EXTERNAL COOLING SYSTEM - A thermal barrier/cooling system for controlling a temperature of an outer case of a gas turbine engine. The thermal barrier/cooling system includes an internal insulating layer supported on an inner case surface, the internal insulating layer extending circumferentially along the inner case surface and providing a thermal resistance to radiated energy from structure located radially inwardly from the outer case. The thermal barrier/cooling system further includes a convective cooling channel defined by a panel structure located in radially spaced relation to an outer case surface of the outer case and extending around the circumference of the outer case surface. The convective cooling channel forms a flow path for an ambient air flow cooling the outer case surface. | 06-13-2013 |
20130149169 | COMPONENT HAVING COOLING CHANNEL WITH HOURGLASS CROSS SECTION - A cooling channel ( | 06-13-2013 |
20130156579 | AMBIENT AIR COOLING ARRANGEMENT HAVING A PRE-SWIRLER FOR GAS TURBINE ENGINE BLADE COOLING - A gas turbine engine including: an ambient-air cooling circuit ( | 06-20-2013 |
20130186585 | COMPOSITE CORE DIE, METHODS OF MANUFACTURE THEREOF AND ARTICLES MANUFACTURED THEREFROM - A composite core die includes a reusable core die; and a disposable core die. The disposable core die is in physical communication with the reusable core die and surfaces of communication between the disposable core die and the reusable core die serve as barriers to prevent the leakage of a slurry that is disposed in the composite core die. | 07-25-2013 |
20130236330 | TURBINE AIRFOIL WITH AN INTERNAL COOLING SYSTEM HAVING VORTEX FORMING TURBULATORS - A turbine airfoil usable in a turbine engine and having at least one cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels having a plurality of turbulators protruding from an inner surface and positioned generally nonorthogonal and nonparallel to a longitudinal axis of the airfoil cooling channel. The configuration of turbulators may create a higher internal convective cooling potential for the blade cooling passage, thereby generating a high rate of internal convective heat transfer and attendant improvement in overall cooling performance. This translates into a reduction in cooling fluid demand and better turbine performance. | 09-12-2013 |
20130294898 | TURBINE ENGINE COMPONENT WALL HAVING BRANCHED COOLING PASSAGES - A component wall in a turbine engine includes a substrate and at least one cooling passage that extends through the substrate for delivering cooling fluid from a chamber associated with an inner surface of the substrate to an outer surface of the substrate. Each cooling passage is divided into at least two branches that receive cooling fluid from an entrance portion of the cooling passage that is in communication with the chamber. The branches each include an intermediate portion that extends transversely from the entrance portion and that receives cooling fluid from the entrance portion, and an exit portion that extends transversely from the respective intermediate portion. The exit portions receive the cooling fluid from the intermediate portions and deliver the cooling fluid out of the respective branch through exit portion outlets. | 11-07-2013 |
20130298400 | METHOD OF PROVIDING A TURBINE BLADE TIP REPAIR - A method of repairing a turbine blade having a radially extending outer wall defining an internal cavity width and a blade tip. The method comprises removing at least a portion of the blade tip to form a repair surface and providing a tip cap having a radially outer side with an outer width that may be less than the internal cavity width, and having a radially inner side with an inner width that is substantially equal to or greater than the internal cavity width. The tip cap is positioned at the repair surface, and the tip cap is welded to the repair surface using a ductile welding material. A cap peripheral portion is formed by build-up welding around the tip cap, and a squealer portion is formed by build-up welding on the cap peripheral portion. | 11-14-2013 |
20130302166 | TURBINE BLADE WITH CHAMFERED SQUEALER TIP FORMED FROM MULTIPLE COMPONENTS AND CONVECTIVE COOLING HOLES - A squealer tip usable in repair systems and formed from a pressure side outer weld rib and a suction side outer weld rib extending radially outward from a tip of the turbine blade and resting upon pressure side and suction side weld members separated by a mid-chord member is disclosed. The pressure and suction side outer weld ribs may be positioned along the pressure side and the suction side of the turbine blade, respectively. The pressure side outer weld rib may include a chamfered pressure side with film cooling holes having exhaust outlets positioned therein. The pressure and suction side weld members may be configured to retain the mid-chord member in position with over extending side surfaces. | 11-14-2013 |
20130302167 | Near-Wall Serpentine Cooled Turbine Airfoil - Certain exemplary embodiments can provide a serpentine coolant flow path formed by inner walls in a cavity between pressure and suction side walls of a turbine airfoil and/or can be adapted to provide cooling matched to the heating topography of the airfoil, minimize differential thermal expansion, revive the coolant, and/or minimize the flow volume needed. | 11-14-2013 |
20140003919 | FINNED SEAL ASSEMBLY FOR GAS TURBINE ENGINES | 01-02-2014 |
20140037461 | COOLING SYSTEM IN A TURBINE AIRFOIL ASSEMBLY INCLUDING ZIGZAG COOLING PASSAGES INTERCONNECTED WITH RADIAL PASSAGEWAYS - An airfoil in a gas turbine engine includes an outer wall, a cooling fluid cavity, a plurality of cooling fluid passages, and a plurality of radial passageways. The outer wall has leading and trailing edges, pressure and suction sides, and radially inner and outer ends. The cooling fluid cavity is defined in the outer wall and receives cooling fluid for cooling the outer wall. The cooling fluid passages are in fluid communication with the cooling fluid cavity and include alternating angled sections, each section having both a radial component and a chordal component. The cooling fluid passages extend from the cooling fluid cavity toward the trailing edge of the outer wall and receive cooling fluid from the cooling fluid cavity for cooling the outer wall near the trailing edge. The radial passageways interconnect radially adjacent cooling fluid passages. | 02-06-2014 |
20140060068 | METHOD FOR OPERATING A GAS TURBINE ENGINE INCLUDING A COMBUSTOR SHELL AIR RECIRCULATION SYSTEM - During full load operation of gas turbine engine operation, a valve system is maintained in a closed position to substantially prevent air from passing through a piping system of a shell air recirculation system. Upon initiation of a turn down operation, which is implemented to transition the engine to a turning gear state or a shut down state, the valve system is opened to allow air to pass through the piping system. A blower is operated to extract air through at least one outlet port of the shell air recirculation system from an interior volume of an engine casing portion associated with the combustion section, to convey the extracted air through the piping system, and to inject the air into the interior volume of the engine casing portion through at least one inlet port of the shell air recirculation system to circulate air within the engine casing portion. | 03-06-2014 |
20140060082 | COMBUSTOR SHELL AIR RECIRCULATION SYSTEM IN A GAS TURBINE ENGINE - A shell air recirculation system for use in a gas turbine engine includes one or more outlet ports located at a bottom wall section of an engine casing wall and one or more inlet ports located at a top wall section of the engine casing wall. The system further includes a piping system that provides fluid communication between the outlet port(s) and the inlet port(s), a blower for extracting air from a combustor shell through the outlet port(s) and for conveying the extracted air to the inlet port(s), and a valve system for selectively allowing and preventing air from passing through the piping system. The system operates during less than full load operation of the engine to circulate air within the combustor shell but is not operational during full load operation of the engine. | 03-06-2014 |
20140061476 | INFRARED NON-DESTRUCTIVE EVALUATION METHOD AND APPARATUS - A method of nondestructive evaluation and related system. The method includes arranging a test piece ( | 03-06-2014 |
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 |
20140109577 | DUCTING ARRANGEMENT FOR COOLING A GAS TURBINE STRUCTURE - A ducting arrangement ( | 04-24-2014 |
20140110559 | CASTING CORE FOR A COOLING ARRANGEMENT FOR A GAS TURBINE COMPONENT - A ceramic casting core, including: a plurality of rows ( | 04-24-2014 |
20140112799 | COOLING ARRANGEMENT FOR A GAS TURBINE COMPONENT - A cooling arrangement ( | 04-24-2014 |
20140123657 | EXTERNAL COOLING FLUID INJECTION SYSTEM IN A GAS TURBINE ENGINE - A cooling fluid air injection system for use in a gas turbine engine includes at an external cooling fluid source, at least one rotor cooling pipe, which is used to inject cooling fluid from the source into a rotor chamber, a piping system that provides fluid communication between the source and the rotor cooling pipe(s), a blower system for conveying the cooling fluid through the piping system and the rotor cooling pipe(s) into the rotor chamber, and a valve system. The valve system is closed during full load engine operation to prevent cooling fluid from the source from passing through the piping system, and open during less than full load engine operation to allow cooling fluid from the source to pass through the piping system. | 05-08-2014 |
20140123675 | AIR INJECTION SYSTEM IN A GAS TURBINE ENGINE - An air injection system for use in a gas turbine engine includes at least one outlet port through which air is extracted from the engine only during less than full load operation, at least one rotor cooling pipe, which is used to inject the air extracted from the outlet port(s) into a rotor chamber, a piping system that provides fluid communication between the one outlet port(s) and the rotor cooling pipe(s), a blower system for extracting air from the engine through the outlet port(s) and for conveying the extracted air through the piping system and the rotor cooling pipe(s) into the rotor chamber, and a valve system. The valve system is closed during full load engine operation to prevent air from passing through the piping system, and open during less than full load engine operation to allow air to pass through the piping system. | 05-08-2014 |
20140147250 | TURBINE BLADE ANGEL WING WITH PUMPING FEATURES - A gas turbine engine, including: a plurality of blades ( | 05-29-2014 |
20140169962 | TURBINE BLADE WITH INTEGRATED SERPENTINE AND AXIAL TIP COOLING CIRCUITS - An air cooled turbine blade including leading and trailing edges, and pressure and suction side walls extending between the leading and trailing edges. Leading and trailing edge cooling circuits extend spanwise adjacent to the leading and trailing edges, respectively. A forward flow mid-section serpentine cooling circuit extends spanwise and is located between the leading and trailing edge cooling circuits. An axial tip cooling circuit extends in the chordal direction and is located between a tip cap of the blade and the serpentine cooling circuit at an outer end of the serpentine cooling circuit. The axial tip cooling circuit has a forward end receiving cooling air from a final channel of the serpentine cooling circuit and discharges the cooling air adjacent to the trailing edge. | 06-19-2014 |
20140205441 | SEAL ASSEMBLY INCLUDING GROOVES IN A RADIALLY OUTWARDLY FACING SIDE OF A PLATFORM IN A GAS TURBINE ENGINE - A seal assembly between a disc cavity and a turbine section hot gas path includes a stationary vane assembly and a rotating blade assembly downstream from the vane assembly and including a plurality of blades that are supported on a platform and rotate with a turbine rotor and the platform during operation of the engine. The platform includes a radially outwardly facing first surface, a radially inwardly facing second surface, a third surface, and a plurality of grooves extending into the third surface. The grooves are arranged such that a space is defined between adjacent grooves. During operation of the engine, the grooves guide purge air out of the disc cavity toward the hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path. | 07-24-2014 |
20140205443 | SEAL ASSEMBLY INCLUDING GROOVES IN AN AFT FACING SIDE OF A PLATFORM IN A GAS TURBINE ENGINE - A seal assembly between a disc cavity and a hot gas path in a gas turbine engine includes a stationary vane assembly and a rotating blade assembly axially upstream from the vane assembly. A platform of the blade assembly has a radially outwardly facing first surface, an axially downstream facing second surface defining an aft plane, and a plurality of grooves extending into the second surface such that the grooves are recessed from the aft plane The grooves are arranged such that a circumferential space is defined between adjacent grooves During operation of the engine, the grooves impart a circumferential velocity component to purge air flowing out of a disc cavity through the grooves to guide the purge air toward a hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path. | 07-24-2014 |
20140219809 | CASTING CORE FOR TWISTED GAS TURBINE ENGINE AIRFOIL HAVING A TWISTED RIB - A casting core ( | 08-07-2014 |
20140219811 | TWISTED GAS TURBINE ENGINE AIRFOIL HAVING A TWISTED RIB - A gas turbine engine blade ( | 08-07-2014 |
20140219818 | Turbine Component Cooling Channel Mesh with Intersection Chambers | 08-07-2014 |
20140234076 | OUTER RIM SEAL ASSEMBLY IN A TURBINE ENGINE - A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages effect a pumping of cooling fluid from the disc cavity toward the hot gas path during operation of the engine. | 08-21-2014 |
20140271103 | VANE CARRIER THERMAL MANAGEMENT ARRANGEMENT AND METHOD FOR CLEARANCE CONTROL - A thermal management arrangement ( | 09-18-2014 |
20140286760 | SEAL ASSEMBLY INCLUDING GROOVES IN AN INNER SHROUD IN A GAS TURBINE ENGINE - A seal assembly between a disc cavity and a hot gas path in a gas turbine engine includes a rotating blade assembly having a plurality of blades that rotate with a turbine rotor during operation of the engine, and a stationary vane assembly having a plurality of vanes and an inner shroud. The inner shroud includes a radially outwardly facing first surface, a radially inwardly facing second surface, and a plurality of grooves extending into the second surface. The grooves are arranged such that a space having a component in a circumferential direction is defined between adjacent grooves. During operation of the engine, the grooves guide purge air out of the disc cavity toward the hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path. | 09-25-2014 |
20140286791 | Component Cooling Channel - A cooling channel ( | 09-25-2014 |
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 |
20140338866 | COOLING PASSAGE INCLUDING TURBULATOR SYSTEM IN A TURBINE ENGINE COMPONENT - A cooling passage defined between first and second spaced apart sidewalls of a turbine engine component includes a turbulator system including a plurality of rows of turbulator members. Each row includes a first side turbulator member extending from the first sidewall, and a second side turbulator member extending from the second sidewall. The first and second side turbulator members are arranged such that a space is defined therebetween. The first and second side turbulator members are staggered with respect to one another such that respective forward and aft ends thereof are offset from one another. Each row further includes at least one elongate intermediate turbulator member located at least partially in the space between the respective first and second side turbulator members. | 11-20-2014 |
20150033697 | REGENERATIVELY COOLED TRANSITION DUCT WITH TRANSVERSELY BUFFERED IMPINGEMENT NOZZLES - A cooling arrangement ( | 02-05-2015 |
20150071763 | OUTER RIM SEAL ASSEMBLY IN A TURBINE ENGINE - A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly axially adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages each include a portion that is curved as the passage extends radially outwardly to effect a scooping of cooling fluid from the disc cavity into the flow passages and toward the hot gas path. | 03-12-2015 |
20150075746 | Method of Casting a Component Having Interior Passageways - A method of casting a component ( | 03-19-2015 |
20150078898 | Compound Cooling Flow Turbulator for Turbine Component - Multi-scale turbulation features, including first turbulators ( | 03-19-2015 |