CIRCULAR PROPULSION JET COMPRESSOR-ENGINE

Heat energy recapture and recycle and its new applications

CIRCULAR PROPULSION JET COMPRESSOR-ENGINE

CIRCULAR PROPULSION JET COMPRESSOR-ENGINE

Heat energy recapture and recycle and its new applications

HALF OF A HOLLOW FAN BLADE, HOLLOW FAN BLADE FOR GAS-TURBINE ENGINE, AS WELL AS APPROPRIATE MANUFACTURING PROCESS

PROCEDURE FOR THE PRODUCTION OF A SHOVEL FOR A GAS TURBINE COMPONENT AND PROCEDURE FOR THE PRODUCTION OF A GAS TURBINE COMPONENT

SHOVEL FOR A FLUID-FLOW MACHINE

TURBINE OF A GAS TURBINE

COOLING CIRCUIT TO WOULD DRIVE OFF FROM WASTE HEAT OF AN ELECTROMECHANICAL TRANSDUCER AND POWER PLANT WITH A SUCH COOLING CIRCUIT

METHOD AROUND THE THICKNESS OF THE SIDE PANEL OF TURBINE CONTROL DEVICE SEGMENTS TO ADAPT AROUND THE COOLING TO IMPROVING

COOLING OF A TURBINE CONTROL DEVICE SEGMENT

COOLINGCASH SEGMENT FOR TURBOMACHINES AND TURBINES WITH COMBUSTION CHAMBER

System and method for oxidant compression in a stoichiometric exhaust gas recirculation gas turbine system

SYSTEM AND METHOD FOR OXIDANT COMPRESSION IN A STOICHIOMETRIC EXHAUST GAS RECIRCULATION GAS TURBINE SYSTEM [00212] A system includes a gas turbine system having a turbine combustor, a turbine driven by combustion products from the turbine combustor, and an exhaust gas compressor driven by the turbine. The exhaust gas compressor is configured to compress and supply an exhaust gas to the turbine combustor. The gas turbine system also has an exhaust gas recirculation (EGR) system. The EGR system is configured to recirculate the exhaust gas along an exhaust recirculation path from the turbine to the exhaust gas compressor. The system further includes a main oxidant compression system having one or more oxidant compressors. The one or more oxidant compressors are separate from the exhaust gas compressor, and the one or more oxidant compressors are configured to supply all compressed oxidant utilized by the turbine combustor in generating the combustion products. + 5/24 I-: C'", D ( Cn L U O ...

OFFSET CORIOLIS TURBULATOR BLADE

A turbine rotor blade (10) includes an airfoil (12) having pressure and suction sidewalls (26,28) extending longitudinally in span from root (30) to tip (32). The sidewalls (26,28) are spaced apart between leading and trailing edges (34,36) and joined together by longitudinal partitions (38) defining flow channels (1-8) therein. Rows of first and second slant turbulators (40-46) extend from one of the sidewalls in one of the channels and are offset longitudinally. The first and second turbulators (40-46) overlap chordally to eliminate an axial gap therebetween while maintaining a radial gap for tripping cooling air channeled along the span of the channel.

PLASMA LIFTED BOUNDARY LAYER GAS TURBINE ENGINE VANE

A plasma boundary layer lifting system (11) includes at least one gas turbine engine vane (32) having a spanwise extending airfoil (39) with an outer surface (54) extending in a chordwise direction (C) between opposite leading and trailing edges (LE, TE) and chordwise spaced apart plasma generators (2) for producing a plasma (90) extending in the chordwise direction (C) along the: outer surface (54). Each plasma generator (2) may include inner and outer electrodes (3, 4) separated by a dielectric material (5) disposed within a spanwise extending groove (6) in the outer surface (54). The airfoil (39) may be hollow having an outer wall (26) and the plasma generators (2) being mounted on the outer wall (26). A method for operating the system (11) includes forming a plasma (90) extending in the chordwise direction (C) along the outer surface (54) of the airfoil (39). The method may further include operating the plasma generators (2) in steady state or unsteady modes.

TRIFORIAL TIP CAVITY AIRFOIL

BOLTED JOINT FOR ROTOR DISKS AND METHOD OF REDUCING THERMAL GRADIENTS THEREIN

Rotor disk stress is reduced in a bolted joint (52) for connecting adjacent rotor disks in a gas turbine engine (10). The bolted joint (52) includes a bolt hole (70) formed in the first rotor disk (40) and a tube (82) disposed in the bolt hole (70) such that a channel (84) is defined between the tube (82) and the bolt hole (70). A bolt (66) is disposed in the tube (82) such that a gap (88) is defined between the bolt (66) and the tube (82). The gap (88) thermally insulates the bolt (66) from hot fluid in the channel (84). A first passage (96) provides fluid communication between the channel (84) and a forward cavity, and a second passage (98) provides fluid communication between the channel (84) and an aft cavity. Hot fluid passing through the channel (84) reduces thermal gradients in the first rotor disk (40). The tube (82) thermally shields the bolt (66) from the hot fluid to minimize differential thermal growth.

TURBINE BLADE WITH IMPROVED COOLING CHARACTERISTICS AND USEFUL LIFE

... ²²²²La présente invention concerne le domaine des aubes de turbine de ²turbomachine, notamment une aube de turbine (1), comportant une paroi ²intrados (2), une paroi extrados (3), au moins une première cavité radiale ²de bord de fuite (4), au moins une seconde cavité radiale (5) en amont de ²la cavité de bord de fuite (4), une paroi interne (6) séparant les cavités ²radiales (4 et 5) et comprenant au moins un canal (7) reliant les cavités (4 ²et 5) entre elles, ledit canal (7) étant orienté selon un axe (71) coupant la ²surface interne (42) de la paroi intrados (2).² ...

TURBINE VANE WITH DUSTING HOLE AT THE BASE OF THE BLADE

Aube de turbine refroidie pour turbomachine comprenant une pale (2) montée sur une plate-forme (6) portée par un pied (5), ladite pale étant creusée d'une ou plusieurs cavités pour la circulation d'air de refroidissement, la cavité (11) s'étendant le long du bord de fuite étant alimentée en air de refroidissement par un conduit d'alimentation (10) reliant une entrée d'air (12) située en partie basse du pied (5) à la cavité (11) de bord de fuite en faisant un coude (13) au sein dudit pied, caractérisée en ce que le conduit (10) comporte, sur un axe sensiblement radial par rapport à l'entrée d'air (12), une niche (14) située sous la plate- forme (6) ayant une forme en cloche, ladite niche débouchant à son sommet par un trou de dépoussiérage (19) traversant ladite plate-forme et étant délimitée à l'intérieur du pied (5) par des parois s'étendant sensiblement radialement à partir de la plate-forme (6) pour la refermer latéralement.

WET COMPRESSION APPARATUS AND METHOD

... ²²²This wet compression invention with a vaporizable fluid mist demonstrates ²major performance improvements over the relevant art in achieving a high ²degree of saturation, providing sensible cooling, strongly reducing the ²temperature increase due to compression work, reducing excess diluent air flow ²for downstream combustion, reducing compression noise, and increasing the ²achievable compressor pressure ratio . These improvements are obtained by one ²or more of: high mist or overspray from a) progressive axial injection of ²vaporizable fluid along the streamwise compression flow path, and b) ²transverse vaporizable fluid delivery from stators, rotors, perforated tubes, ²and/or duct walls, matching the gaseous fluid flow distribution across the ²compressor stream; c) reducing the compressor cross-sectional flow area of ²downstream compressor stages relative to up-stream stages, and d) increasing ²the rate of downstream vaporizable fluid injection relative to the rate of ²upstream injection ...

AIRFOIL LEADING EDGE END WALL VORTEX REDUCING PLASMA

A leading edge vortex reducing system (11) includes a gas turbine engine airfoil (39) extending in a spanwise direction (S) away from an end wall (88), one or more plasma generators (2) extending in the spanwise direction (S) through a fillet (34) between the airfoil (39) and the end wall (88) in a leading edge region (89) near and around a leading edge (LE) of the airfoil (39) and near the fillet (34). The plasma generators (2) being operable for producing a plasma (90) extending over a portion of the fillet (34) in the leading edge region (89). The plasma generators (2) may be mounted on an outer wall (26) of the airfoil (39) with a first portion of the plasma generators (2) on a pressure side (46) of the airfoil (39) and a second portion of the plasma generators (2) on a suction side (48) of the airfoil (39). A method for operating the system (11) includes energizing one or more plasma generators (2) to form the plasma (90) in steady state or unsteady modes.

COOLING CIRCUITS FOR GAS TURBINE BLADE

Aube (1) de turbine à gaz d'un moteur d'avion, comportant dans sa partie centrale au moins un premier circuit de refroidissement central (A) comprenant au moins une première (2) et une deuxième (4) cavités s'étendant radialement du côté intrados (1a) de l'aube (1), au moins une cavité (6) s'étendant du côté extrados (1b) de l'aube, une ouverture d'admission d'air à une extrémité radiale de la première cavité intrados (2) pour alimenter le premier circuit de refroidissement central (A) en air de refroidissement, un premier passage faisant communiquer l'autre extrémité radiale de la première cavité intrados (2) à une extrémité radiale voisine de la cavité extrados (6), un second passage faisant communiquer l'autre extrémité radiale de la cavité extrados avec une extrémité radiale voisine de la deuxième cavité intrados (4), et des orifices de sortie s'ouvrant dans la deuxième cavité intrados et débouchant sur la face intrados (1a) de l'aube.

TURBINE ENGINE FOR EXAMPLE A GAS TURBINE ENGINE FOR AN AIRCRAFT

Turbomachine comprenant une roue à aubes creuses refroidies intérieurement par circulation forcée d'air. Selon une caractéristique importante de l'invention, un flasque- crochet annulaire (70) est intercalé entre le bord extérieur du disque- labyrinthe (52) et le disque de rotor (36), le flasque-crochet comportant un épaulement de retenue (72) formant crochet annulaire extérieur dans lequel s'encastre le bord extérieur du disque-labyrinthe et un épaulement d'appui (74) formant crochet intérieur, engagé dans un évidemment circonférentiel du disque de rotor.

METHOD OF FABRICATING LIQUID COOLED GAS TURBINE COMPONENTS

GAS TURBINE COOLED SHROUD ASSEMBLY WITH HOT GAS INGESTION SUPPRESSION

A cooled shroud assembly (19) includes an angled slot (25) and a plurality of dilution jet openings (26). The shroud forward cavity is modified such that at least one recirculation zone (21) is produced. The angled slot (25) forces; an axial change in momentum of the hot gas flow (37) and increases radial and axial pressure variation attenuation. The cooled shroud assembly (19) isolates the shroud structure and seals from the hot flow path and a cooling flow (42) from the dilution jet openings (26) dilutes the hot gas flow. A series of recirculation zones shields the shroud carrier (33) and high pressure seals from the hot gas flow.

GAS TURBINE ENGINE COOLING SYSTEM AND METHOD

Fuel (12) supplied to a rotary fluid trap (42) is centrifugally accelerated within a first cavity (46) adjacent a first side (48) of a rotor (24), and is then directed though a plurality of first passages (66) extending through the rotor (24) between and proximate to the blades (26), and shaped so as to at least partially conform to the shape of the blades (26). Second passages (100) extend within the blades (26) from the first passages (66) and terminate within associated cavities (110) proximate to the tips (112) of the blades (26). Relatively cooler fuel (12.2) in the first passages (66) is thermosiphon exchanged for relatively hotter fuel (12.3) in the second passages (100) so as to cool the blades (26). The heated fuel (12.3) flows into a second cavity (74) adjacent to a second side (72) of the rotor (24) and is discharged from the rotating frame of reference directly into the combustion chamber (16) through a second rotary fluid trap (96). A separate fuel distribution circuit (130 ...

VENTILATION SYSTEM FOR COMBUSTION CHAMBER WALL

Système de ventilation de paroi de chambre de combustion dans une turbomachine comprenant un compresseur centrifuge (10) alimentant par un diffuseur (12) une chambre de combustion (14), ce système comportant un caisson annulaire (90) agencé radialement entre la chambre de combustion et un flasque aval (26) du diffuseur et qui comporte une paroi radialement externe (92) pour le guidage d'un flux d'air sortant du diffuseur, et une paroi radialement interne (94) délimitant avec le flasque du diffuseur une voie (88) de passage d'air sortant du diffuseur.

COOLING SYSTEM FOR A CENTRIFUGAL COMPRESSOR IMPELLER

Système de refroidissement du rouet d'un compresseur centrifuge dans une turbomachine, ce compresseur (10) alimentant un diffuseur annulaire (12) comportant un flasque (26) qui s'étend en aval et le long du rouet du compresseur et qui est recouvert du côté amont par une tôle annulaire (90) qui délimite, avec le rouet du compresseur, un premier passage annulaire (39) d'écoulement d'air prélevé en sortie du compresseur et, avec le flaque du diffuseur, un second passage annulaire (98) d'écoulement d'une partie du débit d'air sortant du diffuseur.

PROCESS FOR DRILLING HOLES IN A METALLIC WORKPIECE HAVING A THERMAL BARRIER COATING

A method is provided for drilling a hole through a metallic workpiece (1) having a thermal barrier coating (3) with a ceramic top coat (2) by laser drilling a counterbore to a depth which extends through the ceramic top coat but not substantially into the metallic workpiece and then laser drilling the hole through the workpiece aligned with the counterbore, the counterbore having a diameter larger than the hole.

TURBINE ENGINE WITH AIR COOLED TURBINE

In a turbine engine (10), low temperature air is diverted from a low pressure section of the compressor section of the engine to cool the high pressure turbine of the engine. Low pressure air is diverted from the compressor section, and its pressure is thereafter be increased. the pressure is increased in an intermediate cavity (80) in the engine, where rotational energy of the diverted air is converted to static pressure by way of an obstruction (88) that converts dynamic head of the air in the cavity into static pressure.

COOLABLE WALL

METHODS AND APPARATUS FOR COOLING GAS TURBINE ENGINE NOZZLE ASSEMBLIES

A method for fabricating a nozzle (51) for a gas turbine engine (10) facilitates extending a useful life of the nozzles. The nozzle includes an airfoil (52). The method includes forming the airfoil to include a suction side (60) and a pressure side (62) connected at a leading edge (64) and a trailing edge (66) such that a cooling cavity (82) and a cooling circuit (80) are defined within the airfoil, wherein the suction side and the pressure side extend radially between a tip (72) and a root (70). The method also includes forming a plurality of cooling slots (96) within the airfoil that extend from the cooling circuit towards the airfoil trailing edge, and forming a control vane (110) within the cooling circuit to facilitate maintaining a substantially constant cooling effectiveness within the cooling circuit.

PROCESS FOR DRILLING HOLES IN A METALLIC WORKPIECE HAVING A THERMAL BARRIER COATING

A method is provided for drilling a hole through a metallic workpiece (1) having a thermal barrier coating (3) with a ceramic top coat (2) by laser drilling a counterbore to a depth which extends through the ceramic top coat but not substantially into the metallic workpiece and then laser drilling the hole through the workpiece aligned with the counterbore, the counterbore having a diameter larger than the hole.

TURBINE ENGINE FOR EXAMPLE A GAS TURBINE ENGINE FOR AN AIRCRAFT

Turbomachine comprenant une roue à aubes creuses refroidies intérieurement par circulation forcée d'air. Selon une caractéristique importante de l'invention, un flasque- crochet annulaire (70) est intercalé entre le bord extérieur du disque- labyrinthe (52) et le disque de rotor (36), le flasque-crochet comportant un épaulement de retenue (72) formant crochet annulaire extérieur dans lequel s'encastre le bord extérieur du disque-labyrinthe et un épaulement d'appui (74) formant crochet intérieur, engagé dans un évidemment circonférentiel du disque de rotor.

LIGHTER TURBOMACHINE BLADE AND PROCESS FOR MANUFACTURING THE SAME

L'invention propose une aube allégée de turbomachine (10) comportant une pale (40) en alliage métallique, cette pale (40) comportant elle-même une cavité (70) obturée par un couvercle (80) sur l'un des deux flancs (50) appelé flanc creusé (50a), ce couvercle (80) assurant la continuité aérodynamique du flanc (50a), ce couvercle (80) étant lié par le bord (85) au reste de la pale (40) par un cordon de soudure (100). Une telle aube est remarquable en ce que le cordon de soudure (100) débouche sur le flanc creusé (50a) et pénètre dans la pale (40) avec une profondeur P au moins égale à l'épaisseur EC du bord du couvercle (85) afin d'assurer la continuité de la matière entre le bord du couvercle (85) et le reste de la pale (40) sur une profondeur au moins égale à l'épaisseur EC du bord du couvercle (85). L'invention propose également un procédé de fabrication d'une telle aube (10).

COOLING SYSTEM FOR A DOWNSTREAM CAVITY IN A CENTRIFUGAL COMPRESSOR IMPELLER

... ²²²Système de refroidissement d'une cavité aval de rouet de ²compresseur centrifuge dans une turbomachine, ce compresseur (10) étant ²raccordé à un diffuseur (12) comportant un flasque annulaire aval, une tôle ²annulaire (70) étant montée coaxialement autour du flasque du diffuseur et ²s'étendant sensiblement sur toute la dimension axiale du flasque pour ²définir un passage annulaire (72) de ventilation alimenté en air qui est plus ²frais que l'air (61) sortant du compresseur centrifuge.² ...

TURBOMACHINE TURBINE BLADE COMPRISING A COOLING CIRCUIT WITH IMPROVED HOMOGENEITY

Le refroidissement interne des aubes mobiles des turbines dans les turbomachines d'aéronefs est d'une efficacité limitée en raison d'inhomogénéités de ce refroidissement sur chacune des parois d'intrados et d'extrados. Pour remédier à ce problème, il est proposé une aube comprenant un circuit de refroidissement (50) de sa pale (34), dans lequel les cavités interconnectées en série sont telles que le flux d'air circule radialement vers l'extérieur le long de la paroi d'intrados (40) au sein de cavités d'intrados (52, 56), et radialement vers l'intérieur le long de la paroi d'extrados (42) au sein d'une cavité d'extrados (54) séparée des cavités d'intrados par une paroi interne (58) de la pale. Ainsi, la force de Coriolis dévie le flux d'air vers chacune des parois d'intrados et d'extrados, limitant ainsi les inhomogénéités.

TURBINE ASSEMBLY OF AN AIRCRAFT TURBINE ENGINE

La présente invention concerne un ensemble de turbine (10) de turbomachine (1), comprenant au moins un premier rotor aubagé (12), un stator aubagé (13) et un deuxième rotor aubage (14) disposés successivement, les rotors (12, 14) étant montés sur un arbre (2), une platine d'étanchéité (20) s'étendant entre le stator (13) et l'arbre (2) et séparant une première cavité (C1) disposée entre le premier rotor (12) et le stator (13), d'une deuxième cavité (C2) disposée entre le stator (13) et le deuxième rotor (14), des moyens (300, 31) de diminution de la pression étant positionnés au sein de la première cavité (C1), l'ensemble étant caractérisé en ce que lesdits moyens (300, 31) de diminution de la pression comprennent une pluralité d'ailettes de recompression (300) sensiblement radiales s'étendant dans la première cavité (C1).

CASING FOR A PROPULSION ASSEMBLY

Ensemble propulsif d'aéronef, comportant un moteur et une nacelle comprenant un carter(16) de révolution délimitant une veine d'écoulement d'un flux d'air, caractérisé en ce que ce carter comporte au moins deux ouvertures obturées par des panneaux (18) amovibles et interchangeables, au moins l'un de ces panneaux portant un équipement (24) de l'ensemble propulsif.

COOLED VANE OF A HIGH-PRESSURE TURBINE

La présente invention porte sur une aube mobile de turbomachine comportant une pale (12) avec des cavités internes de refroidissement et un pied (11) par lequel l'aube peut être montée sur un disque de rotor, le pied comprenant au moins deux canaux (11c) communiquant avec lesdites cavités internes et débouchant sur sa base (11b), ladite base comprenant au moins deux ouvertures (11b1, 11b2) à travers lesquelles débouchent les canaux, une plaquette de calibrage (20) pourvue de perçages calibrés (21,22 ) correspondant auxdites ouvertures étant fixée sur la base (11b) du pied. L'aube est caractérisée par le fait qu'un moyen mécanique (25, 26; 11m1, 11m2) formant barrière d'étanchéité entre les deux ouvertures est ménagé entre la plaquette et la base du pied.

SYSTEM AND METHOD FOR IMPROVED ENGINE COOLING

... 13DV-10766 THE A method and system for cooling a gas turbine engine. The gas turbine engine has a multi-stage compressor which discharges compressor discharge air for expansion by a combustor. A cavity which is aerodynamically linked to a turbine stage on its downstream end is forced by a metal structure the upstream end of which is connected to a seal which prevents compressor discharge air from entering the cavity. A conduit links an intermediate stage of the multi-stage compressor to the cavity for purposes of cooling the cavity and the turbine stage components located downstream of the cavity.

FILM COOLING OF JET ENGINE COMPONENTS

Patent 13DV-10548 A jet engine component, such as an aircraft gas turbine engine rotor blade or a scramjet engine fuel injector. The component has a wall portion including a first surface exposable to a cooler, higher static pressure fluid and a second surface exposable to a hotter, lower static pressure gas flow flowing across the second surface. The component further includes a film coolant passageway having an inlet on the first surface and an outlet on the second surface. The second surface has an open groove extending from the outlet along the gas flow for improved film cooling of the second surface.

[...][...] PAR jET D'AIR ET PAR tRANSPIRATION.

PRODUCT FROM A SUPERALLOY.

COMPOUND PRODUCT FROM AT LEAST TWO SUPERALLOYS.

Turbine blade with a chamber for receiving a coolant flow.

Eine Turbinenschaufel (16), die ein Schaufelblatt beinhaltet, das von einer konkav geformten druckseitigen Aussenwand (26) und einer konvex geformten saugseitigen Aussenwand (27) definiert wird, die an Vorder- und Hinterkante (28, 29) entlang miteinander verbunden sind und dazwischen eine radial verlaufende Kammer zur Aufnahme eines Kühlmittelstroms bilden. Die Turbinenschaufel beinhaltet ferner eine Rippenanordnung (60, 62, 63, 64, 66, 67, 68, 69), die die Kammer in radial verlaufende Strömungsdurchgänge untergliedert, und eine Schaufelaussenschale, die eine Aussenfläche des Schaufelblatts definiert. Die Rippenanordnung ist ein nichtintegrales Bauteil der Schaufelaussenschale.

Turbine blade with a cooling device.

Turbinenlaufschaufel (14) mit einem Schaufelblatt (25), das eine Kühlungsanordnung beinhaltet, die mehrere längliche Strömungsdurchgänge (43, 44) zum Aufnehmen und Leiten eines Kühlmittels an einem Weg durch das Schaufelblatt (25) entlang hat. Die Kühlungsanordnung beinhaltet Folgendes: einen zentralen Strömungsdurchgang, der an jeder Seite von wandnahen Strömungsdurchgängen (43, 44) flankiert wird, zu denen ein druckseitiger wandnaher Strömungsdurchgang (43) und ein saugseitiger wandnaher Strömungsdurchgang (44) zählen, ein erstes Loch (46), das den mittleren Strömungsdurchgang mit dem druckseitigen wandnahen Strömungsdurchgang (43) in Strömungsverbindung setzt, ein zweites Loch (46), das den mittleren Strömungsdurchgang mit dem saugseitigen wandnahen Strömungsdurchgang (44) in Strömungsverbindung setzt, und Prallverbinder (48), die den mittleren Strömungsdurchgang mit einem Vorderkanten-Strömungsdurchgang (42) in Strömungsverbindung setzen.

Turbine blade with a chamber for receiving a coolant flow.

Eine Turbinenschaufel (25) mit einem Schaufelblatt und einer radial verlaufenden Kammer darin für ein Kühlmittel. Die Turbinenschaufel (25) beinhaltet auch eine Rippenanordnung (60, 62, 63, 64, 66, 67, 68, 69), welche die Kammer in radial verlaufende Strömungsdurchgänge (40) untergliedert. Die Rippenanordnung beinhaltet: eine Skelettlinienrippe (62, 63, 64), die eine wandnahe Strömungskammer definiert, und Querrippen (66, 67, 68, 69), die zwischen der Skelettlinienrippe (62, 63, 64) und einer der Aussenwände verlaufen, um die wandnahe Strömungskammer in nacheinander stapelförmig angeordnete Strömungsdurchgänge (40) zu unterteilen, die jeweils ein Segment der Skelettlinienrippe (62, 63, 64) beinhalten. Das Segment der Skelettlinienrippe (62, 63, 64) von einem der Strömungsdurchgänge (40) hat ein sich verschmälerndes Profil, das ein Profil beinhaltet, das sich von entgegengesetzten Enden zu einem dazwischen angeordneten Hals hin verschmälert.

Turbine blade with a chamber for receiving a coolant flow.

Eine Turbinenschaufel (16), umfassend ein Schaufelblatt, das von einer konkav geformten druckseitigen Aussenwand (26) und einer konvex geformten saugseitigen Aussenwand (27) definiert wird, die an Vorder- und Hinterkante (28, 29) entlang miteinander verbunden sind und dazwischen eine radial verlaufende Kammer zur Aufnahme, eines Kühlmittelstroms bilden. Die Turbinenschaufel (16) beinhaltet ferner Folgendes: eine Rippenanordnung (60, 62, 63, 64, 66, 67, 68, 69), die die Kammer in radial verlaufende Strömungsdurchgänge (40) unterteilt, die einen ersten Strömungsdurchgang und einen zweiten Strömungsdurchgang beinhalten, und einen Verbindungsdurchgang, der einen im ersten Strömungsdurchgang gebildeten Einlass mit einem im zweiten Strömungsdurchgang gebildeten Auslass in Strömungsverbindung setzt. Der Verbindungsdurchgang beinhaltet eine abgeschrägte Anordnung relativ zum zweiten Strömungsdurchgang.

Turbine blade with a chamber for receiving a coolant flow.

Eine Turbinenschaufel (25) mit einem Schaufelblatt, das von Aussenwänden definiert wird, bei denen eine konkav geformte druckseitige Aussenwand (26) und eine konvex geformte saugseitige Aussenwand (27) an einer Vorder- und Hinterkante (28, 29) entlang miteinander verbunden sind und eine Kammer zur Aufnahme eines Kühlmittelstroms bilden. Die Turbinenschaufel (25) beinhaltet eine Rippenanordnung (60, 62, 63, 64, 66, 67, 68, 69), die die Kammer in radial verlaufende Strömungsdurchgänge (40) unterteilt. Die Rippenanordnung beinhaltet eine Rippe mit einem welligen Profil, das durch einen der Strömungsdurchgänge einer Zielfläche gegenüberliegt. Relativ zur Zielfläche beinhaltet das wellige Profil der Rippe einen Rückenteil und einen Furchenteil. Die Rippe beinhaltet durch den Rückenteil ausgebildete Prallöffnungen.

Method and system for cooling of blade- angel wings.

Es wird ein System (202) zum Kühlen eines Engelsflügels (224), der mit einer Laufschaufel (216) in einer Gasturbine (200) verbunden ist, geschaffen. Ein Engelsflügel (224), der mit einem Schaft (220) einer Laufschaufel (216) verbunden ist, weist mindestens einen Kühlkanal (230) auf, der sich von mindestens einer Einlassöffnung (236), die mit einem Radinnenraum (244) der Gasturbine in Strömungsverbindung steht, zu mindestens einer Auslassöffnung (232) erstreckt, die mit einem äusseren Rotor-Stator-Hohlraum der Gasturbine in Strömungsverbindung steht. Die mindestens eine Auslassöffnung (232) ist in einer oberen Oberfläche (234) des Engelsflügels angeordnet. Der mindestens eine Kühlkanal (230) empfängt unter Druck stehende Kühlluft, die aus dem Radinnenraum geleitet wird, so dass die unter Druck stehende Kühlluft in die mindestens eine Einlassöffnung (236) geleitet wird und aus der mindestens einen Auslassöffnung (232) ausgetragen wird.

Turbine blade with a cooling channel, a method for increasing the life of a turbine blade.

Eine Turbinenschaufel (100) weist einen Kühlkanal (128) auf, der zwischen einer Druckseitenwand (120) und einer Saugseitenwand (122) definiert ist. Ein Stift (132) ist in dem Kühlkanal (128) angeordnet und weist ein erstes Ende, das mit der Druckseitenwand (120) verbunden ist, und ein zweites Ende auf, das mit der Saugseitenwand (122) verbunden ist. Eine radial ausgerichtete Hohlkehle, die einen maximalen Krümmungsradius aufweist, ist in einem Bereich maximaler Dauerschwingbeanspruchung entlang eines Umfangs wenigstens eines von dem ersten Ende oder dem zweiten Ende angeordnet. Eine axial ausgerichtete Hohlkehle, die einen maximalen Krümmungsradius aufweist, ist in einem Bereich maximaler Schlagbelastung entlang eines Umfangs wenigstens eines von dem ersten Ende oder dem zweiten Ende angeordnet. Der maximale Krümmungsradius der axial ausgerichteten Hohlkehle ist grösser als der maximale Krümmungsradius der radial ausgerichteten Hohlkehle.

Turbine blade with a chamber for receiving a coolant flow.

Eine Turbinenschaufel (16), die ein Schaufelblatt beinhaltet, das von einer konkav geformten druckseitigen Aussenwand (26) und einer konvex geformten saugseitigen Aussenwand (27) definiert wird, die an Vorder- und Hinterkante (28, 29) entlang miteinander verbunden sind und dazwischen eine radial verlaufende Kammer zur Aufnahme eines Kühlmittelstroms bilden. Die Turbinenschaufel (16) beinhaltet ferner eine Rippenanordnung (60, 62, 63, 64, 66, 67, 68, 69), die die Kammer des Schaufelblatts in radial verlaufende Strömungsdurchgänge (40) untergliedert. Ein erster Strömungsdurchgang beinhaltet eine erste Seite, an der Turbulenzerzeuger positioniert sind, wobei die Turbulenzerzeuger jeweils eine abgeschrägte Anordnung aufweisen.

Turbine blade with a chamber for receiving a coolant flow.

Turbinenschaufel (16) mit einem Schaufelblatt, das von einer konkav geformten druckseitigen Aussenwand (26) und einer konvex geformten saugseitigen Aussenwand (27) definiert wird, die an Vorder- und Hinterkanten (28, 29) entlang miteinander verbunden sind und dazwischen eine radial verlaufende Kammer zur Aufnahme eines Kühlmittelstroms bilden. Die Turbinenschaufel (16) beinhaltet eine Rippenanordnung (60, 62, 63, 64, 66, 67, 68, 49), die die Kammer unterteilt und einen Strömungsdurchgang mit einer ersten Seite und einer zweiten Seite definiert. Der Strömungsdurchgang beinhaltet ein Loch, das durch die erste Seite ausgebildet ist. Eine Projektion einer Mittelachse des Lochs durch den Strömungsdurchgang definiert einen Anprallpunkt auf der zweiten Seite des Strömungsdurchgangs, und auf der zweiten Seite des Strömungsdurchgangs ist eine Rückschlagaussparung zum Fassen des Anprallpunkts positioniert.

Turbine blade with a chamber for receiving the flow of a coolant.

Eine Turbinenschaufel (16) weist ein Profil auf, das durch eine konkav geformte Aussenwand (26) auf der Druckseite und eine konvex geformte Aussenwand (27) auf der Saugseite definiert ist, die entlang einer Vorder- und Hinterkante (28, 29) verbunden sind und dazwischen eine radial verlaufende Kammer zum Aufnehmen des Stroms eines Kühlmittels bilden. Die Turbinenschaufel (16) weist ferner eine Rippenanordnung (60, 62, 63, 64, 66, 67, 68, 69) auf, die die Kammer in radial verlaufende Strömungskanäle (40) unterteilt. Die Rippenanordnung weist eine Wölbungslinienrippe mit einem Wellenprofilquerschnitt auf. Der Wellenprofilquerschnitt kann mindestens eine hin- und hergehende «S»-Form aufweisen.

Heat transfer assembly and method of mounting of the same.

Es wird eine Wärmeübertragungsanordnung (102) zur Wärmeübertragungs-Steuerung einer Turbinenmaschine (100) bereitgestellt. Die Turbinenmaschine (100) enthält ein Gehäuse und enthält einen Verdichter (104) eine Brennkammer (110) und eine Turbine (106) die in dem Gehäuse angeordnet sind. Die Wärmeübertragungsanordnung (102) enthält eine Strömungssteuerungsvorrichtung (160) mit einer mit der Turbine (106) verbundenen Seitenwand (164), wobei die Strömungssteuerungsvorrichtung (160) mit der Verdichterleitschaufel in Strömungsverbindung steht. Die Seitenwand (164) ist dafür eingerichtet, einen ersten Strömungspfad (172) von der Verdichterleitschaufel zu einer Turbinenleitschaufel und einen zweiten Strömungspfad (174) von der Verdichterleitschaufel zu einer Turbinenlaufschaufel zu definieren. Ein Wärmetauscher (162) ist mit dem Gehäuse verbunden und befindet sich zwischen dem Verdichter (104) und der Turbine (106), wobei der Wärmetauscher (162) mit wenigstens einem von dem ersten Strömungspfad ...

System with turbine case arrangement.

Ein System enthält eine Turbinengehäuseanordnung, die eine Aussenschale und eine im Wesentlichen konzentrisch in der Aussenschale positionierte Innenschale (30) enthält. Die Innenschale (30) enthält eine von der Aussenschale weg zeigende Innenoberfläche und eine zur Aussenschale hin zeigende Aussenoberfläche (74) und die Aussenoberfläche (74) hat einen oder mehrere Falschflansche (60). Wenigstens einer von dem einen oder den mehreren Falschflanschen (60) enthält eine aus der Aussenoberfläche (74) hervorstehende und zur Aussenschale hin zeigende erste Oberfläche (76) und einen sich zwischen der ersten Oberfläche (76) und der Aussenoberfläche (74) der Innenschale (70) erstreckenden Strömungsumlenkungsabschnitt (78). Der Strömungsumlenkungsabschnitt (78) enthält einen ersten Abschnitt (82); der in einer ersten Umfangsrichtung (84) zwischen der ersten Oberfläche (76) und der Aussenoberfläche (74) divergiert.

Blades with holes arranged at a shallow angle and method for drilling the same.

Ein Schaufelblatt enthält einen in einer Aussenoberfläche (210) des Schaufelblattes ausgebildeten Diffusorbereich (302). Der Diffusorbereich (302) ist durch wenigstens eine Aussenoberfläche (308) und eine Innenoberfläche (306) definiert, die bei einem Strömungsauslass (320) zusammentreffen. Das Schaufelblatt enthält auch einen in einer Innenoberfläche (212) des Schaufelblattes ausgebildeten Strömungseinlassbereich (304). Der Strömungseinlassbereich (304) enthält einen Strömungseinlass (324). Das Schaufelblatt enthält einen sich durch das Schaufelblatt von dem Diffusorbereich (302) zu dem Strömungseinlassbereich (304) erstreckenden Kühlkanal (322). Der Kühlkanal (322) definiert den Strömungsauslass (320) und den Strömungseinlass (324).

Turbine nozzle segment with cooling.

Ein Turbinenleitschaufelsegment weist eine oder mehrere Leitschaufeln (50) auf, die sich zwischen einer radial inneren und äusseren Seitenwand (46, 48) erstrecken, wobei jede Leitschaufel (50) eine Umfangsrandwand aufweist, die sich zwischen einer Vorderkante (54, 56) und einer Hinterkante (55) der Leitschaufel (50) erstreckt. Erfindungsgemäss ist wenigstens ein im Wesentlichen radial ausgerichteter Kühlkanal in der Umfangsrandwand, der an einer von der inneren und äusseren Seitenwand vorgesehen ist. Die Lage und Länge der Kühlkanäle kann in einer Ausführungsform um die Umfangsrandwand herum variieren, und der innere Hohlraum der Leitschaufel (50) kann mit Rippen versehen sein, die sich entlang des einen oder mehreren Kühlkanäle und zu diesen benachbart erstrecken, um die Wand zu verstärken und auch um zusätzliche Kühloberflächen in dem inneren Hohlraum zu schaffen.

Turbine blade with serpentine core.

Die Erfindung betrifft eine Turbinenschaufel (200), zu der gehören: eine Basis (212); und ein Schaufelblatt (202), das an einem ersten Ende des Schaufelblatts (202) mit der Basis (212) verbunden ist, wobei das Schaufelblatt (202) aufweist: ein Gehäuse (203), zu dem gehören: eine Saugseite (204); eine Druckseite (206), die der Saugseite (204) gegenüberliegt; eine Anströmkante (208), die sich zwischen der Druckseite (206) und der Saugseite (204) erstreckt; und eine Abströmkante (210), die der Anströmkante (208) gegenüberliegt und sich zwischen der Druckseite (206) und der Saugseite (204) erstreckt, wobei das Gehäuse (203) an der Anströmkante (210) eine Öffnung (218) aufweist; und einen Kern in dem Gehäuse (203), wobei der Kern eine Serpentinengestalt zum Stützen des Gehäuses (203) und einen Anströmkantendurchlasskanal aufweist, der mit der Öffnung (218) an der Anströmkante (208) des Gehäuses (203) strömungsmässig verbunden ist.

Gas Turbine Assembly.

Die Erfindung betrifft eine Gasturbinenanordnung (200). Die Gasturbinenanordnung (200) weist ein massives Laufrad (214), eine hintere Turbinenwelle (228) und eine hintere Verbindung (226), die das massive Laufrad (214) mit der hinteren Turbinenwelle (228) verbindet, auf. Die Gasturbinenanordnung (200) enthält ferner eine Abdeckblende (230), die zwischen dem massiven Laufrad (214) und der hinteren Turbinenwelle (228) angeordnet ist. Die Abdeckblende (230) ist eingerichtet, um eine Strömung eines Verdichterentnahmefluids nach innen durch die hintere Verbindung (226) und aus der hinteren Turbinenwelle (228) nach aussen zu leiten.

Impeller means for providing a cooling of a turbine rotor disk.

Eine Laufradanordnung (100) zur Schaffung einer Kühlung eines Turbinenlaufrades (44) enthält ein Laufrad (44), das mehrere Schwalbenschwanznuten (56) enthält, die längs des Umfangs um eine Umfangsfläche des Laufrades (44) beabstandet sind. Jede der Schwalbenschwanznuten (56) enthält ein Paar gegenüberliegender oberer Nutzapfen (58) und ein Paar gegenüberliegender unterer Nutzapfen (58). Die Laufradanordnung (100) enthält ferner wenigstens eine Turbinenschaufel (32). Die Turbinenschaufel (32) enthält ein Schaufelblatt (40), eine Plattform (48) und einen Schwalbenschwanz (46). Der Schwalbenschwanz (46) enthält ein Paar gegenüberliegender oberer Schwalbenschwanzzapfen (54) und ein Paar gegenüberliegender unterer Schwalbenschwanzzapfen (54). Der Schwalbenschwanz (46) enthält ferner wenigstens eine Einlassöffnung (72), die sich in Längsrichtung durch diesen hindurch erstreckt. Das Paar oberer Schwalbenschwanzzapfen (54) enthält ein erstes Kühlloch (74), das sich durch dieses hindurch erstreckt ...

Turbine with a rotor cooling.

In einem Ausführungsbeispiel schafft die Erfindung eine Turbine (100), zu der gehören: ein Rotor (40) mit einem ersten Schaufelfuss (70); und ein Statorglied (10), das aufweist: eine Rotorbohrung (20), in der wenigstens ein Abschnitt des Rotors (40) angeordnet ist; ein Stirnende (12), das benachbart zu dem ersten Schaufelfuss (70) des Rotors (40) angeordnet ist; mehrere Dichtungen (26, 28) in der Rotorbohrung (20), die dazu dienen, gegen den Rotor (40) abzudichten, wobei die mehreren Dichtungen (26, 28) eine erste Dichtung (26), die dem Stirnende (12) am nächsten ist, und eine zweite Dichtung (28) beinhalten, die benachbart zu der ersten Dichtung (26) angeordnet ist; und mehrere Durchlasskanäle (30, 38), wobei sich jeder ausgehend von einer Oberfläche der Rotorbohrung (20) an einer Stelle zwischen der ersten Dichtung (26) und der zweiten Dichtung (28) erstreckt und sich durch das Stirnende (12) erstreckt.

Turbine assembly and methods of providing a purge airflow and an adjustable flow of cooling air to a cavity in a gas turbine.

Die Erfindung betrifft eine Turbinenanordnung und ein Verfahren zur Lieferung einer Spülluftströmung und einer einstellbaren Kühlluftströmung zu einem Laufradhohlraum oder einem Statorhohlraum in einer Gasturbine. Die Turbinenanordnung weist eine Rotoranordnung (118), eine Statoranordnung (120), die benachbart zu der Rotoranordnung (118) positioniert ist, und einen Laufradhohlraum (122), der zwischen der Rotoranordnung (118) und der Statoranordnung (120) ausgebildet ist. Wenigstens eine unveränderliche Spülluftdurchlassöffnung (130) ist der Statoranordnung (120) zugeordnet. Die unveränderliche Spülluftdurchlassöffnung (130) ist konfiguriert, um eine Spülluftströmung (134) zu dem Laufradhohlraum (122) zu liefern. Ausserdem ist wenigstens eine veränderliche Kühlluftdurchlassöffnung (132) der Statoranordnung (120) zugeordnet. Die wenigstens eine veränderliche Kühlluftdurchlassöffnung (132) ist konfiguriert, um eine Kühlluftströmung (136) zu dem Laufradhohlraum (122) zu liefern.

hot gas path construction unit for turbine system.

Es wird ein Heissgaspfad-Bauteil (30) für ein Turbinensystem offengelegt. Das Heissgaspfad-Bauteil (30) enthält eine Schale (32) und eine oder mehrere poröse Stoffanordnungen (60) mit einer Aussenoberfläche (66) und einer Innenoberfläche (68) und an die Schale (32) angrenzend positioniert. Die eine oder die mehreren porösen Stoffanordnungen (60) sind dafür eingerichtet, eine variierende Durchlässigkeit in einer von einer Axialrichtung, einer Radialrichtung, einer Axial- und einer Radialrichtung, einer Axial- und einer Umfangsrichtung, einer Radial- und einer Umfangsrichtung oder einer Axial-, einer Radial- und einer Umfangsrichtung zu enthalten, wobei die porösen Stoffanordnungen (60) an die Schale (32) angrenzend positioniert sind. Die eine oder die mehreren porösen Stoffanordnungen (60) sind ferner dafür eingerichtet, eine von einer Axial-, einer Radial-, einer Axial- und einer Radial-, einer Axial- und einer Umfangs-, einer Radial- und einer Umfangs- oder einer Axial-, einer Radial- ...

Method and apparatus for improving the heat transfer in turbine sections gas turbine engines.

Es ist ein Gasturbinensystem mit einem Verbrennungsabschnitt und einem Turbinenabschnitt geschaffen. Der Turbinenabschnitt weist mindestens eine Turbinenstufe mit mehreren Turbinenschaufeln, die mit einem Rotor gekoppelt sind, und ein Innengehäuse (62) auf, das längs des Umfangs um die mehreren Turbinenschaufeln herum angeordnet ist. Der Turbinenabschnitt weist ein Aussengehäuse (64) auf, das längs des Umfangs um zumindest einen Abschnitt des Innengehäuses (62) herum angeordnet ist. Das Innengehäuse (62) und das Aussengehäuse (64) definieren einen Hohlraum (66), der ein Volumen umfasst, das so ausgelegt ist, dass es die Luftverteilung in dem Hohlraum (66) begünstigt, damit eine Aussenfläche (72) des Innengehäuses (62) und eine Innenfläche (74) des Aussengehäuses (64) gekühlt werden. Das Aussengehäuse (64) umfasst mindestens einen Lufteinlass (80), und das Innengehäuse umfasst mindestens einen Luftauslass (84). In dem Hohlraum (66) ist mindestens ein Bund vorgesehen, und der mindestens eine ...

Gas Turbine System.

Die vorliegende Erfindung betrifft ein Gasturbinensystem mit einer Turbine, welche eine Komponente aufweist, die interne Kühlkanäle (235) enthält, die sich in der Nähe wenigstens einer Oberfläche (230) befinden. In einer Ausführungsform enthält der gekühlte Gegenstand ein Basismaterial (200), eine erste Schicht (205) und eine zweite Schicht (220). Hier ist die erste Schicht (205) mit dem Basismaterial verbunden, und die zweite Schicht (210) ist mit der ersten Schicht (205) verbunden, wobei wenigstens ein geschlossener Kühlkanal (235) in einem Abschnitt der ersten Schicht (205) und einem Abschnitt der zweiten Schicht (220) angeordnet ist.

Cooling of a turbine blade.

Es ist eine Kühlanordnung in einer Plattform (110) in einer Laufschaufel (100) oder einer Seitenwand in einer Leitschaufel in einer Turbine einer Verbrennungsturbine beschrieben. Die Kühlanordnung enthält: eine Kühlkammer (130), die konfiguriert ist, um ein Kühlmittel von einem Einlass (132) zu einem Auslass (134) weiterzuleiten; und eine Rippe (135), die innerhalb der Kühlkammer (130) angeordnet ist. Die Rippe (135) unterteilt die Kühlkammer teilweise, um eine Serpentine zu bilden. Die Rippe (135) verläuft in Bezug auf die Kühlkammer (130) derart schräg, dass die Serpentine einen sich zunehmend verengenden Kanal aufweist.

Zigzag cooling of the guide vane final wall.

Ein Gasturbinenleitschaufelabschnitt (100) einer Gasturbine weist eine innere Endwand (114) mit einer Vorderkante (115) auf. Ein Serpentinenkanal (116) ist im Wesentlichen innerhalb der Vorderkante (115) eingerichtet. Der Serpentinenkanal (116) weist einen Einlass und einen Auslass auf. Luft kann an dem Einlass (118) aufgenommen und unter Kühlung der Vorderkante (117) an dem Auslass (119) ausgegeben werden.

Cooling arrangement for platform region of turbine rotor blade.

Die Erfindung betrifft eine Plattformkühlanordnung in einer Turbinenlaufschaufel, die an einer Übergangsstelle zwischen einem Schaufelblatt (102) und einer Wurzel eine Plattform (110) hat. Die Plattform (110) weist eine druckseitige Schlitzseitenwand und eine saugseitige Schlitzseitenwand auf. Die Plattformkühlanordnung umfasst Folgendes: einen im Inneren der Plattform (110) ausgebildeten Kühlkanal, wobei der Kühlkanal von einem ersten Ende in Richtung auf die druckseitige Schlitzseitenwand oder die saugseitige Schlitzseitenwand verläuft. An einem zweiten Ende weist der Kühlkanal eine Tasche (104) auf. Die Tasche weist, knapp bevor der Kühlkanal die Schlitzseitenwand erreicht, eine abrupte Vergrösserung des Strömungsquerschnitts auf.

Enhanced serpentine cooling the guide vane final wall.

Ein Gasturbinenleitschaufelabschnitt (100) einer Gasturbine weist eine innere Endwand (114) mit einer Vorderkante (115) auf. Ein Serpentinenkanal (116) ist im Wesentlichen innerhalb der Vorderkante (115) eingerichtet. Der Serpentinenkanal (116) weist einen Einlass und einen Auslass auf. Luft kann an dem Einlass (118) aufgenommen und unter Kühlung der Vorderkante (117) an dem Auslass (119) ausgegeben werden.

Cooling arrangement for platform region of turbine blade.

Die Erfindung betrifft eine Plattformkühlanordnung in einer Turbinenlaufschaufel, die an einer Übergangsstelle zwischen einem Schaufelblatt (102) und einer Wurzel eine Plattform (110) hat. Die Plattform (110) weist eine druckseitige Schlitzseitenwand und eine saugseitige Schlitzseitenwand auf. Die Plattformkühlanordnung umfasst Folgendes: einen im Inneren der Plattform (110) ausgebildeten Kühlkanal, wobei der Kühlkanal von einem ersten Ende in Richtung auf die druckseitige Schlitzseitenwand oder die saugseitige Schlitzseitenwand verläuft. An einem zweiten Ende weist der Kühlkanal eine Tasche (104) auf. Die Tasche weist, knapp bevor der Kühlkanal die Schlitzseitenwand erreicht, eine abrupte Vergrösserung des Strömungsquerschnitts auf.

Serpentine cooling the vane end wall.

Ein Gasturbinenleitschaufelabschnitt (110) einer Gasturbine weist eine innere Endwand (114) mit einer Vorderkante (115) auf. Ein Serpentinenkanal (116) ist im Wesentlichen innerhalb der Vorderkante (115) eingerichtet. Der Serpentinenkanal (116) weist einen Einlass und einen Auslass auf. Luft kann an dem Einlass (118) aufgenommen und unter Kühlung der Vorderkante (117) an dem Auslass (119) ausgegeben werden.

Gas Turbines-Platform Cooling Of.

Ein Deckbandsegment (130) für ein Gehäuse einer Gasturbine enthält einen Körper, der zur Befestigung an dem Gehäuse in der Nähe einer lokalisierten kritischen Prozessstelle (144) in dem Gehäuse eingerichtet ist. Der Körper hat eine Vorderkante (132), eine Hinterkante (134) und zwei Seitenkanten (136, 138). Die kritische Prozessstelle (144) befindet sich zwischen der Vorderkante (132) und der Hinterkante (134), wenn der Körper an dem Gehäuse befestigt ist. Ein Kühlkanal (146, 148, 158, 160) ist in dem Körper entlang einer von den Seitenkanten (136, 138) mit einem Einlass (166, 154) oder einem Auslass (152, 164) in der Nähe der kritischen Prozessstelle (144) definiert. Der Kühlkanal (146, 148, 158, 160) ist gross genug dimensioniert, um die eine an den Kühlkanal (146, 148, 158, 160) angrenzende Seitenkante (136, 138) auf ein gewünschtes Niveau während des Betriebs der Gasturbine zu kühlen. Die kritischen Prozessstellen (144) können auf Temperaturen, Drücke oder andere messbare Merkmale bezüglich ...

System for cooling the wall of a gas turbine combustion chamber.

In einem Ausführungsbeispiel umfasst ein System eine Triebwerkswand (58). Die Triebwerkswand (58) weist eine Kaltseite (72) und eine Heissseite (74) auf. Die Triebwerkswand (58) ist mit einem oder mehreren Verdünnungslöchern (70) ausgebildet, wobei zu jedem der Verdünnungslöcher (70) gehört: eine erste Mündung auf der Kaltseite (72), eine zweite Mündung auf einer Heissseite (74), und ein Abschnitt einer Wärmebarrierenbeschichtung (TBC) (78), die auf der Kaltseite (72) angebracht ist und eine Öffnung aufweist, die die erste Mündung im Wesentlichen umgibt.

Air-cooled shovel for a turbine.

Eine Schaufel (20) für eine Turbine ist vorgelegt. Die Schaufel (20) weist einen Flügel (40) auf, welcher einen Fussabschnitt, einen Spitzenabschnitt, eine Flügelform und ein Nennprofil, im Wesentlichen entsprechend X-, Y- und Z-Werten in kartesischen Koordinaten, dargestellt in Tabelle I, aufweist, wobei Z eine Distanz ist, von einer Plattform (42) aus, von welcher der Flügel sich nach aussen erstreckt, und X und Y Koordinaten sind, welche das Profil bei jeder Distanz Z, von der Plattform aus, definieren, und mehrere Kühlkanäle sich zwischen dem Fussabschnitt und dem Spitzenabschnitt des Flügels erstrecken, wobei jeder der Kühlkanäle an dem Spitzenabschnitt nach aussen führt, wobei die mehreren Kühlkanäle in einem wölbungslinienförmigen Muster positioniert sind.

Rotor blade for turbine in gas turbine system, has each cooling channel exiting at point section, cooling channels arranged in pattern of curved line, and dove tail section equipped for fastening blade to turbine

The blade (22) has a dove tail section (46) equipped for fastening the blade to a turbine. A turbine blade (40) includes a foot section (120), a point section (122), a wing form and a nominal profile essentially in accordance with cartesian coordinate values. A set of cooling channels (100) includes cooling channels (102-110) extending between the foot section and the point section of the turbine blade, where each of the cooling channel (102-110) exits at the point section and the set of cooling channels (100) are arranged in a pattern of a curved line. An independent claim is also included for a gas turbine system exhibits a rotor with a rotor wheel.

Turbine blade.

Eine Turbinenschaufel (20) für eine Turbine ist vorgelegt. Die Turbinenschaufel (20) weist einen Flügel (40) auf, welcher einen Fussabschnitt, einen Spitzenabschnitt, eine Flügelform und ein Nennprofil, im Wesentlichen entsprechend X-, Y- und Z-Werten in kartesischen Koordinaten, aufweist. Hierbei ist Z eine Distanz, normal zu einer Plattform (42), von welcher der Flügel (40) sich nach aussen erstreckt, und sind X und Y Koordinaten, welche das Profil bei jeder Distanz Z, von der Plattform aus, definieren. Die Y-Achse erstreckt sich parallel zu einer Achse, um welche die Turbinenschaufel (20) rotiert. Ferner erstrecken sich mehrere Kühlkanäle zwischen dem Fussabschnitt und dem Spitzenabschnitt des Flügels (40), wobei jeder der Kühlkanäle an dem Spitzenabschnitt nach aussen führt, wobei die mehreren Kühlkanäle in einem wölbungslinienförmigen Muster positioniert sind.

Procedure for the supply of Turbulation on the inner surface of holes in an article and used articles.

Stator rotor arrangement, fluid-flow machine and procedure for making a structure of reverse Turbulatoren

Eine Stator-Rotor-Anordnung (21) wird beschreiben, die mindestens einen Verbindungsbereich (92) zwischen dem Stator (18) und dem Rotor (22) umfasst. Mindestens eine Stator- oder Rotoroberfläche im Verbindungsbereich umfasst eine Struktur aus umgekehrten Turbulatoren. Die umgekehrten Turbulatoren drosseln den Gasstrom durch einen Spalt (76) zwischen dem Stator (18) und dem Rotor (22). Verschiedene Strömungsmaschinen, die solch eine Stator-Rotor-Anordnung (21) umfassen, werden ebenfalls beschrieben. Die Offenbarung erläutert ausserdem ein Verfahren zum Herstellen einer Struktur aus umgekehrten Turbulatoren zum Drosseln des Gasstromes durch einen Spalt in einer Stator-Rotor-Anordnung mit Hilfe der umgekehrten Turbulatoren.

Component with more laser-solitaryly material layer to form of cooling passages and method for manufacture.

Es ist ein Herstellungsverfahren geschaffen. Das Herstellungsverfahren beinhaltet die Verwendung eines Laserbeschichtungsverfahrens, um ein laserabgeschiedenes Material (36) auf einer Aussenfläche (34) eines Substrats (32) aufzubringen, um auf der Aussenfläche (34) eines Substrats (32) eine oder mehrere Nuten (30) auszubilden. Jede Nut (30) weist einen Boden und eine Öffnung auf und verläuft zumindest zum Teil längs der Aussenfläche (34) des Substrats (32), wobei das Substrat (32) eine Innenfläche aufweist, die wenigstens einen hohlen Innenraum bildet. Das Herstellungsverfahren beinhaltet zudem den Schritt des Abscheidens eines zusätzlichen Materials (46) über dem laserabgeschiedenen Material (36), um einen oder mehrere Kanäle zur Kühlung des Bauteils zu bilden. Das zusätzliche Material (46) kann zusätzliche laserabgeschiedene Materialschichten oder eine Beschichtung beinhalten. Darüber hinaus betrifft die Erfindung ein mit dem Herstellungsverfahren geschaffenes Bauteil.

turbomachine shovel arrangement.

Die vorliegende Erfindung betrifft ein System, zu dem gehören: eine Turbomaschinenschaufelanordnung mit einem Schaufelblattabschnitt (50) einem Schaftabschnitt (50) und einem Befestigungsabschnitt, wobei der Schaufelblattabschnitt (50), der Schaftabschnitt und der Befestigungsabschnitt eine erste Anzahl von Lagen (200) aufweisen, die sich ausgehend von einer Spitze (202) des Schaufelblatts zu einer Basis (204) des Schwalbenschwanzes (56) erstrecken.

Fabrication method for a component with cooling slots and cooling channels.

Ein Herstellungsverfahren für eine Komponente (30) beinhaltet den Schritt der Bereitstellung eines Substrats (32) und die Ausbildung einer oder mehrerer Kanäle (40) in einer Aussenoberfläche (34) des Substrats (32) oder in einer auf der Aussenoberfläche (34) des Substrats (32) angeordneten Beschichtungslage (42) und die Ausbildung einer oder mehrerer Nuten (50) in einer Innenoberfläche (36) des Substrats (32) oder in einer auf der Innenoberfläche (36) des Substrats (32) angeordneten Beschichtungslage (46), um eine oder mehrere Kühlnuten (50) auf der Innenoberfläche (36) des Substrats (32) zu definieren. Das Verfahren beinhaltet ferner den Schritt der Aufbringung einer strukturellen Beschichtung (44) über wenigstens einem Abschnitt der Aussenoberfläche (34) des Substrats (32) oder einem Abschnitt der auf der Aussenoberfläche des Substrats angeordneten Beschichtung (42), um einen oder mehrere Kanäle (54) auf der Aussenoberflache der Komponente (30) zu definieren. Eine gemäss dem Verfahren ...

Method to form channels on a metallic substrate.

Es ist ein Verfahren zur Bildung von Kanälen (46) auf einem metallischen Substrat (30) beschrieben. Das Verfahren enthält die Schritte des Aufbringens wenigstens einer Schicht (34, 36, 44) aus einem metallischen Beschichtungsmaterial auf eine Oberfläche des Substrates (30) mittels einer Kaltspritztechnik, um Grenzwände (48) für die Kanäle (46) zu definieren und die Grenzwände (48) bis zu einer gewünschten Höhe aufzubauen. Dann wird zusätzliches Beschichtungsmaterial (50) auf eine oder mehrere Oberflächen der Grenzwände (48) mittels der Kaltspritztechnik aufgebracht, um so die Gestalt der Kanäle (46) zu modifizieren. Das Substrat (30) kann eine Hochtemperaturkomponente oder Heissgaspfadkomponente einer beliebigen Bauart sein. In einigen Fällen ist das Substrat eine Gasturbinenwand.

Method to form channels on a metallic substrate.

Es ist ein Verfahren zur Bildung von Kanälen (46) auf einem metallischen Substrat (30) beschrieben. Das Verfahren enthält die Schritte des Aufbringens wenigstens einer Schicht (34, 36, 44) aus einem metallischen Beschichtungsmaterial auf eine Oberfläche des Substrates (30) mittels einer Kaltspritztechnik, um Grenzwände (48) für die Kanäle (46) zu definieren und die Grenzwände (48) bis zu einer gewünschten Höhe aufzubauen. Dann wird zusätzliches Beschichtungsmaterial (50) auf eine oder mehrere Oberflächen der Grenzwände (48) mittels der Kaltspritztechnik aufgebracht, um so die Gestalt der Kanäle (46) zu modifizieren. Das Substrat (30) kann eine Hochtemperaturkomponente oder Heissgaspfadkomponente einer beliebigen Bauart sein. In einigen Fällen ist das Substrat eine Gasturbinenwand.

Turbine blade with a cooling channel with a turn.

Die Frequenzabstimmung, die Fluiddynamikeffizienz und das Leistungsverhalten einer Turbine können unter Verwendung eines speziellen Profils für eine Wendung (366, 367) eines Kühlkanals (364) in einem Schaufelblatt verbessert werden. Durch Vermengung von Aspekten von Basislinien- und Knollenkonturen zu einer vermengten Wendung (366, 367) mit einem ungleichförmigen Profil kann eine mechanische und/oder thermische Belastung in der Wendung (366, 367) und in einem Schaufelblatt, das die Wendung (366, 367) enthält, insbesondere auf einer Abströmungsseite der Wendung, reduziert werden. Belastungen an dem Schaufelblatt können unter Verwendung eines Wendungsprofils reduziert werden, das eine Vermengung aus einem Basislinienprofil und einem Knollenprofil ist und das durch das Schaufelblattkernprofil (301) beschrieben werden kann.

Method for modification of the shape of a channel in a metallic substrate.

Es ist ein Verfahren zur Modifikation der Gestalt eines Kanals (20) in einem metallischen Substrat beschrieben. Das Verfahren enthält den Schritt des Aufbringens wenigstens einer metallischen Beschichtung auf ausgewählte Abschnitte einer inneren Oberfläche (22, 24) des Kanals (20), um so die Wärmeübertragungseigenschaften des Kanals (20) während eines Durchgangs eines Kühlmittelfluids durch diesen zu verändern. Entsprechende Gegenstände, die die modifizierten Kanäle enthalten, sind ebenfalls beschrieben, wie beispielsweise Gasturbinenkomponenten.

Method for modification of the shape of a channel in a metallic substrate.

Es ist ein Verfahren zur Modifikation der Gestalt eines Kanals (20) in einem metallischen Substrat beschrieben. Das Verfahren enthält den Schritt des Aufbringens wenigstens einer metallischen Beschichtung auf ausgewählte Abschnitte einer inneren Oberfläche (22, 24) des Kanals (20), um so die Wärmeübertragungseigenschaften des Kanals (20) während eines Durchgangs eines Kühlmittelfluids durch diesen zu verändern. Entsprechende Gegenstände, die die modifizierten Kanäle enthalten, sind ebenfalls beschrieben, wie beispielsweise Gasturbinenkomponenten.

Components with multilayer cooling structures and method of making the same.

Das erfindungsgemässe Herstellungsverfahren beinhaltet das Bereitstellen eines Substrats (32) mit einer oder mehreren in ihm gebildeten Nuten. Auf dem Substrat (32) sind eine oder mehrere Beschichtungen (42) mit einer oder mehreren in ihnen gebildeten Nuten angeordnet und in Strömungsverbindung mit der einen oder den mehreren Nuten im Substrat. Eine Deckbeschichtung (44) ist auf einem Teil einer äussersten Oberfläche (46) der einen oder mehreren Beschichtungen (42) angeordnet, die einen oder mehrere in ihr gebildete Kühlungsauslässe (64) hat und in Strömungsverbindung mit der einen oder den mehreren Nuten in der einen oder den mehreren Beschichtungen (42) ist. Das Substrat (32), die eine oder mehreren Beschichtungen (42) und die Deckbeschichtung (44) definieren in ihnen ein Kühlungsnetz (41) zum Kühlen eines Bauteils (30). Weiter betrifft die Erfindung ein Bauteil (30) mit einem Kühlungsnetz (41), in dem ein Substrat (32), eine oder mehrere auf wenigstens einem Teil des Substrats (32) angeordnete ...

Airfoil with more vortex generators in inner cooling cavities.

Ein Schaufelblatt (10) enthält eine Vorderkante (12), eine Hinterkante (14), eine Saugseite (16) und eine Druckseite (18), wobei sich mehrere innere Hohlräume (20, 22, 24) innerhalb des Schaufelblattes (10) in Radialrichtung erstrecken, wobei sich ein erster innerer Hohlraum (24) entlang der Hinterkante (14) erstreckt. Die Hinterkante (14) ist mit mehreren entlang dieser angeordneten Kühlmittelaustrittsöffnungen versehen. Mehrere Wirbelgeneratoren sind auf einer Innenfläche der Druckseite (18) und/oder der Saugseite (16) des Schaufelblattes (10) ausgebildet. Die Wirbelgeneratoren sind erfindungsgemäss in einer radial voneinander beabstandeten Anordnung im hinteren inneren Hohlraum (24) angeordnet und erstrecken sich im Wesentlichen parallel zu der Hinterkante (14).

Airfoil with more vortex generators in inner cooling cavities.

Ein Schaufelblatt (10) enthält eine Vorderkante (12), eine Hinterkante (14), eine Saugseite (16) und eine Druckseite (18), wobei sich mehrere innere Hohlräume (20, 22, 24) innerhalb des Schaufelblattes (10) in Radialrichtung erstrecken, wobei sich ein erster innerer Hohlraum (24) entlang der Hinterkante (14) erstreckt. Die Hinterkante (14) ist mit mehreren entlang dieser angeordneten Kühlmittelaustrittsöffnungen versehen. Mehrere Wirbelgeneratoren sind auf einer Innenfläche der Druckseite (18) und/oder der Saugseite (16) des Schaufelblattes (10) ausgebildet. Die Wirbelgeneratoren sind erfindungsgemäss in einer radial voneinander beabstandeten Anordnung im hinteren inneren Hohlraum (24) angeordnet und erstrecken sich im Wesentlichen parallel zu der Hinterkante (14).

Platform cooling equipment in a turbine rotor blade as well as procedure for producing such.

Eine Plattformkühleinrichtung (110) in einer Turbinenlaufschaufel, die enthält: einen Plattformschlitz, der durch wenigstens entweder die druckseitige Schlitzseitenwand und/oder die saugseitige Schlitzseitenwand hindurchführend ausgebildet ist; einen lösbar eingesetzten Pralleinsatz (130), der die Plattform (110) in zwei radial gestapelte Plenumkammern unterteilt, wobei sich eine erste Plenumkammer (139) weiter innen von einer zweiten Plenumkammer (140) befindet; eine Hochdruck-Verbindungseinrichtung (148), die die erste Plenumkammer (139) mit einem Hochdruck-Kühlmittelbereich eines inneren Kühlkanals verbindet; eine Niederdruck-Verbindungseinrichtung (149), die die zweite Plenumkammer (140) mit einem Niederdruck-Kühlmittelbereich des inneren Kühlkanals verbindet.

Component with at least a curved film cooling hole and procedure for their production.

Es wird eine Komponente (10) bereitgestellt, die wenigstens eine Wand (12) aufweist mit einer ersten Oberfläche und einer zweiten Oberfläche (14, 16). Wenigstens ein Filmkühlungsloch (18) erstreckt sich durch die Wand zwischen den ersten und zweiten Oberflächen hindurch und hat einen Austrittsbereich (20) an der zweiten Oberfläche der Komponentenwand. Die zweite Oberfläche der Komponente hat in der Nähe des Austrittsbereichs eine nicht-ebene Krümmung. Das Filmkühlungsloch ist an dem Austrittsbereich dergestalt abgeschrägt, dass sich die Krümmung des Filmkühlungsloches in dem Austrittsbereich an die nicht-ebene Krümmung der zweiten Oberfläche der Komponentenwand anpasst und dadurch einen gekrümmten Austrittsbereich erzeugt. Ein Verfahren wird ebenfalls zum Erzeugen wenigstens eines Filmkühlungsloches (18) in der Komponente (10) bereitgestellt. Das Verfahren beinhaltet die Erzeugung eines geraden Abschnittes (26) in der Komponentenwand (12) dergestalt, dass sich der gerade Abschnitt durch ...

System for the cooling of turbine blades.

In einer Ausführungsform enthält ein System (10) eine Turbinenlaufschaufel (40) mit einer radialen Schaufelspitze (62). Das System (10) enthält weiterhin eine Hinterkantennut (94), die in der radialen Schaufelspitze (62) ausgebildet ist und sich zu einer Hinterkante (68) der Turbinenlaufschaufel (40) hin erstreckt. Die Hinterkantennut (94) enthält weiterhin eine erste Gruppe von Kühlkanälen, von denen jede eine erste Nut aufweist, die entlang einer ersten Seitenwand der Hinterkantennut ausgebildet ist, wobei die Nut mit einer zugehörigen ersten Öffnung verbunden ist, die sich durch einen Boden der Hinterkantennut hindurch erstreckt.

Turbine blade with a platform cooling equipment and process for their preparation.

Die Erfindung betrifft eine Turbinenlaufschaufel mit einer Plattformkühleinrichtung für eine Plattform (110) der Turbinenlaufschaufel, umfassend: einen Plattformschlitz, der in wenigstens entweder einer druckseitigen Schlitzseitenwand und/oder einer saugseitigen Schlitzseitenwand der Turbinenlaufschaufel ausgebildet ist; einen lösbar eingesetzten Pralleinsatz (130), der den Plattformschlitz in zwei radial gestapelte Plenumkammern unterteilt, wobei sich eine erste Plenumkammer (139) radial weiter innen von einer zweiten Plenumkammer (140) befindet; eine Hochdruck-Verbindungseinrichtung (148), die die erste Plenumkammer (139) mit einem Hochdruck-Kühlmittelbereich eines im Inneren der Turbinenlaufschaufel ausgebildeten Kühlkanals verbindet; eine Niederdruck-Verbindungseinrichtung (149), die die zweite Plenumkammer (140) mit einem Niederdruck-Kühlmittelbereich des Kühlkanals verbindet.

Platform cooling arrangement in a turbine rotor blade as well as procedure for the production of such.

Eine Plattformkühlanordnung einer Turbinenrotorschaufel weist eine Plattform auf, die einen in ihr ausgebildeten Kühlkanal (116) hat. Die Plattformkühlanordnung weist auf: Ein Hauptvolumen (132), das etwas innerbords bezüglich der planaren Oberseite angeordnet ist und sich innerhalb der Druckseite und/oder der Saugseite der Plattform von einer hinteren Position in eine vordere Position erstreckt, wobei das Hauptvolumen (132) eine Längsachse aufweist, die ungefähr parallel zu der planaren Oberseite ausgerichtet ist; ein Liefervolumen (134), das sich zwischen dem Hauptvolumen (132) und dem inneren Kühlkanal (116) erstreckt; sowie eine Anzahl von Kühlöffnungen (136), wobei sich jede Kühlöffnung (136) von der druckseitigen und/oder der saugseitigen Spaltfläche (126) weg zu einer Verbindung mit dem Hauptvolumen (132) erstreckt.

Shovel sheet with several eddy generators in internal cooling cavities.

Ein Schaufelblatt (10) enthält eine Vorderkante (12), eine Hinterkante (14), eine Saugseite und eine Druckseite, wobei sich mehrere innere Hohlräume (20, 22, 24) innerhalb des Schaufelblattes in Radialrichtung erstrecken, wobei sich einer der mehreren inneren Hohlräume (24) entlang der Hinterkante (14) erstreckt. Die Hinterkante ist mit mehreren entlang dieser angeordneten Kühlmittelaustrittsöffnungen versehen. Mehrere Wirbelgeneratoren sind auf einer Innenfläche von wenigstens entweder der Druckseite und/oder der Saugseite des Schaufelblattes ausgebildet. Die Wirbelgeneratoren sind in einer radial voneinander beabstandeten Anordnung in einem der mehreren inneren Hohlräume (24) angeordnet und erstrecken sich im Wesentlichen parallel zu den und in der Nähe der mehreren Kühlmittelaustrittsöffnungen.

Turbine blade, Turbinenrotor, and procedure for the cooling of a turbine blade.

Die Erfindung betrifft eine Turbinenschaufel mit einem Blattkühlungsdurchgang, einem Deckband (120), einer Wandung mit Austrittsöffnung für Kühlfluid aus dem Blattkühlungsdurchgang, einer Deckband-Kühlkammer (142), verbunden mit der Austrittsöffnung, wobei die Austrittsöffnung auf eine Ziel-Wandfläche (134) der Deckband-Kühlkammer (142) gerichtet ist, wodurch die Austrittsöffnung eine Prallöffnung (132) zur Prallkühlung der Ziel-Wandfläche (134) als eine Prallzone definiert und zumindest eine Auslassöffnung (136, 138) zum Ausleiten von verbrauchtem Prallkühlungsfluid aus der Deckband-Kühlkammer.

Turbine blade with shovel point cooling.

Die Erfindung betrifft eine Turbinenlaufschaufel (40) mit einer Schaufelspitze (62). Die Schaufelspitze (62) enthält weiterhin eine Schaufelspitzennut (94), die in der radialen Schaufelspitze (62) ausgebildet ist und sich zu einer Hinterkante (68) der Turbinenlaufschaufel (40) hin erstreckt. Die Schaufelspitzennut (94) enthält weiterhin erste Kühlkanäle, welche sich durch den Nutboden und zur Hinterkante (68) geneigt erstrecken. In Verlängerung der Kühlkanäle können Kühlkanalnuten in den Nutseitenwänden vorgesehen sein.

COMPONENT FOR A GAS TURBINE ENGINE WITH A FILM HOLE

A component is provided and comprises at least one wall comprising a first and a second surface. At least one film cooling hole extends through the wall between the first and second surfaces and has an outlet region at the second surface. The film cooling hole includes a first expansion section being a side diffusion portion and a second expansion section being a layback diffusion portion, wherein the side diffusion portion is upstream and spaced from the layback diffusion portion. 1. A component for a gas turbine engine comprising:a hot side exposed to a hot air flow;a cool side exposed to a cooling air flow;a film hole passage extending between the cool side and the hot side with an inlet on the cool side and an outlet on the hot side, the film hole passage defining a diameter, the film hole passage further defining a side diffusion portion defining a side diffusion length between a start of the side diffusion portion and the outlet, and a layback diffusion portion defining a layback length between a start of the layback diffusion portion and the outlet, wherein the side diffusion length is greater than the layback diffusion length.2. The component of wherein the side diffusion portion is upstream and spaced from the layback diffusion portion.3. The component of wherein the side diffusion portion defines a side diffusion angle claim 1 , α claim 1 , relative to a centerline for the film hole passage claim 1 , and the side diffusion angle is less than 12.5 degrees.4. The component of wherein the layback diffusion portion defines a layback diffusion angle claim 3 , ß claim 3 , relative to a centerline for the film hole passage claim 3 , and the layback diffusion angle is less than 12 degrees.5. The component of wherein the layback diffusion length is less than 4 times the diameter.6. The component of wherein the layback diffusion length is equal to or less than zero.7. The component of wherein the layback diffusion length is less than four times the diameter and the ...

NON-CONTACT SEAL ASSEMBLY FOR ROTATIONAL EQUIPMENT

Assemblies are provided for rotational equipment. One of these assemblies includes a first bladed rotor assembly, a second bladed rotor assembly, a stator vane assembly, a stator structure and a seal assembly. The second bladed rotor assembly includes a rotor disk structure. The stator vane assembly is axially between the first and the second bladed rotor assemblies. The stator structure is mated with and radially within the stator vane assembly. The seal assembly is configured for sealing a gap between the stator structure and the rotor disk structure, wherein the seal assembly includes a non-contact seal. 1. An assembly for rotational equipment , the assembly comprising:a first bladed rotor assembly;a second bladed rotor assembly including a rotor disk structure;a stator vane assembly axially between the first and the second bladed rotor assemblies;a stator structure mated with and radially within the stator vane assembly; anda seal assembly configured for sealing a gap between the stator structure and a seal land of the rotor disk structure, wherein the seal assembly includes a non-contact seal, and the seal land is configured as a cantilevered tubular body.2. The assembly of claim 1 , wherein the non-contact seal is a hydrostatic non-contact seal.3. The assembly of claim 1 , wherein the non-contact seal comprises:an annular base;a plurality of shoes arranged around and radially adjacent the rotor disk structure; anda plurality of spring elements, each of the spring elements radially between and connecting a respective one of the shoes to the base.4. The assembly of claim 3 , wherein the base is configured with a monolithic full hoop body.5. The assembly of claim 1 , wherein the stator structure is floating radially within the stator vane assembly.6. The assembly of claim 1 , whereinthe first and the second bladed rotor assemblies are turbine rotor assemblies; andthe first bladed rotor assembly is upstream of the second bladed rotor assembly.7. The assembly of ...

THERMALLY DRIVEN SPRING VALVE FOR TURBINE GAS PATH PARTS

A thermally driven spring valve for turbine gas path parts is disclosed herein. A thermally driven spring valve includes a bimetallic sheet comprising a base, a first finger portion extending from the base and a second finger portion extending from the base, the first finger portion having a first curvature vector and the second finger portion have a second curvature vector, wherein an exterior surface extends from the base through the first finger portion and the second finger portion and an interior surface extends from the base through the first finger portion and the second finger portion, wherein the exterior surface of the first finger portion is disposed proximate the interior surface of the base wherein the exterior surface of the second finger portion is disposed proximate the interior surface of the base. A thermally driven spring valve may include perforations through a finger portion. 1. A thermally driven spring valve comprising:a metallic sheet comprising a base mount portion and a floating portion having a curvature vector, wherein the base mount portion is coupled to a wall of a chamber, wherein the floating portion is disposed proximate an aperture in the wall.2. The thermally driven spring valve of , wherein the metallic sheet is coupled to the wall of the chamber by at least one of brazing or welding. The thermally driven spring valve of , wherein the metallic sheet is a bimetallic sheet.4. The thermally driven spring valve of claim 1 , wherein the metallic sheet comprises at least one of steel claim 1 , titanium claim 1 , titanium alloy claim 1 , cobalt claim 1 , cobalt alloy claim 1 , platinum claim 1 , or platinum alloy.5. The thermally driven spring valve of claim 1 , wherein the metallic sheet has a coefficient of thermal expansion of between about 0.6×10/K to about 25×10/K.6. The thermally driven spring valve of claim 1 , wherein the chamber is coupled to a baffle. This application is a divisional of, and claims priority to, and the benefit ...

MANUFACTURING METHODS FOR MULTI-LOBED COOLING HOLES

A method for producing a diffusion cooling hole extending between a wall having a first wall surface and a second wall surface includes forming a cooling hole inlet at the first wall surface, forming a cooling hole outlet at the second wall surface, forming a metering section downstream from the inlet and forming a multi-lobed diffusing section between the metering section and the outlet. The inlet, outlet, metering section and multi-lobed diffusing section are formed by laser drilling, particle beam machining, fluid jet guided laser machining, mechanical machining, masking and combinations thereof. 1. A method for producing a diffusion cooling hole extending between a wall having a first wall surface and a second wall surface , the method comprising:forming a cooling hole inlet at the first wall surface;forming a cooling hole outlet at the second wall surface;forming a metering section downstream from the inlet; andforming a multi-lobed diffusing section between the metering section and the outlet,wherein the inlet, outlet, metering section and multi-lobed diffusing section are formed by a technique selected from the group consisting of laser drilling, particle beam machining, fluid jet guided laser machining, mechanical machining, masking and combinations thereof.2. The method of claim 1 , wherein the wall comprises a metal or superalloy substrate.3. The method of claim 1 , wherein the second wall surface comprises a coating claim 1 , and wherein at least a portion of the cooling hole extends through the coating.4. The method of claim 3 , wherein the coating comprises:a bond coating; anda thermal barrier coating.5. The method of claim 4 , wherein a portion of the diffusing section is located within the coating.6. The method of claim 5 , wherein the entire diffusing section is located within the coating.7. The method of claim 6 , wherein a portion of the metering section is located within the coating.8. The method of claim 1 , wherein the inlet and metering section ...

TURBINE ENGINE THERMAL MANAGEMENT

A gas turbine engine including core engine is provided. Air may enter the core engine through an inlet and travel through and engine air flowpath extending through the core engine, e.g., generally along an axial direction of the gas turbine engine. The gas turbine engine additionally includes a cooling air flowpath extending outwardly generally along the radial direction of the gas turbine engine. The cooling air flowpath extends between an inlet in flow communication with engine air flowpath and an outlet defined by an opening in an outer casing of the core engine. Moreover, the gas turbine engine includes a heat exchanger positioned at least partially within the outer casing the core engine with the cooling air flowpath extending over or through the heat exchanger. 1. A gas turbine engine defining a radial direction , the gas turbine engine comprising:a core engine including an outer casing;an engine air flowpath extending through the core engine;a cooling air flowpath extending between an inlet in flow communication with the engine air flowpath and an outlet defined by an opening in the outer casing of the core engine; anda heat exchanger positioned in thermal communication with the cooling air flowpath.2. The gas turbine of claim 1 , wherein the core engine includes a vent over the opening in the outer casing claim 1 , the vent configured to adjust an amount of airflow allowable through the cooling air flowpath.3. The gas turbine of claim 1 , wherein the heat exchanger is rigidly attached to the outer casing and is configured as an air cooled oil cooler.4. The gas turbine of claim 3 , wherein the core engine includes an annular compressor frame positioned within the outer casing claim 3 , and wherein the heat exchanger is also rigidly attached to the annular compressor frame such that the heat exchanger provides structural support between the outer casing and the annular compressor frame.5. The gas turbine of claim 1 , wherein the core engine includes a ...

SYSTEM AND METHOD OF FABRICATING AND REPAIRING A GAS TURBINE COMPONENT

A method of fabricating and repairing a gas turbine component having a plurality of cooling holes defined therein is provided. The method includes determining a parameter of a first cooling hole defined in the gas turbine component, and generating a tool path for forming a protective cap around the first cooling hole. The tool path is based at least partially on the parameter of the first cooling hole. The method also includes directing a robotic device to follow the tool path, and discharging successive layers of ceramic slurry towards the gas turbine component as the tool path is followed such that the protective cap is formed around the first cooling hole. 1. A method of fabricating and repairing a gas turbine component having a plurality of cooling holes defined therein , said method comprising:determining a parameter of a first cooling hole defined in the gas turbine component;generating a tool path for forming a protective cap around the first cooling hole, the tool path based at least partially on the parameter of the first cooling hole;directing a robotic device to follow the tool path; anddischarging successive layers of ceramic slurry towards the gas turbine component as the tool path is followed such that the protective cap is formed around the first cooling hole.2. The method in accordance with claim 1 , wherein determining a parameter comprises determining at least one of a size of the first cooling hole claim 1 , an edge profile of the first cooling hole claim 1 , or a location of the first cooling hole on the gas turbine component.3. The method in accordance with claim 1 , wherein determining a parameter comprises conducting a non-destructive inspection of the gas turbine component.4. The method in accordance with claim 1 , wherein generating a tool path comprises determining an arrangement of a plurality of individual layers for forming a three-dimensional representation of the protective cap claim 1 , wherein the plurality of individual layers ...

ENHANCED HEAT SINK AVAILABILITY ON GAS TURBINE ENGINES THROUGH THE USE OF COOLERS

A cooling assembly for a gas turbine engine may include a heat source, a heat sink, and a heat pump coupled to the heat source and the heat sink, wherein the heat pump may be configured to convey a quantity of heat from the heat source to the heat sink. The heat pump may be mounted to an inner surface of an inner fan casing, and the heat sink may be mounted to an outer surface of the inner fan casing such that heat may convey from the heat sink to a bypass air stream passing over the inner fan casing through or past the heat sink. 120-. (canceled)21. A cooling assembly for a gas turbine engine comprising:a heat source;a heat sink; anda heat pump coupled to the heat source and the heat sink, wherein the heat pump is configured to convey a quantity of heat from the heat source to the heat sink;wherein the heat pump is mounted to an inner surface of an inner fan casing, the heat sink is mounted to an outer surface of the inner fan casing such that heat conveys from the heat sink to a bypass air stream passing over the inner fan casing through or past the heat sink.22. The cooling assembly of claim 21 , wherein the heat source is a heat exchanger having a material passing therethrough.23. The cooling assembly of claim 22 , wherein the material is a heat transfer fluid or gas.24. The cooling assembly of claim 23 , wherein the heat transfer fluid or gas is one of compressor air claim 23 , oil claim 23 , fuel claim 23 , or coolant.25. The cooling assembly of claim 21 , wherein the heat pump is a generator.26. The cooling assembly of claim 21 , wherein the heat sink is a surface cooler heat exchanger.27. A heat transfer system for a gas turbine engine comprising:a surface cooler heat exchanger;a generator; anda heat exchanger heat source coupled to the generator, wherein the heat exchanger heat source is configured to remove heat from a material passing therethrough, and wherein the generator is configured to cause a quantity of heat to pass from the heat exchanger heat ...

COOLED FUEL INJECTOR SYSTEM FOR A GAS TURBINE ENGINE AND A METHOD FOR OPERATING THE SAME

A cooled fuel injector system of a combustor section of a gas turbine engine is provided. At least a part of the fuel injector system is exposed to core gas flow traveling through the engine. The cooled fuel injector system includes a source of a first cooling fluid and a fuel injector system component. The first cooling fluid is at a temperature lower than a temperature of the core gas flow proximate the fuel injector system. The fuel injector system component includes a vascular engineered structure lattice (VESL) structure, which VESL structure is in fluid communication with the source of the cooling fluid. 1. A cooled fuel injector system of a combustor section of a gas turbine engine , at least a part of the fuel injector system exposed to a core gas flow traveling through the gas turbine engine , the cooled fuel injector system comprising:a first source of a first cooling fluid, the first cooling fluid at a temperature lower than a temperature of the core gas flow proximate the fuel injector system; and a plurality of nodes, each node of the plurality of nodes comprising a solid node structure; and', 'a plurality of branches extending from each node of the plurality of nodes, each branch of the plurality of branches comprising a solid branch structure, wherein only a single branch of the plurality of branches extends between and connects adjacent nodes of the plurality of nodes,, 'a fuel injector system component including a vascular engineered structure lattice (VESL) structure disposed between a first wall and a second wall of the fuel injector system component, the first wall spaced from the second wall, the VESL structure in fluid communication with the source of the first cooling fluid, the VESL structure comprisingwherein a space between the first wall and the second wall of the fuel injector system component and exterior surfaces of the plurality of nodes and the plurality of branches defines a plurality of open passages and wherein the plurality of ...

COATING SYSTEM INCLUDING OXIDE NANOPARTICLES IN OXIDE MATRIX

In some examples, an article may include a substrate and a coating on the substrate. The substrate may include a superalloy, a ceramic, or a ceramic matrix composite. The coating may include a layer comprising a matrix material and a plurality of nanoparticles. The matrix material may include at least one of silica, zirconia, alumina, titania, or chromia, and the plurality of nanoparticles may include nanoparticles including at least one of yttria, zirconia, alumina, or chromia. In some examples, an average diameter of the nanoparticles is less than about 400 nm. 1. A system comprising: the oxide matrix comprises at least one of silica, zirconia, alumina, titania, or chromia;', 'the plurality of oxide nanoparticles comprises at least one of yttria, zirconia, alumina, or chromia;', 'a chemical composition of the plurality of oxide nanoparticles is different from the chemical composition of the oxide matrix such that the plurality of oxide nanoparticles form a second, distinct phase in the oxide matrix;', 'an average diameter of the plurality of oxide nanoparticles is less than 400 nm; and', 'the layer is formed from a mixture that comprises a precursor of the oxide matrix and between about 0.7 volume percent and about 13 volume percent of the plurality of oxide nanoparticles, wherein the volume percent of the plurality of oxide nanoparticles is based on the volume of oxide nanoparticles divided by a total volume of the plurality of oxide nanoparticles plus the precursor of the oxide matrix; and, 'a first component comprising an alloy substrate comprising an alloy substrate coated with a coating, wherein the coating comprises a layer comprising an oxide matrix and a plurality of oxide nanoparticles, and whereina second component comprising a ceramic or a CMC substrate, wherein at least a portion of the second component is in contact with at least a portion of the coating.2. The system of claim 1 , wherein the mixture from which the layer is formed comprises between ...

CMAS RESISTANT THERMAL BARRIER COATINGS

The present application provides Calcia-Magnesia-Alumina-Silica (CMAS) (or molten silicate) resistant thermal barrier coatings (IBC). The coatings include elongate growth domains of non-equiaxed, randomly arranged overlapping grains or splats. The elongate growth domains include overlapping individual, randomly distributed splats of tough and soft phases. In some embodiments, the elongate growth domains are formed via air plasma spray. In some embodiments, the tough phases are at least partially stabilized zirconia and/or hafnia compositions, and the soft phases are CMAS (or molten silicate) reactive or resistant compositions. Within each elongate growth domain, the mixture of the tough and soft phases act together to limit penetration of CMAS and also provide sufficient domain toughness to minimize cracking forces produced during crystallization of infiltrated CMAS. The soft phases may react with the CMAS and increase its melting point, increase its viscosity, and reduce the destabilization of the tough phases. 119.-. (canceled)20. A method of forming a thermal barrier coating on a substrate , the method comprising:obtaining a substrate;obtaining a feedstock consisting of about micron or sub-micron ceramic particles of tough and soft phases suspended in a liquid agent, wherein the tough phases are at least one of partially stabilized zirconia compositions and partially stabilized hafnia compositions, and the soft phases are at least one of CMAS reactive compositions and CMAS resistant compositions; andutilizing an air plasma spray apparatus to heat and deposit the tough and soft phases of the feedstock on the substrate in randomly distributed overlapping splats that form a plurality of elongate material growth domains of at least about 75% density defined between domain boundaries.21. The method of claim 20 , wherein at least about 75% of the splats of the domains include a width to length aspect ratio of greater than or equal to about 3:1 and a substantially ...

TURBINE BLADE

An airfoil comprises one or more internal cooling circuits. The cooling circuit can further comprise a near wall cooling mesh, fluidly coupling a supply passage to a mesh plenum. The mesh plenum can be disposed adjacent to the external surface of the airfoil having a plurality of film holes extending between the mesh plenum and the external surface of the airfoil. The mesh plenum can further comprise a cross-sectional area sized to facilitate machining of the film holes without damage to the interior of the airfoil. 1. An airfoil for a gas turbine engine , the airfoil comprising:an outer surface defining a pressure side and a suction side extending axially between a leading edge and a trailing edge and extending radially between a root and a tip; and a radially extending supply passage fluidly coupled to a cooling air inlet passage,', 'a near wall cooling mesh located adjacent to and extending along a portion of the outer surface,', 'a radially extending opening fluidly coupling the supply passage to the near wall cooling mesh to define a fluid inlet for the near wall cooling mesh, and', 'a plenum fluidly coupled to the near wall cooling mesh to define a fluid outlet for the near wall cooling mesh;, 'a cooling circuit located within the airfoil and comprisingwherein cooling air flows through the near wall cooling mesh from the inlet to the outlet and the cross-sectional area of the plenum is greater than the cross-sectional area of the near wall cooling mesh in an airflow direction.2. The airfoil according to further comprising a pin bank located within the near wall cooling mesh.3. The airfoil according to wherein one of the fluid inlet and fluid outlet for the near wall cooling mesh is located closer to the leading edge than the other of the fluid inlet and fluid outlet for the near wall cooling mesh.4. The airfoil according to wherein the near wall cooling mesh is located adjacent to one of the pressure side and suction side.5. The airfoil according to wherein ...

ANGLED IMPINGEMENT INSERTS WITH COOLING FEATURES

An engine component assembly for impingement cooling. The engine component assembly includes an engine first component having a cooled surface. The engine first component having a flow path on one side of the cooled surface. A second component is a disposed adjacent to the engine first component between the flow path and the engine first component, and has a plurality of openings forming an array through the second component. The cooling flow path passes through the plurality of openings to cool the cooled surface. The second component having a surface facing the cooled surface of the engine first component. A plurality of discrete cooling features that have at least one wall that has a curved cross-section extend from the second component surface into a gap between and toward the cooled surface of the engine first component and defining an array. 1. An engine component assembly for impingement cooling , comprising:an engine first component having a cooled surface;said engine first component having a flow path on one side of said cooled surface;a second component disposed adjacent to said engine first component between said flow path and said engine first component, said second component having a plurality of openings forming an array through said second component, said cooling flow path passing through said plurality of openings to cool said cooled surface;said second component having a surface facing said cooled surface of said engine first component;a plurality of discrete cooling features extending from said second component surface into a gap between and toward said cooled surface of said engine first component and defining an array, wherein a forward wall of at least one cooling feature of the plurality of cooling features has a curved cross-section; andsaid openings extending through said second component at a non-orthogonal angle to said second component surface; andwherein the array comprises a staggered array, the staggered array comprising a first row and ...

SEAL ASSEMBLY FOR TURBINE ENGINE COMPONENT

A seal assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a housing mountable to an engine static structure, a seal carrier secured to the housing and configured to be selectively biased from the housing, and a wedge seal secured to the seal carrier. The wedge seal abuts against sealing surfaces of adjacent blade outer air seals in response to movement of the seal carrier relative to the housing. The wedge seal is separate and distinct from the seal carrier. A method of sealing between adjacent components of a gas turbine engine is also disclosed. 1. A seal assembly for a gas turbine engine comprising:a housing mountable to an engine static structure;a seal carrier secured to the housing and configured to be selectively biased from the housing; anda wedge seal secured to the seal carrier, wherein the wedge seal abuts against sealing surfaces of adjacent blade outer air seals in response to movement of the seal carrier relative to the housing, and wherein the wedge seal is separate and distinct from the seal carrier.2. The seal assembly as recited in claim 1 , wherein the wedge seal span across a gap defined by opposed mate faces of the adjacent blade outer air seals.3. The seal assembly as recited in claim 2 , wherein the seal carrier defines a spring cavity receiving a spring that is configured to bias the seal carrier away from the housing.4. The seal assembly as recited in claim 2 , wherein the wedge seal defines a first engagement surface and a second engagement surface joined at an apex.5. The seal assembly as recited in claim 1 , further comprising a spring that is configured to bias the seal carrier away from the housing.6. The seal assembly as recited in claim 1 , wherein the wedge seal defines a first engagement surface and a second engagement surface joined at an apex that faces away from the housing.7. The seal assembly as recited in claim 1 , comprising an insulation member arranged between ...

TURBINE SHROUD SEGMENT WITH LOAD DISTRIBUTION SPRINGS

A turbine shroud adapted for use in a gas turbine engine includes a plurality of metallic carrier segments and a plurality of blade track segments mounted to corresponding metallic carrier segments. Cooling air is directed onto the blade track segments to cool the blade track segments when exposed to high temperatures in a gas turbine engine. 1. A turbine shroud segment adapted for use in a gas turbine engine having a central axis , the turbine shroud segment comprisinga carrier segment comprising metallic materials, the carrier segment formed to include an attachment-receiving space,a blade track segment comprising ceramic matrix composite materials, the blade track segment formed to include a runner shaped to extend at least partway around the central axis and an attachment portion that extends radially outward from the runner into the attachment-receiving space formed by the carrier segment,an attachment assembly including a first attachment post that extends from the carrier segment through an attachment hole formed in the attachment portion of the blade track segment, a first attachment support arranged inside a cavity formed by the attachment portion of the blade track segment that is shielded by the runner of the blade track segment from the central axis and coupled to the first attachment post to block withdrawal of the attachment post through the attachment hole, and a first spring member arranged outside of the attachment-receiving space and configured to pull the first attachment support radially outward away from the central axis, wherein the first attachment support contacts the attachment portion of the blade track segment at locations spaced apart from the attachment hole formed in the attachment portion of the blade track segment.2. The turbine shroud segment of claim 1 , wherein the first spring member is arranged radially outward of the carrier segment.3. The turbine shroud segment of claim 2 , wherein the first spring member is a coil spring that ...

GEARED TURBOFAN ARRANGEMENT WITH CORE SPLIT POWER RATIO

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section including a fan having a plurality of fan blades, and a nacelle surrounding the plurality of fan blades, a compressor section including a low pressure compressor and a high pressure compressor, the low pressure compressor including a plurality of stages, and the high pressure compressor including 6 or more stages. A turbine section includes a fan drive turbine that drives the fan section through a gear arrangement, and including a second turbine that drives the high pressure compressor. A power ratio is provided by the combination of a first power input of the low pressure compressor and a second power input of the high pressure compressor, the power ratio defined by the second power input divided by the first power input, and the power ratio is less than or equal to 1.0 1. A gas turbine engine comprising:a fan section including a fan having a plurality of fan blades, and a nacelle surrounding the plurality of fan blades;a compressor section including a low pressure compressor and a high pressure compressor, the low pressure compressor including a plurality of stages, and the high pressure compressor including 6 or more stages;wherein the fan section delivers a portion of air into the compressor section, and a portion of air into a bypass duct, and a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section, is equal to or greater than 12;a turbine section including a fan drive turbine that drives the fan section through a gear arrangement, and including a second turbine that drives the high pressure compressor, anda power ratio provided by the combination of a first power input of the low pressure compressor and a second power input of the high pressure compressor, the power ratio defined by the second power input divided by the first power input, and the power ratio ...

DIVERSION OF FAN AIR TO PROVIDE COOLING AIR FOR GAS TURBINE ENGINE

A gas turbine engine section includes a plurality of spaced rotor stages, with a static guide vane intermediate the spaced rotor stages. The static guide vane provides swirl into air passing toward a downstream one of the spaced rotor stages, and an outer housing surrounding the spaced rotor stages. A diverter diverts a portion of air radially outwardly through the outer housing, and across at least one heat exchanger. The diverted air passes back into a duct radially inwardly through the outer housing, and is exhausted toward the downstream one of the spaced rotor stages. 1. A gas turbine engine section comprising:a plurality of spaced rotor stages, with a static guide vane intermediate said spaced rotor stages, said static guide vane providing swirl into air passing toward a downstream one of said spaced rotor stages, and an outer housing surrounding said spaced rotor stages, a diverter diverting a portion of air radially outwardly through said outer housing, and across at least one heat exchanger, with the diverted air passing back into a duct radially inwardly through said outer housing, and being exhausted toward said downstream one of said spaced rotor stages.2. The gas turbine engine as set forth in claim 1 , wherein said exhausted air passing through an injector claim 1 , and said injector imparting swirl into the air exhausting toward the downstream one of the two spaced turbine rotors.3. The gas turbine engine as set forth in claim 2 , wherein a swirl angle imparted by said injector is greater than a swirl angle imparted by said static guide vane.4. The gas turbine engine as set forth in claim 3 , wherein a swirl angle imparted by said static guide vane is greater than 40 degrees.5. The gas turbine engine as set forth in claim 4 , wherein said swirl angle is greater than 55 degrees.6. The gas turbine engine as set forth in claim 1 , wherein said at least one heat exchanger cooling an electronic component.7. The gas turbine engine as set forth in claim 1 , ...

CROSS-STREAM HEAT EXCHANGER

A heat exchanger system for a gas turbine engine is disclosed. The heat exchanger system may include a first structure at least partially defining a first plenum configured to receive a first air stream, a second structure at least partially defining a second plenum configured to receive a second air stream having lower pressure than the first air stream, a third structure at least partially defining a third plenum configured to receive a third air stream having lower pressure than the second air stream, and a heat exchanger configured for operative communication with the first air stream, the second air stream, and the third air stream while disposed between the second air stream and the third air stream. The heat exchanger may be configured to transfer heat from the first air stream to the third air stream. 1. A heat exchanger system for a gas turbine engine , comprising:a first structure at least partially defining a first plenum configured to receive a first air stream;a second structure at least partially defining a second plenum configured to receive a second air stream having lower pressure than the first air stream;a third structure at least partially defining a third plenum configured to receive a third air stream having lower pressure than the second air stream; anda heat exchanger configured for operative communication with the first air stream, the second air stream, and the third air stream, the heat exchanger configured to transfer heat from a portion of the first air stream to a portion of the second air stream at the heat exchanger, the portion of the second air stream flowing to the third air stream, the heat exchanger disposed between the second structure and the third structure;wherein the heat exchanger includes a flow metering device configured to control flow through the heat exchanger, the flow metering device including a plurality of slats in a louver arrangement.2. The heat exchanger system of claim 1 , wherein the heat exchanger is disposed ...

METHOD OF ASSEMBLY OF BI-CAST TURBINE VANE

One aspect of the present disclosure includes a turbine vane assembly comprising a vane made from ceramic matrix composite material having an outer wall extending between a leading edge and a trailing edge and between a first end and an opposing second end; an endwall made at least partially from a ceramic matrix composite material configured to engage the first end of the vane; and a retaining region including corresponding bi-cast grooves formed adjacent the first end of the vane and a receiving aperture formed in the endwall; wherein a bond is formed in the retaining region to join the vane and endwall together. 120.-. (canceled)21. A method of assembling a gas turbine engine vane including an airfoil having an outer surface extending between a leading edge and a trailing edge and between a first end and a second end; a through slot extending between the first and second ends of the airfoil; and a spar slidingly engaged within the slot of the airfoil , the spar including a pair of extensions with at least one bi-cast groove formed on opposing ends thereof , wherein the extensions of the spar are configured to engage with corresponding apertures formed in a pair of opposing endwalls , the method comprising:inserting at least one of the extensions into one of the apertures of the end walls,injecting a pre-cursor into the at least one bi-cast groove,heating the pre-cursor to form a fixed connection.22. The method of claim 21 , wherein the pre-cursor is initially a powder or a liquid.23. The method of claim 21 , wherein a cross sectional shape of the spar substantially conforms with the cross sectional shape of the through slot and at least one of the pair of extensions protrudes from the through slot.24. The method of claim 21 , wherein heating the pre-cursor including forming a fixed connection between the spar relative to at least one of the endwalls while the airfoil remains slidingly free within the through slot.25. The method of claim 21 , wherein injecting the ...

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. 1. A finishing apparatus for a film cooled article comprising:a central control machine including a computerized controller;at least one articulating device controlled by said central control machine;a machining tool mounted to said articulating device, such that said articulating device is operable to move said machining tool;at least one of a touch sensor apparatus and a visual sensor apparatus mounted to said machining tool and communicatively coupled to the computerized controller; andwherein said computerized controller stores instructions operable to cause said finishing apparatus to perform the steps of:positioning a machining element in contact with an exterior surface of a film cooled article;automatically moving said machining element along said exterior surface while maintaining contact between the machining element and the surface;identifying an actual position of at least one film opening based on sensory feedback from the machining element using a controller while maintaining contact between the machining element and the surface;removing material from said exterior surface at said film opening using said machining element, thereby creating a depression at said at least one film opening ...

AIRFOIL WITH DUAL PROFILE LEADING END

An airfoil includes first and second endwalls and an airfoil section having a dual airfoil profile. The airfoil section includes a double wall that has an outer wall that defines a primary leading end of the dual airfoil profile and an inner wall spaced from the outer wall. The outer wall is trapped between the first and second endwalls in mortise and tenon joints, and the inner wall defines a secondary leading end of the dual airfoil profile. 1. An airfoil comprising:first and second endwalls; an outer wall that defines a primary leading end of the dual airfoil profile, the outer wall being trapped between the first and second endwalls in mortise and tenon joints, and', 'an inner wall spaced from the outer wall, and the inner wall defines a secondary leading end of the dual airfoil profile., 'an airfoil section having a dual airfoil profile, the airfoil section including a double wall having'}2. The airfoil as recited in claim 1 , wherein the outer wall is formed of a first material composition claim 1 , and the inner wall is formed of a second claim 1 , different material composition.3. The airfoil as recited in claim 2 , wherein the first material composition is ceramic and the second material composition is metal.4. The airfoil as recited in claim 1 , wherein the outer wall has an exterior side and an interior side claim 1 , and the inner wall has a plurality of cooling holes that open to the interior side of the outer wall.5. The airfoil as recited in claim 1 , wherein the outer wall is formed of a first material composition claim 1 , the inner wall is formed of a second claim 1 , different material composition claim 1 , the first material composition is ceramic claim 1 , the second material composition is metal claim 1 , the outer wall has an exterior side and an interior side claim 1 , and the inner wall has a plurality of cooling holes that open to the interior side of the outer wall.6. An airfoil comprising:first and second endwall sections; an outer wall ...

COOLING HOLE WITH ENHANCED FLOW ATTACHMENT

A gas turbine engine component includes a wall having first and second wall surfaces, a cooling hole extending through the wall and a convexity. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally and laterally from the metering section and a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section. The convexity is located near the outlet. 1. A gas turbine engine component comprising:a wall having first and second wall surfaces; an inlet located at the first wall surface;', 'an outlet located at the second wall surface;', 'a metering section extending downstream from the inlet; and', a first lobe diverging longitudinally and laterally from the metering section; and', 'a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section; and, 'a diffusing section extending from the metering section to the outlet, the diffusing section comprising], 'a cooling hole extending through the wall and comprisinga convexity located near the outlet.2. The component of claim 1 , further comprising:an indentation located downstream from the convexity on the second wall surface and downstream from the outlet.3. The component of claim 2 , wherein the indentation has a lateral width that is greater than or equal to a lateral width of the outlet.4. The component of claim 2 , wherein the indentation has a lateral width that is between about 100% and about 150% of the lateral width of the convexity.5. The component of claim 2 , wherein the indentation has a longitudinal length that is substantially equal to a longitudinal length of the convexity.6. The component of claim 1 , wherein the diffusing section further comprises:an interlobe region ...

GAS TURBINE ENGINE STATOR VANE BAFFLE ARRANGEMENT

A method of flowing cooling fluid through a stator vane in a gas turbine engine includes the step of providing an airfoil that has an exterior wall that provides a cooling cavity. The exterior surface has an interior surface that has multiple pin fins that extend therefrom. A baffle is arranged in the cooling cavity and supported by the pin fins. A perimeter cavity is provided between the baffle and the exterior wall. The pin fins are arranged in the perimeter cavity. Cooling fluid flows through a region in the perimeter cavity. The pin fins are arranged in the region having a low Reynolds number and through which the cooling fluid 1. A method of flowing cooling fluid through a stator vane in a gas turbine engine , comprising the steps of:providing an airfoil having an exterior wall providing a cooling cavity, the exterior surface has an interior surface having multiple pin fins extending therefrom;providing a baffle arranged in the cooling cavity and supported by the pin fins, wherein a perimeter cavity is provided between the baffle and the exterior wall, the pin fins arranged in the perimeter cavity; andflowing cooling fluid through a region in the perimeter cavity, wherein the pin fins are arranged in the region having a low Reynolds number and through which the cooling fluid flows.2. The method according to claim 1 , wherein the baffle is sheet steel.3. The method according to claim 2 , wherein the exterior wall provides pressure and suction sides joined at leading and trailing edges claim 2 , and the baffle includes impingement holes configured to provide impingement cooling fluid onto the exterior wall at the leading edge.4. The method according to claim 2 , wherein the baffle includes a generally smooth outer contour free of protrusions.5. The method according to claim 4 , wherein the outer contour is provided by plastic deformation.6. The method according to claim 4 , wherein cooling holes are provided by at least one of drilling claim 4 , laser drilling ...

BONDED MULTI-PIECE GAS TURBINE ENGINE COMPONENT

A gas turbine engine component includes an airfoil that has first and second structural airfoil segments that are bonded to each other in at least one diffusion joint. The first and second structural airfoil segments are formed of, respectively, first and second materials. The first and second materials are: different base-metal metallic alloys, a metallic alloy and a ceramic-based material, or ceramic-based materials that differ by at least one of composition and microstructure. The first structural airfoil segment is a first skin and the second structural airfoil segment is a hollow core that has an airfoil shape. 1. A gas turbine engine component comprising: different base-metal metallic alloys,', 'a metallic alloy and a ceramic-based material, or', 'ceramic-based materials that differ by at least one of composition and microstructure,, 'an airfoil including first and second structural airfoil segments that are bonded to each other in at least one diffusion joint, the first and second structural airfoil segments being formed of, respectively, first and second materials, wherein the first and second materials arewherein the first structural airfoil segment is a first skin and the second structural airfoil segment is a hollow core that has an airfoil shape.2. The gas turbine engine component as recited in claim 1 , wherein each of the first and second structural airfoil segments includes a respective wall having an exterior surface and an opposed claim 1 , interior surface claim 1 , with a plurality of ribs extending from the interior surface.3. The gas turbine engine component as recited in claim 1 , wherein the first skin includes a wall having an external surface and an opposed claim 1 , interior surface claim 1 , with a plurality of spaced-apart protrusions extending from the interior surface claim 1 , and free ends of the plurality of spaced-apart protrusions are bonded to the hollow core in diffusion joints.4. The gas turbine engine as recited in claim 3 , ...

TURBINE CLEARANCE CONTROL SYSTEM AND METHOD FOR IMPROVED VARIABLE CYCLE GAS TURBINE ENGINE FUEL BURN

A method of assembling a gas turbine engine includes setting a build clearance at assembly in response to a running tip clearance defined with a cooled cooling air. A method of operating a gas turbine engine includes supplying a cooled cooling air to a high pressure turbine in response to an engine rotor speed. 120-. (canceled)21. A gas turbine engine comprising:a flow circuit from a second stream airflow path of the gas turbine engine to communicate a portion of an airflow from said second stream airflow path to a heat exchanger;a flow circuit from said heat exchanger to eject said portion of said airflow of said second stream airflow path into a third stream airflow path;a flow circuit from a primary airflow path of the gas turbine engine to communicate a portion of said core airflow from said primary airflow path to said heat exchanger; anda flow circuit from said heat exchanger to eject said portion of said core airflow from said cooled cooling air system as a cooled cooling airflow to a high pressure turbine section of the gas turbine engine to limit a transient pinch event thereby reducing a steady state radial tip clearance during engine operation.22. The gas turbine engine as recited in claim 21 , wherein said flow circuit from said primary airflow path of the gas turbine engine communicates with a diffuser in a combustor section.23. The gas turbine engine as recited in claim 21 , wherein the radial running tip clearance is defined between a turbine airfoil and a shroud assembly during engine operation.24. The gas turbine engine as recited in claim 21 , further comprising selectively supplying the cooled cooling air in response to an engine rotor speed.25. The gas turbine engine as recited in claim 21 , wherein the second stream airflow from the second stream airflow path is ejected from the air-to-air heat exchanger system to the third stream airflow path.26. The gas turbine engine as recited in claim 21 , wherein the transient pinch event occurs when the ...

INTERCOOLED COOLING AIR USING COOLING COMPRESSOR AS STARTER

A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. The cooling compressor is connected to be driven with at least one rotor in the main compressor section. A source of pressurized air is selectively sent to the cooling compressor to drive a rotor of the cooling compressor to rotate, and to in turn drive the at least one rotor of the main compressor section at start-up of the gas turbine engine. An intercooling system is also disclosed. 1. A gas turbine engine comprising;a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations;a turbine section having a high pressure turbine;a tap connected for tapping air from at least one of the more upstream locations in the compressor section, and connected for passing the tapped air through a heat exchanger and then to a cooling compressor, the cooling compressor for compressing air downstream of the heat exchanger, and delivering air into at least one of the high pressure turbine and the high pressure compressor;the cooling compressor being connected to be driven with at least one rotor in the main compressor section, and a source of pressurized air that is selectively sent to the cooling compressor, to drive a rotor of the cooling compressor to rotate, to in turn drive the at least one rotor of the main compressor section at start-up of the gas turbine engine; andwherein the rotor of the cooling compressor includes a centrifugal compressor impeller.2. The gas turbine engine as set forth in claim 1 , wherein air temperatures at the ...

FLOW DIRECTING COVER FOR ENGINE COMPONENT

An assembly for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end. The assembly further includes a cover having at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway. 1. An assembly for a gas turbine engine , comprising:an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end; anda cover formed separately from the airfoil and including at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway.2. The assembly as recited in claim 1 , wherein:the cover includes a first turning cavity and a second turning cavity;the airfoil includes a third passageway having an inlet at the radial end;the first turning cavity is configured to direct a first portion of the fluid expelled from the outlet of the first passageway into the inlet of the second passageway; andthe second turning cavity is configured to direct a second portion of the fluid expelled from the outlet of the first passageway into the inlet of the third passageway.3. The assembly as recited in claim 2 , wherein the cover includes a flow divider between the first turning cavity and the second turning cavity.4. The assembly as recited in claim 3 , wherein claim 3 , when viewed from an interior of the cover claim 3 , the flow divider is substantially convex and the first and second turning cavities are substantially concave.5. The assembly as recited in claim 4 , wherein the first and second turning cavities are substantially semi-circular in cross-section.6. The assembly as recited in claim 2 , wherein the first passageway is inward of the second and ...

FLUID SUPERCHARGING DEVICE

A fluid supercharging device, comprising: a rotating shaft; a vane disc coaxially fixed to the rotating shaft; a plurality of fan blades fixed around a perimeter of the vane disc; the back side of the fan blades being provided with at least one fluid guiding inlet, an end of the back side distal from the vane disc is provided with a fluid guiding outlet, a fluid channel communicating the fluid guiding inlet with the fluid guiding outlet is provided along a lengthwise direction inside the fan blades; the fan blades rotate to generate a centrifugal force such that a fluid flows into the fluid channel via the fluid guiding inlet on the back side, and flows out of the fluid guiding outlet along the lengthwise direction of the fan blades. 1. A fluid supercharging device , comprising:a rotating shaft for rotating under action of a driving force;at least one stage of an impeller coaxially fixed to the rotating shaft;a plurality of fan blades fixed around a perimeter of each of the at least one stage of the impeller;wherein a side of each of the plurality of the fan blades facing toward fluid is a windward side and another side of each of the plurality of the fan blades backing on to the fluid is a leeward side, a plurality of fluid guiding inlets are arranged at several locations along a lengthwise direction extending from the impeller to a tip end of the leeward side of each of the plurality of the fan blades, and a fluid channel communicating with the plurality of the fluid guiding inlets is provided along the lengthwise direction inside each of the plurality of the fan blades;wherein fluid supercharging device further comprises a suction motor, a suction port of the suction motor is in communication with the fluid channel;wherein when the plurality of the fan blades rotate, the fluid flows into the fluid channel via the plurality of the fluid guiding inlets on the leeward side, and the suction motor is configured to generate a suction force to enable the suction port to ...

INTERCOOLED COOLING AIR WITH DUAL PASS HEAT EXCHANGER

A gas turbine engine comprises a main compressor section having a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses ng air downstream of the heat exchanger, and delivers air into the high pressure turbine. The heat exchanger has at least two passes, with one of the passes passing air radially outwardly, and a second of the passes returning the air radially inwardly to the compressor. An intercooling system for a gas turbine engine is also disclosed. 120.-. (canceled)21. An intercooling system for a gas turbine engine comprising:a heat exchanger for cooling air drawn from a portion of a main compressor section at a first temperature and pressure for cooling the air to a second temperature cooler than the first temperature;a cooling compressor compressing air communicated from the heat exchanger to a second pressure greater than the first pressure and communicating the cooling air to a portion of at least a turbine section; andsaid heat exchanger having at least two passes, with a first of said passes passing air in a direction having at least a radially outward component, and a second of said passes returning the air in a direction having at least a radially inward component to the compressor.22. The intercooling system as set forth in claim 21 , wherein said first pass is positioned downstream of said second pass in said bypass duct.23. The intercooling system as set forth in claim 22 , wherein said cooling compressor includes a centrifugal compressor impeller.24. The intercooling system as set forth in claim 22 , wherein a main fan delivers bypass air into a bypass duct and into said main compressor section and said heat exchanger positioned within said bypass duct to be cooled by bypass air.25. The intercooling ...

COOLING HOLE WITH SHAPED METER

A gas turbine engine component having a cooling passage includes a first wall defining an inlet of the cooling passage, a second wall generally opposite the first wall and defining an outlet of the cooling passage, a metering section extending downstream from the inlet, and a diffusing section extending from the metering section to the outlet. The metering section includes an upstream side and a downstream side generally opposite the upstream side. At least one of the upstream and downstream sides includes a first passage wall and a second passage wall where the first and second passage walls intersect to form a V-shape. 1. A wall located in a gas turbine engine , the wall comprising:first and second surfaces; and [ a longitudinal first side; and', 'a longitudinal second side opposite the longitudinal first side, wherein the longitudinal first side comprises a first passage wall and a second passage wall, and wherein the second longitudinal side comprises a third passage wall and a fourth passage wall;, 'a metering section commencing at the inlet, the metering section comprising, 'a diffusing section in communication with the metering section and terminating at the outlet;', the first and second passage walls are straight and intersect to form a vertex; and', 'the third passage wall is curved, extending from the first passage wall; and', 'the fourth passage wall is curved, extending from the second passage wall to intersect the third passage wall., 'wherein, viewing the metering section from the diffusion section], 'a cooling passage extending between an inlet at the first surface and an outlet at the second surface, the cooling passage comprising2. The wall of claim 1 , wherein the longitudinal first side is an upstream side of the cooling passage.3. The wall of claim 1 , wherein the longitudinal first side is a downstream side of the cooling passage.4. The wall of claim 1 , wherein the first passage wall and the second passage wall intersect to form an angle at ...

Flared central cavity aft of airfoil leading edge

A blade includes an airfoil defined by a pressure side outer wall and a suction side outer wall connecting along leading and trailing edges and form a radially extending chamber for receiving a coolant flow. A rib configuration may include: a leading edge transverse rib connecting the pressure side outer wall and the suction side outer wall and partitioning the radially extending chamber into a leading edge passage within the leading edge of the airfoil and a central passage adjacent to the leading edge passage. One or both camber line ribs connect to a corresponding pressure side outer wall and suction side outer wall at a point aft of the leading edge transverse rib causing the central passage to extend towards one or both of the pressure side outer wall and the suction side outer wall, resulting in a flared center cavity aft of the leading edge.

AIRFOIL WITH SEAL BETWEEN ENDWALL AND AIRFOIL SECTION

An airfoil includes an endwall section and an airfoil section that defines, at least in part, an airfoil profile. The airfoil section includes an internal passage and a rib that sub-divides the internal passage. At least one of the rib or the endwall section includes a seal cavity. A seal is disposed in the seal cavity. 1. An airfoil comprising:an endwall section;an airfoil section defining, at least in part, an airfoil profile, the airfoil section including an internal passage and a rib sub-dividing the internal passage, at least one of the rib or the endwall section including a seal cavity; anda seal disposed in the seal cavity.2. The airfoil as recited in claim 1 , wherein the seal is rigidly attached with the other one of the rib or the endwall section.3. The airfoil as recited in claim 1 , wherein the seal cavity is in the rib claim 1 , and the seal cavity opens radially.4. The airfoil as recited in claim 3 , wherein the seal is rigidly attached with the endwall section.5. The airfoil as recited in claim 1 , wherein the rib is radially elongated.6. The airfoil as recited in claim 5 , wherein the seal cavity is in a radial face of the rib.7. The airfoil as recited in claim 6 , wherein the seal cavity is in an enlarged radial end of the rib.8. The airfoil as recited in claim 1 , wherein the airfoil section is formed of ceramic and the seal is formed of metal.9. A gas turbine engine comprising:a compressor section;a combustor in fluid communication with the compressor section; anda turbine section in fluid communication with the combustor, an airfoil section defining, at least in part, an airfoil profile, the airfoil section including an internal passage and a rib sub-dividing the internal passage, at least one of the rib or the endwall section including a seal cavity, and', 'a seal disposed in the seal cavity., 'at least one of the turbine section or the compressor section including an airfoil having an endwall section,'}10. The gas turbine engine as recited in ...

GAS TURBINE COMPONENT WITH PLATFORM COOLING

A gas turbine engine component includes an airfoil that has a leading edge and extends from a first side of a platform that has a leading edge overhang. A root portion extends from a second side of the platform opposite the first side. A cooling passage extends through the platform and beneath a trailing edge of the airfoil. The cooling passage includes an inlet located on a second opposite side of the platform. The inlet is located axially upstream of the leading edge of the airfoil and the inlet is located axially upstream of the root portion in the leading edge overhang. 1. A gas turbine engine component comprising:an airfoil having a leading edge and extending from a first side of a platform having a leading edge overhang;a root portion extending from a second side of the platform opposite the first side; anda cooling passage extending through the platform and beneath a trailing edge of the airfoil, the cooling passage including an inlet located on a second opposite side of the platform, wherein the inlet is located axially upstream of the leading edge of the airfoil and the inlet is located axially upstream of the root portion in the leading edge overhang.2. The component of claim 1 , wherein the cooling passage includes an outlet located on the first side of the platform.3. The component of claim 1 , wherein the cooling passage extends beneath a portion of the airfoil.4. The component of claim 3 , wherein a first portion of the cooling passage is located on a pressure side of the platform and a second portion of the cooling passage is located on a suction side of the platform.5. The component of claim 1 , wherein the cooling passage includes a plurality of cooling structures extending from a radially inner side of the cooling passage to a radially outer side of the cooling passage.6. The component of claim 1 , wherein the component is a turbine blade for a gas turbine engine.7. The component of claim 1 , wherein the cooling passage includes an outlet located ...

INTEGRATED INNER CASE HEAT SHIELD

A mid-turbine frame for a gas turbine engine according to an example of the present disclosure includes, among other things, a first frame case, a flange coupled to the first frame case, and a heat shield adjacent to the flange. A method of cooling a portion of a gas turbine engine is also disclosed. 1. A mid-turbine frame for a gas turbine engine , comprising:a first frame case;a flange coupling a conduit to the first frame case;a heat shield adjacent to the flange and spaced apart from adjacent spokes, the heat shield arranged between an airfoil receiving the conduit and surfaces of the flange, the airfoil defining an airfoil cavity, and the heat shield defining an opening for receiving the conduit; andwherein the heat shield defines a cooling cavity, the airfoil cavity and the cooling cavity being separated by the heat shield, the cooling cavity has an inlet and an outlet, the inlet bounded by the first frame case such that the inlet establishes a first fluid passage between the airfoil cavity and the cooling cavity, the outlet defined between an outer perimeter of the conduit and an inner perimeter of the opening such that the opening establishes a second fluid passage along exposed surfaces of the inner perimeter that extends between the cooling cavity and the airfoil cavity.2. The mid-turbine frame as recited in claim 1 , wherein the inlet follows an outer perimeter of the heat shield such that the first fluid passage is defined between the outer perimeter of the heat shield and the first frame case.3. The mid-turbine frame as recited in claim 2 , wherein the heat shield includes a first portion and a second portion claim 2 , the first portion extends in a circumferential direction with respect to a longitudinal axis defined by the first frame case claim 2 , and the second portion abuts the first frame case and extends in a radial direction with respect to the longitudinal axis.4. The mid-turbine frame as recited in claim 3 , wherein the outlet has a ring ...

INTERNALLY COOLED AIRFOIL

A casting core and/or an airfoil may comprise a tip flag cavity having a forward pedestal and a first spear pedestal disposed aft of the forward pedestal. A trailing edge discharge cavity may be separated from the tip flag cavity and include a first row of pedestals. The first row of pedestals may comprise a first racetrack pedestal. A second row of pedestals may be disposed aft of the first row of pedestals and include a second racetrack pedestal. A third row of pedestals may be disposed aft of the second row of pedestals and include a circular pedestal. A fourth row of pedestals may be disposed aft of the third row of pedestals and include a second spear pedestal. 1. A casting core , comprising:a tip flag cavity comprising a forward pedestal and a first spear pedestal disposed aft of the forward pedestal; and a first row of pedestals comprising a first racetrack pedestal,', 'a second row of pedestals aft of the first row of pedestals, the second row of pedestals comprising a second racetrack pedestal,', 'a third row of pedestals aft of the second row of pedestals, the third row of pedestals comprising a circular pedestal, and', 'a fourth row of pedestals aft of the third row of pedestals, the fourth row of pedestals comprising a second spear pedestal., 'a trailing edge discharge cavity separated from the tip flag cavity and comprising,'}2. The casting core of claim 1 , wherein a diameter of the first racetrack pedestal is equal to a diameter of the second racetrack pedestal.3. The casting core of claim 2 , wherein a diameter of the first spear pedestal is greater than the diameter of the first racetrack pedestal.4. The casting core of claim 3 , wherein the diameter of the first spear pedestal is 0.026 inches.5. The casting core of claim 1 , wherein the first row of pedestals comprises 12 racetrack pedestals claim 1 , and the fourth row of pedestals comprises 24 spear pedestals.6. The casting core of claim 1 , wherein a first blockage of the first row of pedestals ...

COOLANT FLOW REDIRECTION COMPONENT

A gas turbine engine includes a compressor section, a combustor fluidly connected to the compressor section, and a turbine section fluidly connected to the combustor and mechanically connected to the compressor section via a shaft. Multiple rotors are disposed in one of the compressor section and the turbine section. Each of the rotors includes a rotor disk portion having a radially inward bore, and is static relative to the shaft. Each rotor is axially adjacent at least one other rotor and a gap is defined between each rotor and an adjacent rotor. A cooling passage for a cooling flow is defined between the shaft and the rotors, and a cooling flow redirection component is disposed at the gap and is operable to redirect the cooling flow in the cooling passage into the gap. 1. A gas turbine engine comprising:a compressor section;a combustor fluidly connected to the compressor section;a turbine section fluidly connected to the combustor and mechanically connected to said compressor section via a shaft;a plurality of rotors disposed in one of said compressor section and said turbine section, each of said rotors including a rotor disk portion including a radially inward bore, and each of said rotors being static relative to said shaft;each rotor in said plurality of rotors being axially adjacent at least one other of said rotors in said plurality of rotors and defining a gap between each of said rotors and said axially adjacent rotors, wherein each of said rotors includes a joint arm partially crossing said gap;a cooling passage for a cooling flow defined between said shaft and said rotors; anda cooling flow redirection component disposed at said gap and operable to redirect said cooling flow in said cooling passage into said gap, wherein the cooling flow redirection component is distinct from said rotors in said plurality of rotors.2. The gas turbine engine of claim 1 , wherein said gap is defined between radially aligned surfaces of adjacent rotor disks.3. The gas ...

INDUSTRIAL VAPOR GENERATOR FOR DEPOSITING AN ALLOY COATING ON A METAL STRIP

The present invention relates to a vacuum deposition facility for depositing a metal alloy coating on a substrate (), said facility being equipped with a vapour generator/mixer comprising a vacuum chamber () in the form of an enclosure provided with means for creating a vacuum state therein relative to the external environment and provided with means for the entry and exit of the substrate (), while still being essentially sealed from the external environment, said enclosure including a vapour deposition head, called the injector (), configured so as to create a jet of metal alloy vapour of sonic velocity towards the surface of the substrate () and perpendicular thereto, said ejector () being in sealed communication with a separate mixer device (), which is itself connected upstream to at least two crucibles () respectively, these containing different metals M and M in liquid form, each crucible () being connected to the mixer () by its own pipe (). 17767373141112121112144414. A method of depositing a metal alloy coating on a substrate () , comprising using a facility for depositing under vacuum a metal alloy coating on a substrate () , equipped with a vapor generator-mixer comprising a vacuum chamber () in the form of an enclosure , provided with means for ensuring a vacuum state therein relative to the external environment and provided with means for the inlet and outlet of the substrate () , while still being substantially sealed relative to the external environment , said enclosure including a vapor deposition head , called ejector () , configured so as to create a jet of metal alloy vapor at sonic velocity towards the surface of the substrate () and perpendicular thereto , said ejector () being in sealed communication with a separate mixer device () , which is itself connected upstream to at least two crucibles ( ,) , respectively , and comprising different metals M and M in liquid form , each crucible ( ,) being connected to the mixer () by its own pipe ( ,′) ...

Airfoil assembly with spacer and tie-spar

An airfoil assembly includes at least one airfoil that has a hollow interior. First and second platforms are disposed between the airfoil. At least one tie-spar extends along an axis through the first platform, the hollow interior of the airfoil, and the second platform. There is a thermal expansion difference between a thermal expansion of the tie-spar in the axial direction and the combined thermal expansion of the airfoil and the first and second platform in the axial direction. At least one spacer portion is arranged on the tie-spar. The spacer portion has a thermal expansion in the axial direction that is greater than the thermal expansion difference such that the spacer portion maintains the tie-spar under tension and clamps the first and second platforms on the airfoil.

PROCESS FOR PRODUCING A THERMAL BARRIER IN A MULTILAYER SYSTEM FOR PROTECTING A METAL PART AND PART EQUIPPED WITH SUCH A PROTECTIVE SYSTEM

The object of the present invention is to produce a metal part equipped with a protection system, particularly for turbine blades for aircraft engines, having a thermal barrier that is improved in terms of thermal properties, adhesion to the part and resistance to oxidation/corrosion. In order to achieve this, the method according to the invention produces in a single step, from specific ceramics, coating layers using SPS technology. 113-. (canceled)14: A metal part made of superalloy , equipped with a thermal barrier being formed by a flash sintering operation in an SPS machine enclosure , a substrate, and', 'a thermal barrier comprising at least two ceramic layers consisting of zirconium-based refractory ceramic, at least one layer is an inner ceramic layer of zirconium-based refractory ceramic, and at least one layer is an outer ceramic layer of zirconium-based refractory ceramic having an outer face and being disposed over the inner ceramic layer,, 'wherein the metal part comprises'}wherein the thermal barrier has a porosity gradient, with porosity increasing from the outer ceramic layer to the inner ceramic layer with regard to the metal part, (i) the outer ceramic layer has at least one physicochemical resistance property to calcium-magnesium-aluminum-silicate oxide pollutants which is higher than that of the inner ceramic layer, or', '(ii) the outer ceramic layer has a higher thermal resistance property than that of the inner ceramic layer, 'wherein'}wherein the thermal barrier formed by the flash sintering operation is a monolayer having a gradient of properties from the metal part to the outer face of the outer ceramic layer corresponding to the gradient of initial properties of the inner ceramic layer and the outer ceramic layer, andwherein the inner ceramic layer has a higher thermal expansion coefficient than the outer ceramic layer.15: The metal part according to claim 14 , wherein the substrate comprises a nickel-based superalloy.16: The metal part ...

ADDITIVELY MANUFACTURED CORE

A method of preparing a casting article for use in manufacturing a gas turbine engine part according to an exemplary aspect of the present disclosure includes, among other things, communicating a powdered material to an additive manufacturing system and preparing a casting article that includes at least one trunk and a skin core that extends from the at least one trunk out of the powdered material. A casting article is also disclosed. 1. A casting article for a gas turbine engine , comprising:a core having the core geometry including at least a trunk that forms a mainbody cooling passage of a cast gas turbine engine part and a skin core that forms an internal cooling feature of the cast gas turbine engine part, the skin core extending from the trunk;wherein the core includes multiple layers of a powdered material deposited onto one another, the multiple layers are joined to one another, and the multiple layers include at least a portion of a second layer of the powdered material that is melted to adhere the second layer to a first layer; andwherein each of the trunk and the skin core includes a respective vascular network, the vascular network includes a plurality of spherical nodes and a plurality of elongated branches that extend between the plurality of nodes to establish a lattice structure, and the vascular network is a hollow vascular network structure having interconnected internal hollow passages extending inside of the plurality of nodes and the plurality of branches.2. The casting article as recited in claim 1 , wherein the powdered material includes at least one of a refractory metal claim 1 , a silica and an alumina.3. The casting article as recited in claim 1 , wherein a portion of the core geometry is at least partially filled with the powdered material.4. The casting article as recited in claim 1 , wherein the trunk includes an outer shell body defining a hollow interior claim 1 , and the hollow interior is at least partially filled with the powdered ...

Airfoil Cooling Using Non-Line of Sight Holes

An airfoil for a gas turbine engine is provided that includes a first portion formed from a first plurality of plies of a ceramic matrix composite material and defining an inner surface of the airfoil, as well as a second portion formed from a second plurality of plies of a ceramic matrix composite material and defining an outer surface of the airfoil. The first portion and the second portion define a non-line of sight cooling aperture extending from the inner surface to the outer surface of the airfoil. In one embodiment, a surface angle that is less than 45° is defined between a second aperture and the outer surface. A method for forming an airfoil for a gas turbine engine also is provided. 120.-. (canceled)21. An airfoil for a gas turbine engine , the airfoil comprising:a first portion defining a first aperture and an inner surface of the airfoil, the first portion formed from a first plurality of plies of a ceramic matrix composite material; anda second portion defining a second aperture and an outer surface of the airfoil, the second portion formed from a second plurality of plies of a ceramic matrix composite material,wherein the first aperture and the second aperture define a non-line of sight cooling aperture extending from the inner surface to the outer surface of the airfoil,wherein the first aperture has a first end defined at the inner surface and a second end defined at the second portion,wherein the second aperture has a first end defined adjacent the second end of the first aperture and a second end defined at the outer surface,wherein the second end of the first aperture has a cross-sectional area that differs from a cross-sectional area of the first end of the second aperture.22. The airfoil of claim 21 , wherein the second end of the first aperture has a cross-sectional shape that differs from a cross-sectional shape of the first end of the second aperture.23. The airfoil of claim 21 , wherein the first aperture has a first minimum cross-sectional ...

BLADE OUTER AIR SEAL FIN COOLING ASSEMBLY AND METHOD

A blade outer air seal according to an exemplary aspect of the present disclosure includes, among other things, a body to be distributed circumferentially about a blade array. The body has a plurality of grooves, which can, for example, improve the aerodynamic efficiency of a turbine. A fin is between a first groove and a second groove of the plurality of grooves. The fin extends radially from the body and terminates at a radially inner fin face that provides one or more cooling outlets. A sacrificial structure for forming internal cooling passages within a blade outer air seal and a method of cooling an interface between a blade outer air seal and a rotating blade array is also disclosed. 1. A sacrificial structure for forming internal cooling passages within a blade outer air seal , comprising:a refractory metal core to form an internal cooling passage within a fin of a blade outer air seal;wherein the fin extends between an opposed pair of grooves in a body of the blade outer air seal, the fin extends radially from the body and terminates at a radially inner fin face having one or more cooling outlets;wherein the internal cooling passage extends from a cooling cavity, through the fin, to at least one of the one or more cooling outlets; andwherein the refractory metal core includes a first radially extending portion dimensioned to extend from the cooling cavity, a second radially extending portion dimensioned to extend from a respective one of the one or more cooling outlets, and a circumferentially extending portion connecting the first and second radially extending portions such that the first radially extending portion is circumferentially offset from the second radially extending portion.2. The sacrificial structure of claim 1 , wherein the refractory metal core extends to a radially inner core face to align with the radially inner fin face.3. The sacrificial structure of claim 2 , wherein the fin is one of a plurality of fins extending radially from the body. ...

GAS TURBINE ENGINES WITH IMPROVED LEADING EDGE AIRFOIL COOLING

An airfoil for a gas turbine engine includes a body with a first side wall and a second side wall joined at a leading edge and a trailing edge, the first side wall having a first interior surface and the second side wall having a second interior surface. The airfoil further includes an internal wall disposed within of the body and extending between the first interior surface and the second interior surface to define a supply passage and a leading edge passage. The internal wall defines a plurality of cooling holes to direct cooling air from the supply passage to the leading edge passage such that the cooling air impinges upon the leading edge. The airfoil further includes a first plurality of grooves formed in the first interior surface, each the first plurality of grooves extending in a chordwise direction within the leading edge passage. 1. An airfoil for a gas turbine engine , comprising:a body comprising a first side wall and a second side wall joined at a leading edge and a trailing edge, the first side wall having a first interior surface and the second side wall having a second interior surface;an internal wall disposed within the body and extending between the first interior surface and the second interior surface to define a supply passage and a leading edge passage,wherein the internal wall defines a plurality of cooling holes to direct cooling air from the supply passage to the leading edge passage such that the cooling air impinges upon the leading edge; anda first plurality of grooves formed into the first interior surface by removing portions of the first interior surface such that the first side wall has a reduced thickness within respective grooves of the first plurality of grooves relative to areas in between the respective grooves and areas along and adjacent to the leading edge, each of the first plurality of grooves extending in a chordwise direction within the leading edge passage, wherein each of the first plurality of grooves is formed by a ...

MULTI-SOURCE TURBINE COOLING AIR

A gas turbine engine comprises a compressor section and a turbine section, with the turbine section having a first stage blade row and a downstream blade row. A higher pressure tap is tapped from a higher pressure first location in the compressor. A lower pressure tap is tapped from a lower pressure location in the compressor which is at a lower pressure than the first location. The higher pressure tap passes through a heat exchanger, and then is delivered to cool the first stage blade row in the turbine section. The lower pressure tap is delivered to at least partially cool the downstream blade row. 1. A gas turbine engine comprising:a compressor section and a turbine section, with said turbine section having a first stage blade row and a downstream blade row;a higher pressure tap tapped from a higher pressure first location in said compressor; anda lower pressure tap tapped from a lower pressure location in said compressor which is at a lower pressure than said first location, said higher pressure tap passing through a heat exchanger, and then being delivered to cool said first stage blade row in said turbine section, and said lower pressure tap being delivered to at least partially cool said downstream blade row.2. The gas turbine engine as set forth in claim 1 , wherein said higher pressure tap passing from said heat exchanger toward said turbine section claim 1 , and split into a first path heading radially outwardly to cool an upstream end of said first stage blade row claim 1 , and a second path moving radially inwardly of a hub mounting said first stage blade row and then moving radially outwardly to cool a downstream end of said first stage blade row.3. The gas turbine engine as set forth in claim 2 , wherein radially outwardly extending air from said higher pressure tap also cooling a vane mounted intermediate said first stage blade row and said downstream blade row.4. The gas turbine engine as set forth in claim 4 , wherein said radially outwardly ...

GEARED TURBOFAN ARRANGEMENT WITH CORE SPLIT POWER RATIO

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including a low pressure compressor and a second compressor section, and a turbine section including a fan drive turbine and a high pressure turbine. The fan drive turbine drives the low pressure compressor and a gear arrangement to drive the fan section. A core split power ratio is provided by power input to the high pressure compressor divided by a power input to the low pressure compressor measured in horsepower. 1. A turbofan engine comprising:a fan section including a plurality of fan blades and a fan case surrounding the fan blades to define a bypass duct, the fan section defining a fan pressure ratio of less than 1.40 across the fan blade alone, the fan section having a low corrected fan tip speed of less than 1150 ft/sec, and the low corrected fan tip speed being an actual fan tip speed divided by [(Tram ° R)/(518.7° R)]{circumflex over ( )}0.5;a gear arrangement including a gear reduction ratio greater than 2.3;a compressor section including a low pressure compressor and a high pressure compressor, the low pressure compressor including 3 or more stages, and the high pressure compressor including greater than or equal to 8 stages and less than or equal to stages;a turbine section including a fan drive turbine and a two-stage high pressure turbine, the high pressure turbine driving the high pressure compressor, the fan drive turbine driving the low pressure compressor and driving the gear arrangement to drive the fan section, and the fan drive turbine including between three and six stages;wherein the fan drive turbine includes an inlet, an outlet, and a fan drive turbine pressure ratio greater than 5, and the fan drive turbine pressure ratio is pressure measured prior to the inlet as related to pressure at the outlet prior to any exhaust nozzle;wherein a core split power ratio is provided by a second power input to the ...

COOLING SYSTEM

A thermal management system for a gas turbine engine and/or an aircraft is provided including a thermal transport bus having a heat exchange fluid flowing therethrough. The thermal management system also includes a plurality of heat source exchangers and at least one heat sink exchanger. The plurality of heat source exchangers and the at least one heat sink exchanger are in thermal communication with the heat exchange fluid in the thermal transport bus. The plurality of heat source exchangers are arranged along the thermal transport bus and configured to transfer heat from one or more accessory systems to the heat exchange fluid, and the at least one heat sink exchanger is located downstream of the plurality of heat source exchangers and configured to remove heat from the heat exchange fluid. 1. A thermal management system for incorporation at least partially into at least one of a gas turbine engine or an aircraft , the thermal management system comprising:a thermal transport bus having a heat exchange fluid flowing therethrough;a pump for generating a flow of the heat exchange fluid in the thermal transport bus;a plurality of heat source exchangers in thermal communication with the heat exchange fluid in the thermal transport bus, the plurality of heat source exchangers arranged along the thermal transport bus; andat least one heat sink exchanger permanently or selectively in thermal communication with the heat exchange fluid in the thermal transport bus at a location downstream of the plurality of heat source exchangers.2. The thermal management system of claim 1 , wherein the plurality of heat source exchangers are configured to transfer heat from an accessory system of the gas turbine engine to the heat exchange fluid in the thermal transport bus claim 1 , and wherein the at least one heat sink exchanger is configured to remove heat from the heat exchange fluid in the thermal transport bus.3. The thermal management system of claim 1 , wherein the at least one ...

ADDITIVE MANUFACTURING METHOD FOR MAKING HOLES BOUNDED BY THIN WALLS IN TURBINE COMPONENTS

A method of forming a passage in a turbine component includes: using an additive manufacturing process to form a first support structure on a first surface of the turbine component; forming a second support structure on a second surface of the turbine component, the second support structure being spaced apart from the first support structure; and forming a passage in the turbine component between the first and second support structures. 1. A method of forming a passage in a turbine component , comprising:using an additive manufacturing process to form a first support structure on a first surface of the turbine component;forming a second support structure on a second surface of the turbine component, the second support structure being spaced apart from the first support structure; andforming a passage in the turbine component between the first and second support structures.2. The method of wherein the additive manufacturing process comprises depositing powder on the first surface of the turbine component; and fusing the powder in a pattern corresponding to a layer of the first support structure.3. The method of further comprising repeating in a cycle the steps of depositing and fusing to build up the support structures in a layer-by-layer fashion.4. The method of further comprising forming the passage such that it is positioned closer to the second support structure than the first support structure.5. The method of further comprising removing at least a portion of the first support structure and the first surface.6. The method of further comprising removing at least a portion of the second support structure.7. The method of further comprising removing at least a portion of the first and second support structures.8. The method of further comprising completely removing the first and second support structures.9. A method of forming a passage in a turbine airfoil claim 1 , comprising:using an additive manufacturing process to form a first support structure on a first ...

Ceramic Matrix Composite Airfoil Cooling

Ceramic matrix composite airfoils for gas turbine engines are provided. In an exemplary embodiment, an airfoil includes opposite pressure and suction sides extending radially along a span. The pressure and suction sides define an outer surface of the airfoil. The airfoil further includes opposite leading and trailing edges extending radially along the span, the pressure and suction sides extending axially between the leading and trailing edges. The airfoil also includes a filler pack defining the trailing edge; the filler pack comprises a ceramic matrix composite material. Moreover, the airfoil includes a plenum defined within the airfoil for receiving a flow of cooling fluid, and a cooling passage defined within the filler pack for directing the flow of cooling fluid from the plenum to the outer surface of the airfoil. Methods for forming airfoils for gas turbine engines also are provided. 115.-. (canceled)16. A method for forming an airfoil for a gas turbine engine , the method comprising: opposite pressure and suction sides extending radially along a span,', 'opposite leading and trailing edges extending radially along the span, the pressure and suction sides extending axially between the leading and trailing edges, and', 'a plenum defined within the airfoil preform assembly; and, 'laying up a ceramic matrix composite material to form an airfoil preform assembly, the airfoil preform assembly including'}processing the airfoil preform assembly to produce the airfoil,wherein a cooling passage is defined within the airfoil, the cooling passage defined from the plenum to the trailing edge of the airfoil.17. The method of claim 16 , wherein the cooling passage comprises a crossover aperture defined from the plenum to a cavity defined by a plenum preform claim 16 , and wherein the cooling passage further comprises an ejection aperture defined within a filler pack preform from the trailing edge to the cavity.18. The method of claim 16 , further comprising:performing a ...

MODULATED TURBINE COOLING SYSTEM

A flow transfer apparatus for transferring cooling flow from a primary gas flowpath to a turbine rotor. The apparatus includes a first supply plenum communicating with the primary gas flowpath and first inducers, the first inducers configured to accelerate a first fluid flow from the first supply plenum and discharge it toward the rotor with a tangential velocity; a second supply plenum communicating with the primary gas flowpath and second inducers, the second inducers configured to accelerate a second fluid flow from the second supply plenum towards the rotor with a tangential velocity; and a cooling modulation valve operable to selectively permit or block the second fluid flow from the primary gas flowpath to the second supply plenum. The valve includes a flow control structure disposed in the primary gas flowpath and an actuation structure extending to a location radially outside of a casing defining the primary gas flowpath. 1. A method of transferring a cooling flow from a primary gas flowpath to a turbine rotor of a gas turbine engine that has at least two different operating conditions , the method comprising:during all conditions of engine operation, flowing a first fluid flow from the primary gas flowpath through a first supply plenum to a plurality of first inducers, each of the first inducers configured to accelerate the first fluid flow and discharge the first fluid flow toward the turbine rotor with a tangential velocity component; andduring some but not all of the operating conditions, flowing a second fluid flow from the primary gas flowpath through a cooling modulation valve and a second supply plenum to a plurality of second inducers, each of the second inducers configured to accelerate the second fluid flow and discharge the second fluid flow towards the turbine rotor with a tangential velocity component;wherein the cooling modulation valve comprises a flow control structure disposed in the primary gas flowpath and a valve actuation structure ...

PLASMA-RESISTANT MEMBER

According to an aspect of the invention, there is provided a plasma-resistant member including: a base member; and a layer structural component formed at a surface of the base member, the layer structural component including an yttria polycrystalline body and being plasma resistant, the layer structural component including a first uneven structure, and a second uneven structure formed to be superimposed onto the first uneven structure, the second uneven structure having an unevenness finer than an unevenness of the first uneven structure. 1. A plasma-resistant member , comprising:a base member; anda layer structural component formed at a surface of the base member, the layer structural component including an yttria polycrystalline body and being plasma resistant, a first uneven structure, and', 'a second uneven structure formed to be superimposed onto the first uneven structure, the second uneven structure having an unevenness finer than an unevenness of the first uneven structure., 'the layer structural component including'}2. The plasma-resistant member according to claim 1 , whereinthe first uneven structure has voids made in a portion of a surface of the layer structural component, the voids being where groups of crystal particles detached, andthe second uneven structure has an unevenness formed in the entire surface of the layer structural component, a size of the crystal particles of the unevenness being fine.3. The plasma-resistant member according to claim 1 , whereinthe arithmetic average Sa of a surface of the layer structural component is not less than 0.025 μm and not more than 0.075 μm,{'sup': 3', '2', '3', '2, 'the core material volume Vmc determined from a load curve of the surface of the layer structural component is not less than 0.03 μm/μmand not more than 0.08 μm/μm,'}{'sup': 3', '2', '3', '2, 'the core void volume Vvc determined from the load curve of the surface of the layer structural component is not less than 0.03 μm/μmand not more than 0.1 μ ...

FILLET OPTIMIZATION FOR TURBINE AIRFOIL

A blade for a gas turbine engine comprises an airfoil having a pressure side and a suction side, with a root and a tip wall. The pressure side and suction side extend beyond the tip wall to define a tip channel, defining a plurality of internal and external corners. The corners comprise fillets to define a thickness being greater than the thickness for the pressure, suction, or tip walls. A film hole can extend through the fillet, such that the length of the film hole at the fillet can be increased to define an increased length-to-diameter ratio for the film hole to improve film cooling through the film hole. 1. A blade for a gas turbine engine comprising:an airfoil having an outer wall defining a pressure side and a suction side, the outer wall extending chord-wise from a leading edge to a trailing edge and span-wise from a root toward a tip;a tip wall spanning the pressure side and the suction side of the outer wall and intersecting the outer wall to form at least one corner, with the outer wall having a first thickness at the corner and the tip wall having a second thickness at the corner;a cooling passage having a portion located along the tip wall and at least partially defined by the tip wall and the outer wall;a fillet located at the corner and having an effective radius of at least 1.5 times larger than the greater of the first and second thicknesses; andat least one film hole extending through the fillet to fluidly couple the cooling passage to an exterior of the airfoil.2. The blade according to wherein the at least one film hole includes an outlet provided on at least one of the pressure side or the suction side.3. The blade according to wherein the tip wall at least partially defines a tip channel between extensions of the pressure side and the suction side of the outer wall claim 1 , and the at least one film hole extends to the tip channel.4. The blade according to further comprising a tip shelf disposed within the outer wall wherein the at least one ...

GAS TURBINE ENGINE BUFFER SYSTEM

A gas turbine engine includes a buffer system that communicates a buffer cooling air to at least one bearing structure and at least one shaft of the gas turbine engine. The buffer system includes a first bleed air supply and a conditioning device that conditions the first bleed air supply to render the first buffer supply air at an acceptable temperature to pressurize the at least one bearing structure and cool the at least one shaft. 1. A gas turbine engine , comprising:a buffer system that communicates a buffer cooling air to at least one bearing structure and at least one shaft of the gas turbine engine, wherein said buffer system includes:a first bleed air supply; anda conditioning device that conditions said first bleed air supply to render said buffer cooling air at an acceptable temperature for pressurizing said at least one bearing structure and cooling said at least one shaft.2. The gas turbine engine as recited in claim 1 , wherein said at least one shaft is an inner shaft that interconnects a low pressure compressor and a low pressure turbine of the gas turbine engine.3. The gas turbine engine as recited in claim 1 , wherein said at least one shaft is an outer shaft that interconnects a high pressure compressor and a high pressure turbine of the gas turbine engine.4. The gas turbine engine as recited in claim 1 , wherein said at least one shaft includes an outer shaft that surrounds an inner shaft claim 1 , and said buffer cooling air is communicated between said inner shaft and said outer shaft.5. The gas turbine engine as recited in claim 4 , wherein said outer shaft is a tie shaft.6. The gas turbine engine as recited in claim 1 , wherein said buffer cooling air is communicated axially along an outer diameter of said at least one shaft.7. The gas turbine engine as recited in claim 1 , wherein said buffer cooling air is communicated axially through an inner diameter of said at least one shaft.8. The gas turbine engine as recited in claim 1 , wherein said ...

EROSION SUPPRESSION SYSTEM AND METHOD IN AN EXHAUST GAS RECIRCULATION GAS TURBINE SYSTEM

In an embodiment, a method includes flowing an exhaust gas from a turbine of a gas turbine system to an exhaust gas compressor of the gas turbine system via an exhaust recirculation path; evaluating moist flow parameters of the exhaust gas within an inlet section of the exhaust gas compressor using a controller comprising non-transitory media programmed with instructions and one or more processors configured to execute the instructions; and modulating cooling of the exhaust gas within the exhaust recirculation path, heating of the exhaust gas within the inlet section of the exhaust gas compressor, or both, based on the evaluation. 1. A method , comprising:flowing an exhaust gas from a turbine of a gas turbine system to an exhaust gas compressor of the gas turbine system via an exhaust recirculation path;evaluating moist flow parameters of the exhaust gas within an inlet section of the exhaust gas compressor using a controller comprising non-transitory media programmed with instructions and one or more processors configured to execute the instructions; andmodulating cooling of the exhaust gas within the exhaust recirculation path, heating of the exhaust gas within the inlet section of the exhaust gas compressor, or both, based on the evaluation.2. The method of claim 1 , wherein evaluating moist flow parameters of the exhaust gas within the inlet section of the exhaust gas compressor comprises estimating a projected droplet size and flux of the exhaust gas at a portion of the exhaust gas compressor based at least in part on the evaluation of the moist flow parameters of the exhaust gas claim 1 , wherein the cooling modulation of the exhaust gas and/or the heating modulation of the exhaust gas is made based at least partially on the projected droplet size and flux.3. The method of claim 2 , wherein evaluating moist flow parameters of the exhaust gas within the inlet section of the exhaust gas compressor comprises:monitoring relative humidity within the inlet section ...

INTEGRAL CERAMIC MATRIX COMPOSITE FASTENER WITH NON-POLYMER RIGIDIZATION

A method of forming an integral fastener for a ceramic matrix composite component comprises the steps of forming a fiber preform with an opening, forming a fiber fastener, inserting the fiber fastener into the opening, and infiltrating a matrix material into the fiber preform and fiber fastener to form a ceramic matrix composite component with an integral fastener. A gas turbine engine is also disclosed. 1. A gas turbine engine component comprising: a fiber preform with an opening that has a wide portion at one surface of the fiber preform and a narrow portion at an opposite surface of the fiber preform,', 'a fiber fastener having a fastener body extending from a first end to a second end, the fastener body defined by a first dimension and having an enlarged head at the first end of the fastener body that is defined by a second dimension that is greater than the first dimension,', 'wherein the fiber fastener is inserted into the opening to form a dry fiber preform and fiber fastener assembly, the fastener body extending through the dry fiber preform such that the enlarged head portion is received within the wide portion of the opening, and', 'wherein a matrix material is infiltrated into the dry fiber preform and fiber fastener assembly to provide the single-piece structure; and, 'a gas turbine engine component body formed of a ceramic matrix composite material having at least one fastener integrally formed with the gas turbine engine component body as a single-piece structure, wherein the single-piece structure initially comprises'}an engine support structure, wherein the at least one fastener of the single-piece structure connects the gas turbine engine component body to the engine support structure.2. The gas turbine engine component according to wherein the fiber preform comprises a rigidized preform structure having the opening to receive the fiber fastener claim 1 , and wherein the fiber fastener comprises a woven fastener formed from a fiber based material ...

COMPRESSOR INJECTOR APPARATUS AND SYSTEM

Cooling systems for high pressure compressor systems are provided. The cooling systems may comprise tangential on board injectors (“TOBIs”). The TOBIs may comprise one or more fluid channels configured to conduct cooling fluid flow to components of the compressor, including, for example, disk-hub portions of the compressor. In this regard, the TOBI may be configured to exhaust cooling air in a manner such that the exhausted air has a similar linear velocity of the disk-hub portion. The cooling air may also be exhausted in a manner that is substantially parallel to the disk-hub portion. 1. A compressor , comprising:a disk-hub;a stator portion;a rotor portion coupled to the disk hub, the rotor portion adjacent to and aft the stator portion;an exit guide vane adjacent to and aft the rotor portion;a dual tangential on-board injector (“dual TOBI”) disposed radially inward of the exit guide vane and the rotor portion, wherein the dual TOBI is configured to conduct a cooling flow to a section of the disk hub adjacent the stator portion.2. The compressor of claim 1 , wherein the rotor portion is coupled to the disk-hub by a blade attachment.3. The compressor of claim 2 , wherein the blade attachment defines a fluid conduit.4. The compressor of claim 3 , wherein the dual TOBI includes a first channel defined within a portion of a root of the exit guide vane and a second channel defined within a portion of a root of the rotor portion.5. The compressor of claim 4 , wherein the first channel and the second channel are in fluid communication.6. The compressor of claim 4 , wherein the first channel is operatively coupled to the second channel via the fluid conduit.7. The compressor of claim 4 , wherein a first pressure (P) in the first channel is equivalent to a second pressure (P) adjacent and aft the exit guide vane.8. The compressor of claim 7 , wherein a third pressure (P) in the second channel is equivalent to a fourth pressure (P) that is in a section between the stator ...

ROTOR HUB SEAL

A gas turbine engine includes a rotor hub that extends along a longitudinal axis of the gas turbine engine and includes a radially outer surface that has a substantially axially extending projection and a substantially radially extending projection. A seal ring includes a body portion and at least one knife edge that extends outward from the body portion. A recess is in the body portion for accepting the axially extending portion of the rotor hub. A lock ring is in abutting contact with the seal ring and the radially extending projection on the rotor hub for securing the seal ring to the rotor hub. 1. A gas turbine engine comprising:a rotor hub extending along a longitudinal axis of the gas turbine engine including a radially outer surface having a substantially axially extending projection and a substantially radially extending projection;a seal ring including a body portion and at least one knife edge extending outward from the body portion, a recess in the body portion for accepting the axially extending portion of the rotor hub; anda lock ring in abutting contact with the seal ring and the radially extending projection on the rotor hub for securing the seal ring to the rotor hub.2. The gas turbine engine of claim 1 , including a plenum at least partially defined by the rotor hub and the seal ring.3. The gas turbine engine of claim 2 , including a feed passage in the rotor hub in fluid communication with the plenum.4. The gas turbine engine of claim 3 , wherein the substantially axially extending projection is located axially upstream of the substantially radially extending projection.5. The gas turbine engine of claim 4 , wherein a radially inward facing surface on the substantially axially extending projection is located radially outward from a distal end of the substantially radially extending projection.6. The gas turbine engine of claim 5 , wherein the recess in the body portion is at least partially defined by an axially extending projection of the seal ...

GAS TURBINE ENGINE AIRFOIL WITH WISHBONE BAFFLE COOLING SCHEME

A gas turbine engine component includes a structure including spaced apart first and second exterior walls that extend in a first direction to an endwall. The first and second exterior walls are joined at the endwall to provide a cooling cavity. A wishbone baffle is arranged in the cooling cavity and includes first and second interior walls respectively adjacent to the first and second exterior walls. The first and second interior walls extend in the first direction to and are joined by an apex to provide a first cavity. The wishbone baffle separates the first cavity from a second cavity provided between the apex and the endwall. 1. A gas turbine engine component comprising:a structure including spaced apart first and second exterior walls extending in a radial direction to an endwall, the first and second exterior walls are joined at the endwall to provide a cooling cavity, the first and second exterior walls are respectively pressure and suction side walls joined at leading and trailing edges spaced from one another in a chord-wise direction;wherein the first and second exterior walls are not parallel to one another and provide first and second portions of the cooling cavity respectively at first and second radial positions that are different than one another, the first portion has a first portion width in a circumferential thickness direction at a chord-wise location along a chord-wise direction, the second portion has a second portion width in the circumferential thickness direction at the chord-wise location, the second portion width is larger than the first portion width; anda wishbone baffle arranged in the second portion of the cooling cavity and including first and second interior walls respectively adjacent to the first and second exterior walls, the first and second interior walls extending in a radial direction and joined by an apex to provide a first cavity, the wishbone baffle separates the first cavity from a second cavity provided between the apex ...

TURBINE ENGINE COMPONENT INCLUDING AN AXIALLY ALIGNED SKIN CORE PASSAGE INTERRUPTED BY A PEDESTAL

A component for a turbine engine includes a fore edge connected to an aft edge via a first surface and a second surface. Multiple cooling passages are defined within the turbine engine component. A skin core passage is defined immediately adjacent the first surface, and at least one pedestal interrupts a flow path through the skin core passage. 1. A turbine engine component comprising:a fore edge connected to an aft edge via a first surface and a second surface;a plurality of cooling passages defined within the turbine engine component;a skin core passage defined immediately adjacent said first surface; andat least one pedestal interrupting a flow path through said skin core passage.2. The turbine engine component of claim 1 , wherein said at least one pedestal comprises a plurality of pedestals distributed through said skin core passage.3. The turbine engine component of claim 2 , wherein at least one pedestal in said plurality of pedestals includes a film cooling hole defined at least partially within the pedestal claim 2 , the film cooling hole connecting one of said plurality of cooling passages to said first surface.4. The turbine engine component of claim 3 , wherein said film cooling hole connects a first cooling air flow path internal to said turbine engine component to said first surface claim 3 , and wherein said skin core passage receives cooling air from a second cooling air flow path internal to said turbine engine component claim 3 , and wherein said second cooling air flow path is a low pressure cooling air flow path relative to said first cooling air flow path.5. The turbine engine component of claim 3 , wherein one of said plurality of cooling passages is connected to said first surface via said cooling hole has a higher cooling air pressure than a cooling air pressure of said skin core passage.6. The turbine engine component of claim 1 , wherein said at least one pedestal has an oblong cross section relative to a turbine engine including the ...

TURBINE AIRFOIL WITH TRAILING EDGE COOLING CIRCUIT

One aspect of the disclosure provides for a turbine airfoil. The turbine airfoil may include a trailing edge having: a set of cooling channels having a first cooling channel fluidly connected to a second cooling channel; a first section having a first pin bank cooling arrangement, the first section fluidly connected to the first cooling channel; a second section having a second pin bank cooling arrangement, the second section fluidly connected to the second cooling channel and being radially inward of the first section; and a pressure side panel having a third pin bank cooling arrangement, the pressure side panel fluidly connected to the first cooling channel. 1. A trailing edge of a turbine airfoil , comprising:a first section having a first pin bank cooling arrangement, the first section fluidly connected to a first cooling channel; anda second section having a second pin bank cooling arrangement, the second section fluidly connected to a second cooling channel and being radially inward of the first section.2. The trailing edge of the turbine airfoil of claim 1 , further comprising a first set of crossover holes fluidly connecting the second section to the second cooling channel.3. The trailing edge of the turbine airfoil of claim 2 , wherein the first set of crossover holes are angled toward a pressure sidewall of the turbine airfoil.4. The trailing edge of the turbine airfoil of claim 2 , further comprising:a first set of raised features, each raised feature in the first set of raised features corresponding to a crossover hole in the first set of crossover holes and each raised feature in the first set of raised features extending toward a suction sidewall of the turbine airfoil.5. The trailing edge of the turbine airfoil of claim 1 , further comprising a third section radially inward of the second section claim 1 , the third section including a third pin bank cooling arrangement and being fluidly connected to the second cooling channel.6. The trailing edge of ...

Ceramic Matrix Composite Component Cooling

Nozzle segments and methods of cooling airfoils of nozzle segments are provided. For example, a turbine nozzle segment includes an inner band defining an inner band cavity and/or an outer band defining an outer band cavity. The inner band may define an inner band aperture extending from the inner band cavity through the inner band, and the outer band may define an outer band aperture extending from the outer band cavity through the outer band. Inner and/or outer band cooling passages may extend through a trailing edge portion of a CMC airfoil of the nozzle segment. An inlet of any inner band cooling passage is defined adjacent an inner band aperture, and an inlet of any outer band cooling passage is defined adjacent an outer band aperture. The cooling passage inlets are aligned with the adjacent inner or outer band apertures to provide cooling fluid from the respective cavity. 1. A turbine nozzle segment for a gas turbine engine , the turbine nozzle segment comprising:an inner band defining an inner band cavity for receipt of a cooling fluid and an inner band aperture extending from the inner band cavity through the inner band;an outer band; opposite pressure and suction sides extending radially along a span, the pressure and suction sides defining an outer surface of the CMC airfoil,', 'opposite leading and trailing edges extending radially along the span, the pressure and suction sides extending axially between the leading and trailing edges, the leading edge defining a forward end of the CMC airfoil, the trailing edge defining an aft end of the CMC airfoil, and', 'a trailing edge portion defined adjacent the trailing edge at the aft end; and, 'a ceramic matrix composite (CMC) airfoil extending from the inner band to the outer band, the CMC airfoil including'}an inner band cooling passage extending through the CMC airfoil from an inlet defined adjacent the inner band aperture to an outlet defined in the outer surface,wherein the inner band aperture and the inlet ...

SLIDING SEAL

The present disclosure relates generally to a sliding seal between two components. The sliding seal includes a first seal section, a second seal section, and (in some embodiments) a third seal section. Two or three of the seal sections are uncoupled, which allows the uncoupled seal sections to move relative to one another during relative movement between the two components. One or more spring tabs extend from the second seal section and bias the first and third (or in some embodiments, the first and second) seal sections away from one another. 1. A seal for sealing a seal cavity defined by first and second adjacent components disposed about an axial centerline , the seal comprising:a first seal section comprising a first split hoop; anda second seal section comprising a second split hoop and a spring element; a third seal section; wherein the spring element contacts the first seal section and is configured to axially load the first seal section against one of the first and second components; wherein the second split hoop contacts the third seal section and is configured to axially load the third seal section against another one of the first and second components; and wherein at least two of the first, second and third seal sections are configured to move relative to one another.2. The seal of claim 1 , wherein the first seal section is configured to sealingly engage with the first and second components.3. The seal of claim 1 , wherein the third seal section comprises a split hoop.4. The seal of claim 1 , wherein the first seal section comprises a first seal section free-state inner diameter that is smaller than a seal cavity inner diameter claim 1 , such that a radial preload is achieved between the first seal section and at least one of the first and second components.5. The seal of claim 1 , wherein the spring element comprises a plurality of first spring tabs frustoconically but primarily radially extending from said second split hoop.6. The seal of claim 5 , ...

BEVELED COVERPLATE

A gas turbine engine includes a platform that has a gas path side, a non-gas path side, a first mate face, and a second mate face. The second mate face has a beveled edge sloping towards the first mate face. The gas turbine engine also includes a coverplate that includes a first bend, a flat portion substantially parallel to the first mate face and a first wing substantially parallel to the second mate face. 1. A gas turbine engine assembly , comprising:a compressor section;a combustor section; and [ a gas path side,', 'a non-gas path side,', 'a platform floor on the non-gas path side,', 'a first mate face on the non-gas path side offset from the platform floor by, 'a platform having, 'a second mate face on the non-gas path side, the second mate face having a beveled edge sloping towards the platform floor and offset from the platform floor by a second distance that is larger than the first distance, and', 'a first distance,'}, 'a coverplate including a first bend, a flat portion substantially parallel to the first mate face of the platform, and a first wing that is substantially parallel to the second mate face of the platform., 'a turbine section, wherein at least one of the compressor section or the turbine section includes a component that includes2. The assembly of claim 1 , further comprising a duct adapted to channel coolant from the compressor section to the non-gas path side of the platform.3. The assembly of claim 1 , further comprising a cavity defined by the coverplate and the platform claim 1 , wherein the coverplate defines a cooling hole adapted to allow coolant to flow into the cavity.4. A gas turbine engine assembly claim 1 , comprising:a compressor section;a combustor section; and [ a gas path side,', 'a non-gas path side,', 'a platform floor on the non-gas path side,', 'a first mate face on the non-gas path side offset from the platform floor by, 'a platform having, 'a second mate face on the non-gas path side, the second mate face having a ...

GAS TURBINE ENGINE COMPONENT COOLING PASSAGE AND SPACE EATING CORE

A gas turbine engine includes a structure that has walls that provide a cooling passage and a cooling surface. A non-ferrous obstruction is relative to the walls. The obstruction includes a portion spaced from the cooling surface to provide a gap which is configured to receive a cooling fluid. 1. A gas turbine engine component comprising:a structure having walls providing a cooling passage having a cooling surface;a non-ferrous obstruction arranged in the cooling passage engaging the cooling surface to locate the obstruction relative to the walls, and the obstruction including a portion spaced from the cooling surface to provide a gap configured to receive a cooling fluid.2. The gas turbine engine component according to claim 1 , wherein the structure is an airfoil.3. The gas turbine engine component of claim 2 , wherein the structure a turbine blade or a turbine vane.4. The gas turbine engine component according to claim 1 , wherein the structure is a blade outer air seal.5. The gas turbine engine component according to claim 1 , wherein the structure is a platform that supports an airfoil.6. The gas turbine engine component according to claim 1 , wherein the structure is a combustor liner.7. The gas turbine engine component according to claim 1 , wherein the structure is an exhaust liner.8. The gas turbine engine component according to claim 1 , wherein the obstruction is arranged to block a flow of the cooling fluid through the cooling passage.9. The gas turbine engine component according to claim 1 , wherein the obstruction includes a refractory metal.10. The gas turbine engine component according to claim 1 , wherein the obstruction provides multiple gaps adjacent to the cooling surface.11. The gas turbine engine component according to claim 1 , wherein the structure includes a nickel alloy claim 1 , and the obstruction is provided by a material that is different than the nickel alloy.12. The gas turbine engine component according to claim 1 , wherein the ...

HYDROPHOBIC MATERIALS INCORPORATING RARE EARTH ELEMENTS AND METHODS OF MANUFACTURE

This invention relates generally to an article that includes a base substrate, an intermediate layer including at least one element or compound selected from titanium, chromium, indium, zirconium, tungsten, and titanium nitride on the base substrate, and a hydrophobic coating on the base substrate, wherein the hydrophobic coating includes a rare earth element material (e.g., a rare earth oxide, a rare earth carbide, a rare earth nitride, a rare earth fluoride, and/or a rare earth boride). An exposed surface of the hydrophobic coating has a dynamic contact angle with water of at least about 90 degrees. A method of manufacturing the article includes providing the base substrate and forming an intermediate layer coating on the base substrate (e.g., through sintering or sputtering) and then forming a hydrophobic coating on the intermediate layer (e.g., through sintering or sputtering). 1. An article comprising:a base substrate;an intermediate layer on the base substrate, wherein the intermediate layer comprises at least one element or compound selected from the list comprising titanium, chromium, indium, zirconium, tungsten, and titanium nitride; anda hydrophobic layer on the intermediate layer, the hydrophobic layer comprising a rare earth element material.2. The article of claim 1 , wherein the rare earth element material is or comprises rare earth oxide or a lanthanide series rare earth oxide.3. The article of claim 2 , wherein the rare earth oxide is or comprises cerium (IV) oxide (“ceria”).4. The article of any claim 1 , wherein the intermediate layer is or comprises titanium.5. The article of claim 4 , wherein the titanium is pure titanium.6. The article of claim 4 , wherein the titanium is a titanium alloy.7. The article of claim 4 , wherein the titanium is doped with a second material.8. The article of claim 1 , wherein an exposed surface of the article has a dynamic contact angle with water of at least about 90 degrees.9. The article of claim 1 , wherein the ...

METHOD FOR TREATING A GAS TURBINE BLADE AND GAS TURBINE HAVING SAID BLADE

The use of different ceramic layers allows different configurations of gas turbines to be produced each of which is optimized for a respective use of base load operation or peak load operation. 1. A method for operating a gas turbine system , comprising:altering turbine blades or vanes of a first gas turbine to produce a second gas turbine, the first turbine blades or vanes having first ceramic thermal barrier coatings configured for base load operation;removing completely the first ceramic thermal barrier coating of a first turbine blade or vane of the first gas turbine;applying a new, second ceramic thermal barrier coating configured for peak load operation, different from the one application, to the first turbine blade or vane from which the first ceramic thermal barrier coating has been completely removed to produce a second turbine blade or vane by coating with a ceramic and a polymer or with a ceramic with grains of at least 20% greater mean particle diameter to generate pores in the new, second ceramic thermal barrier coating; andstarting the second gas turbine daily instead of continuously operating the second gas turbine;wherein the new, second ceramic thermal barrier coating differs significantly from the completely removed first ceramic thermal barrier coating, in that the porosities of the first and the second ceramic thermal barrier coatings are different, and the absolute difference in the reduced or increased porosity is at least 2%, and difference in coating thicknesses of the first and the second ceramic thermal barrier coatings is at least 50 μm;wherein the porosity of the completely removed first ceramic thermal barrier coating, and the porosity of the second ceramic thermal barrier coating are in the range 12%±4% to 25%±4%,wherein the second turbine blade or vane is incorporated in the second gas turbine.2. The method as claimed in claim 1 , wherein the completely removed first ceramic thermal barrier coating is a two-layer ceramic thermal ...

SYSTEMS AND METHODS FOR OPTIMAL SOURCE MATERIAL DEPOSITION ALONG HOLE EDGES

A method for depositing a coating of a source material onto a panel is disclosed. The method includes providing a cathodic arc, the cathodic arc including a target surface, the target surface disposed along a target deposition axis and able to emit the source material as a generally cloud of source material vapor and a generally conical stream of liquid particles of the source material. The method further includes positioning the panel relative to the target surface based on a deposition angle, the deposition angle being between the target surface and an outer limit of the generally conical stream of liquid particles o the source material. The method may further include emitting the source material from the target surface as the generally conical cloud of source material vapor and coating the edge with the cloud of source material vapor to provide an edge coating. 1. A system for depositing a coating of a source material onto at least one panel , the at least one panel defining an edge and a front panel surface , the system comprising:a cathodic arc including a target surface, the target surface disposed along a target deposition axis, and able to emit the source material as both a cloud of source material vapor and a generally conical stream of liquid particles of the source material, the cloud of source material vapor used to coat the edge with the source material to provide an edge coating; anda coating deposition structure, the coating deposition structure positioning one or both of the cathodic arc and the panel, such that the panel is positioned relative to the target surface based on a deposition plane, the deposition plane being defined by the target deposition axis, wherein the entire panel is positioned above or below the deposition plane and within the generally conical stream of liquid particles.2. The system of claim 1 , wherein the coating deposition structure positions the panel substantially perpendicular to the target deposition axis.3. The system ...

DUAL COOLING AIRFLOW TO BLADES

An airfoil may comprise a root and an airfoil body radially outward of the root. The airfoil body may define a first cooling chamber and a second cooling chamber. A first passage may be defined within the root and configured to direct a first airflow radially outward through the root into the first cooling chamber. A second passage may be defined within the root and configured to direct a second airflow radially outward through the root and into the second cooling chamber. A tangential onboard injector (TOBI) may be disposed in the first airflow path. A radial onboard injector (ROBI) may be disposed in the second airflow path. 1. An airfoil , comprising:a root;an airfoil body radially outward of the root, the airfoil body defining a first cooling chamber and a second cooling chamber;a first passage defined within the root and configured to direct a first airflow received from a tangential onboard injector (TOBI) radially outward through the root into the first cooling chamber; anda second passage defined within the root and configured to direct a second airflow received from a radial onboard injector (ROBI) radially outward through the root and into the second cooling chamber,wherein the TOBI is disposed radially outward from the ROBI, andthe first cooling chamber is disposed forward from the second cooling chamber.2. The airfoil of claim 1 , further comprising a first inlet defined in an axially forward surface of the root.3. The airfoil of claim 2 , further comprising a second inlet defined in an axially aft surface of the root.4. The airfoil of claim 3 , further comprising a leading edge and a trailing edge claim 3 , wherein the first cooling chamber is disposed at the leading edge and the second cooling chamber is disposed at the trailing edge.5. The airfoil of claim 4 , wherein the airfoil body defines a first plurality of holes at the leading edge claim 4 , and wherein the first airflow is directed out the airfoil through the first plurality of holes.6. The ...

INTERNAL COOLING CAVITY WITH TRIP STRIPS

An airfoil is provided. The airfoil may comprise a cross over, an impingement chamber in fluid communication with the cross over, and a first trip strip disposed on a first surface of the impingement chamber. A cooling system is also provided. The cooling system may comprise an impingement chamber, a first trip strip on a first surface of the impingement chamber, and a second trip strip on a second surface of the impingement chamber. An internally cooled engine part is further provided. The internally cooled part may comprise a cross over and an impingement chamber in fluid communication with the cross over. The cross over may be configured to direct air towards a first surface of the impingement chamber. A first trip strip may be disposed on the first surface of the impingement chamber. 1. An airfoil , comprising:an internal cavity;an impingement chamber proximate the internal cavity and in fluid communication with the cavity via a cross over, wherein the cross over is configured to direct a coolant as a jet towards a suction-side surface of the impingement chamber;a first trip strip disposed on the suction-side surface of the impingement chamber, wherein the first trip strip has a first geometry selected from a group consisting of at least one of: a v-shaped geometry having a point of the v-shaped geometry oriented relatively downward along an axial extent of the impingement chamber away from a tip of the airfoil, a circular geometry, a wave geometry, an annular elliptical geometry, or a linear geometry non-parallel and non-orthogonal to the axial extent of the impingement chamber; anda second trip strip disposed on a pressure-side surface of the impingement chamber, wherein the second trip strip has a second geometry selected from a group consisting of at least one of: the v-shaped geometry having the point of the v-shaped geometry oriented relatively downward along the axial extent of the impingement chamber away from the tip, the circular geometry, the wave ...

VENTED TANGENTIAL ON-BOARD INJECTOR FOR A GAS TURBINE ENGINE

An on-board injector that delivers discharge air toward a turbine rotor of a gas turbine engine includes a second wall spaced form a first wall to define an annular inlet about an engine longitudinal axis and a multiple of airfoil shapes between the first wall and the second wall to segregate discharge air from the annular inlet, and a multiple of bypass apertures each along a radial axis transverse to the engine longitudinal axis through each of the multiple of airfoil shapes and the respective first wall, the second wall. 118-. (canceled)19. A method of managing purge air within a turbo machine comprising the steps of:segregating discharge air from an annular inlet with a multiple of airfoil shapes, the annular inlet defined around an engine longitudinal axis; anddirecting purge air through a multiple of bypass apertures each along a radial axis transverse to the engine longitudinal axis and through each of the multiple of airfoil shapes.20. The method according to claim 19 , further comprising:tangentially directing the discharge air through a tangential on board injector.21. The method according to claim 19 , further comprising:directing the discharge air at an angle through an angled on board injector.22. A system for a gas turbine engine comprising:a coverplate for a turbine rotor defined about an engine longitudinal axis, said coverplate including a multiple of coverplate apertures; andan on-board injector with a multiple of airfoil shapes between a first wall and a second wall to define an annular inlet about the engine longitudinal axis, said multiple of airfoil shapes operable to segregate and direct discharge air from the annular inlet toward said multiple of coverplate apertures, said on-board injector including a multiple of bypass apertures each along a radial axis transverse to the engine longitudinal axis, each of said multiple of apertures extends through one of said multiple of airfoil shapes, said first wall, and said second wall, wherein said on- ...

TIP LEAKAGE FLOW DIRECTIONALITY CONTROL

An airfoil according to an example includes, among other things, a suction sidewall and a pressure sidewall. A tip wall extends from a leading edge to a trailing edge and joins respective outer ends of the suction and pressure sidewalls. A tip rib extends along a pressure side of the tip wall. A tip leakage control channel is provided at the tip wall and has a floor that extends between a first control channel vane sidewall and a second control channel vane sidewall established by a corresponding tip leakage control vane. Each of the tip leakage control vanes is contiguous with and extending from a suction side surface of the tip rib. The tip leakage channel extends toward the trailing edge while extending toward the suction sidewall. One or more of an internal cavity, a channel cooling aperture, and a sidewall microcircuit may be provided. 1. An airfoil , comprising:a suction sidewall and a pressure sidewall, each sidewall extending spanwise from an airfoil base and extending chordwise between a leading edge and a trailing edge;a tip wall extending chordwise from said leading edge to said trailing edge and joining respective outer spanwise ends of said suction and pressure sidewalls, including a tip rib extending chordwise along a pressure side of said tip wall;a plurality of tip leakage control channels provided at said tip wall, each of said plurality of tip leakage control channels including a control channel floor that extends between a first control channel vane sidewall and a second control channel vane sidewall established by a corresponding tip leakage control vane, each of said tip leakage control vanes contiguous with a suction side surface of said tip rib, said plurality of tip leakage channels extending toward said trailing edge while extending toward said suction sidewall;an internal cavity disposed between said suction sidewall and said pressure sidewall; andan aperture extending from said internal cavity to at least one of said tip leakage control ...

TEMPERATURE-MODULATED RECUPERATED GAS TURBINE ENGINE

A recuperated gas turbine engine includes an engine core that has a compressor section, a combustor section, and a turbine section. An exhaust duct is located downstream of the turbine section for receiving a hot turbine exhaust stream from the turbine section. The exhaust duct includes a heat exchanger and a temperature-control module upstream of the heat exchanger. A first compressor bleed line portion leads into the heat exchanger, and a second compressor bleed lie portion leads into the exhaust duct upstream of the heat exchanger. A compressor return line leads from the heat exchanger into the engine core upstream of the combustor section. The compressor bleed line is operable to selectively feed compressed air to the heat exchanger, and the temperature-control module is operable to selectively modulate at least one of temperature and flow of the hot turbine exhaust stream with respect to the heat exchanger. 1. A recuperated gas turbine engine comprising:an engine core including a compressor section, a combustor section, and a turbine section;an exhaust duct downstream of the turbine section for receiving a hot turbine exhaust stream from the turbine section, the exhaust duct including a heat exchanger and a temperature-control module upstream of the heat exchanger;a first compressor bleed line portion leading into the heat exchanger and operable to selectively feed compressed air from the compressor section to the heat exchanger;a second compressor bleed line portion leading into the exhaust duct upstream of the heat exchanger and operable to selectively feed compressed air from the compressor section to the exhaust duct upstream of the heat exchanger; anda compressor return line leading from the heat exchanger into the engine core upstream of the combustor section, the compressor bleed line operable to selectively feed compressed air to the heat exchanger;the temperature-control module operable to selectively modulate at least one of temperature and flow of ...

GAS TURBINE ENGINE FLUID COOLING SYSTEMS AND METHODS OF ASSEMBLING THE SAME

A fluid cooling system for use in a gas turbine engine including a fan casing circumscribing a core gas turbine engine includes a heat source configured to transfer heat to a heat transfer fluid and a primary heat exchanger coupled in flow communication with the heat source. The primary heat exchanger is configured to channel the heat transfer fluid therethrough and is coupled to the fan casing. The fluid cooling system also includes a secondary heat exchanger coupled in flow communication with the primary heat exchanger. The secondary heat exchanger is configured to channel the heat transfer fluid therethrough and is coupled to the core gas turbine engine. The fluid cooling system also includes a bypass mechanism coupled in flow communication with the secondary heat exchanger. The bypass mechanism is selectively moveable based on a temperature of a fluid medium to control a cooling airflow through the secondary heat exchanger. 1. A fluid cooling system for use in a gas turbine engine including a core section of the gas turbine engine having an axis of rotation and a fan casing substantially circumscribing the core section of the gas turbine engine , said fluid cooling system comprising:a heat source configured to transfer heat to a heat transfer fluid;a primary heat exchanger coupled in flow communication with said heat source and configured to channel the heat transfer fluid therethrough, said primary heat exchanger coupled to the fan casing;a secondary heat exchanger coupled in flow communication with said primary heat exchanger and configured to channel the heat transfer fluid therethrough, said secondary heat exchanger located within to the core section of the gas turbine engine;a bypass mechanism coupled in flow communication with said secondary heat exchanger, said bypass mechanism being selectively moveable based on a temperature of a fluid medium to control a cooling airflow through said secondary heat exchanger.2. The fluid cooling system in accordance ...

CENTRIFUGAL AIRFOIL COOLING MODULATION

In one example embodiment, a blade includes an attachment region, an airfoil extending from the attachment region, and a blade cooling arrangement. The blade cooling arrangement includes at least a first feed passage disposed through the attachment region, which is connected to a first cooling passage disposed in the airfoil. A passively actuated first coolant valve is disposed in or proximate the first feed passage. A plurality of such blades can be disposed in a turbine section of an engine. 1. A blade comprising:an attachment region;an airfoil extending from the attachment region;a blade cooling arrangement including at least a first feed passage disposed through the attachment region, and connected to a first cooling passage disposed in the airfoil; anda passively actuated first coolant valve disposed in or proximate the first feed passage;wherein the first coolant valve includes a first blocking element attached to a first blade surface; and a first hinge portion; and', 'a first blocking portion extending from the first hinge portion and across a portion of the blade cooling arrangement., 'wherein the first blocking element comprises2. The blade of claim 1 , wherein the first cooling passage is selected from: a serpentine passage claim 1 , a leading edge impingement passage claim 1 , a trailing edge cooling passage claim 1 , and a sidewall microcircuit.3. The blade of claim 1 , wherein the first coolant valve is actuated in response to a centrifugal force applied in a spanwise direction of the blade.4. The blade of claim 1 , wherein the first coolant valve is disposed over an opening to the first root passage.5. The blade of claim 1 , wherein the first coolant valve is disposed within the first root passage or proximate an interface between the first root passage and the first cooling passage.6. The blade of claim 1 , wherein the airfoil also includes a plurality of film cooling holes formed through a sidewall of the airfoil between the first cooling passage ...

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. 1. An airfoil for a gas turbine engine comprising:an outer airfoil wall providing 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 for a trailing edge passage, the supply passage being arranged upstream from and in fluid communication with the trailing edge passage, a cooling hole extending through the outer airfoil wall from the supply passage to the exterior surface on the suction side; andwherein the cooling passages include first, second and third cooling passages are arranged progressively aftward of one another wherein the second cooling passage is arranged aftward of the first cooling passage, and the third cooling passage is arranged aftward of the second cooling passage, the supply passage is provided by the third cooling passage, and wherein the second passage includes a baffle, the cooling hole is arranged downstream from the baffle with respect to a chordwise direction provided between the leading and trailing edges, an end of the baffle arranged in the supply passage.2. The airfoil according to claim 1 , wherein the baffle is spaced apart from an inner surface of the second cooling passage by a rib.3. The airfoil according to claim 1 , wherein the trailing edge passage includes first and second surfaces that are interconnected to one another in a thickness direction by a structure claim 1 , the thickness ...

GEARED ARCHITECTURE TURBOFAN ENGINE THERMAL MANAGEMENT SYSTEM AND METHOD

A method of sizing a heat exchanger for a geared architecture gas turbine engine includes sizing a minimum frontal area of at least one heat exchanger located in communication with a fan bypass airflow such that a ratio of waste heat area to horsepower generation characteristic area is between 1.6 to 8.75. 1. A method of sizing a heat exchanger for a geared architecture gas turbine engine comprising:sizing a minimum frontal area of at least one heat exchanger located in communication with a fan bypass airflow of the geared architecture gas turbine engine, such that a ratio of waste heat area to horsepower generation characteristic area is between 1.6 to 17.5, the waste heat area defined by the frontal area of the at least one heat exchanger, wherein the horsepower generation characteristic area is defined by an exit area of a high pressure compressor of the geared architecture gas turbine engine.2. The method as recited in claim 1 , wherein the geared architecture provides an efficiency of at least 97.7%3. The method as recited in claim 1 , further comprising locating the at least one heat exchanger within a fan bypass airflow path.4. The method as recited in claim 1 , further comprising locating the at least one heat exchanger within a fan bypass airflow path such that the ratio of waste heat area to horsepower generation characteristic area is between 1.6 to 8.75.5. The method as recited in claim 1 , further comprising locating the at least one heat exchanger with respect to a fan duct total pressure profile.6. A geared architecture gas turbine engine comprising:a geared architecture;a fan driven by the geared architecture of the gas turbine engine to generate a fan bypass airflow; and{'sup': 2', '2', '2, 'at least one heat exchanger mounted in communication with the fan bypass airflow, a minimum frontal area of the at least one heat exchanger defines an area less than 420 in(270967 mm) and greater than 0 in.'}7. The geared architecture gas turbine engine as ...

ENGINE ASSEMBLY WITH COMBINED ENGINE AND COOLING EXHAUST

An engine assembly for an aircraft, including an internal combustion engine having a liquid coolant system in fluid communication with a heat exchanger, an exhaust duct in fluid communication with air passages of the heat exchanger, a fan in fluid communication with the exhaust duct for driving a cooling air flow through the air passages of the heat exchanger and into the exhaust duct, and an intermediate duct in fluid communication with an exhaust of the engine and having an outlet positioned within the exhaust duct downstream of the fan and upstream of the outlet of the exhaust duct. The outlet of the intermediate duct is spaced inwardly from a peripheral wall of the exhaust duct. The engine assembly may be configured as an auxiliary power unit. A method of discharging air and exhaust gases in an auxiliary power unit having an internal combustion engine is also discussed. 1. An engine assembly for an aircraft , the engine assembly comprising:an internal combustion engine having a liquid coolant system;a heat exchanger having coolant passages in fluid communication with the liquid coolant system and air passages in heat exchange relationship with the coolant passages;an exhaust duct in fluid communication with the air passages of the heat exchanger, the exhaust duct having an outlet in fluid communication with an environment of the aircraft;a fan in fluid communication with the exhaust duct for driving a cooling air flow through the air passages of the heat exchanger and into the exhaust duct; andan intermediate duct in fluid communication with an exhaust of the internal combustion engine, the intermediate duct having an outlet positioned within the exhaust duct downstream of the fan and upstream of the outlet of the exhaust duct, the outlet of the intermediate duct spaced inwardly from a peripheral wall of the exhaust duct.2. The engine assembly as defined in claim 1 , wherein the intermediate duct is in fluid communication with an exhaust of the internal ...

GAS TURBINE ENGINE MOTORING SYSTEM FOR BOWED ROTOR ENGINE STARTS

A system for a gas turbine engine is provided. The system comprising: a gas turbine engine including rotational components comprising an engine compressor, an engine turbine, and a rotor shaft operably connecting the engine turbine to the engine compressor, wherein each rotational component is configured to rotate when any one of the rotational components is rotated; a permanent magnet alternator operably connected to at least one of the rotational components, the permanent magnet alternator being configured to rotate the rotational components; and a motor controller in electronic communication with the permanent magnet alternator, the motor controller being configured to command the permanent magnet alternator to rotate the rotational components at a selected angular velocity for a selected period of time. 1. A system for cooling a gas turbine engine , the system comprising:a gas turbine engine including rotational components comprising an engine compressor, an engine turbine, and a rotor shaft operably connecting the engine turbine to the engine compressor, wherein each rotational component is configured to rotate when any one of the rotational components is rotated;a permanent magnet alternator operably connected to at least one of the rotational components, the permanent magnet alternator being configured to rotate the rotational components; anda motor controller in electronic communication with the permanent magnet alternator, the motor controller being configured to command the permanent magnet alternator to rotate the rotational components at a selected angular velocity for a selected period of time.2. The system of claim 1 , further comprising:an accessory gearbox operably connecting the permanent magnet alternator to at least one of the rotational components.3. The system of claim 1 , wherein:the permanent magnet alternator is configured to generate electricity when the rotational components are rotating under power of the gas turbine engine.4. The system ...

GAS TURBINE ENGINE COMPONENT COOLING WITH RESUPPLY OF COOLING PASSAGE

A gas turbine engine component with a core includes a first core portion configured to provide a first cooling passage. A second core portion is spaced from the first core portion and configured to provide a second cooling passage. The second core portion includes a longitudinal leg and a resupply leg transverse to and intersecting the longitudinal leg. The resupply leg has a terminal end and is configured to provide a resupply channel A connector interconnects the terminal end to the first core portion. The connector is configured to provide a resupply hole. 1. A gas turbine engine component with a core comprising:a first core portion configured to provide a first cooling passage;a second core portion spaced from the first core portion and configured to provide a second cooling passage, the second core portion including a longitudinal leg and a resupply leg transverse to and intersecting the longitudinal leg, the resupply leg having a terminal end and configured to provide a resupply channel; anda connector interconnects the terminal end to the first core portion, the connector configured to provide a resupply hole.2. The gas turbine engine component with the core according to claim 1 , wherein the second core portion includes multiple resupply legs spaced apart from one another and connected to the same longitudinal leg.3. The gas turbine engine component with the core according to claim 2 , wherein first and second resupply legs are arranged on opposite lateral sides of the longitudinal leg.4. The gas turbine engine component with the core according to claim 3 , wherein the structure includes multiple longitudinal legs each including multiple resupply legs claim 3 , at least some of the multiple resupply legs of adjacent multiple longitudinal legs interleaved with one another.5. The gas turbine engine component with the core according to claim 1 , wherein the resupply leg converges along a length extending from the connector to the longitudinal leg.6. The gas ...

AIRFOIL WITH PANEL AND SIDE EDGE COOLING

An airfoil includes a core structure that defines a cooling passage and first and second panels that are attached with the core structure. Each of the first and second panels includes an exterior gas path side, an opposed interior side, and a side edge. The side edge has a bearing portion that defines a bearing surface. The bearing surface of the first panel abuts the bearing surface of the second panel at a bearing interface. There are channels in the bearing portion of the first panel. Each of the channels has a first end that opens to the cooling passage and a second end that opens to the exterior gas path side at one of the side edges. Each of the channels is defined on a side by the bearing surface of the second panel. 1. An airfoil comprising:a core structure defining a cooling passage;first and second panels attached with the core structure, each of the first and second panels including an exterior gas path side, an opposed interior side, and a side edge, the side edge having a bearing portion that defines a bearing surface, the bearing surface of the first panel abutting the bearing surface of the second panel at a bearing interface; andchannels in the bearing portion of the first panel, each of the channels having a first end opening to the cooling passage and a second end opening to the exterior gas path side at one of the side edges, and each of the channels being defined on a side by the bearing surface of the second panel.2. The airfoil as recited in claim 1 , wherein the channels are notches.3. The airfoil as recited in claim 2 , wherein the each of the notches is elongated and sloped.4. The airfoil as recited in claim 1 , wherein the first and second panels overlap at the bearing interface.5. The airfoil as recited in claim 1 , wherein the airfoil includes an airfoil section and a platform claim 1 , and the first and second panels are on the platform.6. The airfoil as recited in claim 5 , wherein the each of the notches is elongated and sloped.7. The ...

GAS TURBINE ENGINE WITH INTERCOOLED COOLING AIR AND TURBINE DRIVE

A gas turbine engine has a compressor section with a downstream most end and a cooling air tap at a location spaced upstream from the downstream most end. The cooling air tap is passed through at least one boost compressor and at least one heat exchanger, and then passed to a turbine section to cool the turbine section. The boost compressor is driven by a driveshaft which is driven by the turbine section. A boost turbine selectively drives the boost compressor. 1. A gas turbine engine comprising:a compressor section having a downstream most end and a cooling air tap at a location spaced upstream from said downstream most end; andsaid cooling air tap being passed through at least one boost compressor and at least one heat exchanger, and then passed to a turbine section to cool said turbine section, said boost compressor being driven by a driveshaft which is driven by said turbine section, and a boost turbine for selectively driving said boost compressor; anda clutch positioned between said boost turbine and said boost compressor and said clutch being selectively opened or closed to provide said selective drive of said boost compressor by said boost turbine.2. The gas turbine engine as set forth in claim 1 , wherein when said boost turbine is driving said boost compressor claim 1 , rotation passes back into said driveshaft.3. The gas turbine engine as set forth in claim 2 , wherein a second clutch is provided between the driveshaft and said boost compressor claim 2 , such that both said driveshaft and said boost turbine can be selectively connected or disconnected claim 2 , from said boost compressor.4. The gas turbine engine as set forth in claim 1 , wherein a second clutch is provided between the driveshaft and said boost compressor claim 1 , such that both said driveshaft and said boost turbine can be selectively connected or disconnected claim 1 , from said boost compressor.5. The gas turbine engine as set forth in claim 1 , wherein there are a plurality of said ...

ENGINE DRIVEN BY SC02 CYCLE WITH INDEPENDENT SHAFTS FOR COMBUSTION CYCLE ELEMENTS AND PROPULSION ELEMENTS

A gas turbine engine includes a first shaft coupled to a first turbine and a first compressor, a second shaft coupled to a second turbine and a second compressor, and a third shaft coupled to a third turbine and a fan assembly. The turbine engine includes a heat rejection heat exchanger configured to reject heat from a closed loop system with air passed from the fan assembly, and a combustor positioned to receive compressed air from the second compressor as a core stream. The closed-loop system includes the first, second, and third turbines and the first compressor and receives energy input from the combustor. 120-. (canceled)21. A gas turbine engine , comprising:a first shaft coupled to a first turbine and a first compressor;a second shaft offset axially from the first shaft, the second shaft coupled to a second turbine and a second compressor;a heat rejection heat exchanger configured to reject heat from a closed loop system to bypass air of the gas turbine engine;a combustor positioned to receive air from the second compressor; includes the first and second turbines and the first and second compressors;', 'receives energy input from the combustor; and', 'provides power from the combustor to the first and second turbines., 'wherein the closed-loop system22. The gas turbine engine of claim 21 , wherein the first and second shafts are coaxially aligned with one another.23. The gas turbine engine of claim 21 , wherein the closed-loop system includes carbon dioxide as a working fluid.24. The gas turbine engine of claim 21 , further comprising a third shaft coupled to a third turbine and a fan assembly claim 21 , wherein the bypass air is passed from the fan assembly to provide cooling to the heat rejection heat exchanger.25. The gas turbine engine of claim 24 , further comprising a gear coupled to the third shaft that reduces a rotation of a fan blade within the fan assembly relative to a rotational speed of the third shaft.26. The gas turbine engine of claim 24 , ...

INTEGRATED HEAT EXCHANGERS FOR LOW FAN PRESSURE RATIO GEARED TURBOFAN

An oil cooling system and method are provided for use with respect to a lubricated mechanical system within a bypass configured gas turbine engine. A surface cooler is fluidly linked to the lubricated mechanical system to receive oil from the lubricated mechanical system for cooling and reuse. In an embodiment, the surface cooler is mounted on an existing surface within the bypass airflow path of the bypass configured gas turbine engine to provide effective cooling while avoiding the introduction of additional aerodynamic disturbances in the bypass path. In an embodiment, the surface cooler is mounted on the fan casing or on a fan exit guide vane. 1. A cooling system for a bypass configured gas turbine engine having a bypass airflow path and a fan exit guide vane extending through the bypass airflow path , the cooling system comprising:a lubricated mechanical system having an internal cavity adapted to contain lubricating oil;a surface cooler operably connected to the fan exit guide vane and extending into the bypass airflow path beyond a surface of the fan exit guide vane; andone or more fluid conduits linking the internal cavity of the lubricated mechanical system to the surface cooler such that the surface cooler reduces a temperature of the lubricating oil.2. The cooling system for a bypass configured gas turbine engine in accordance with claim 1 , wherein the lubricated mechanical system comprises a fan drive gear reduction architecture powering a fan of the bypass configured gas turbine engine.3. A method for cooling lubricating oil used in a lubricated mechanical system associated with a bypass configured gas turbine engine having a bypass airflow path and the cooling system for the bypass configured gas turbine engine of claim 1 , the method comprising:mounting the surface cooler to a structure within the bypass airflow path, within the bypass configured gas turbine engine; androuting the lubricating oil from the lubricated mechanical system to the surface ...

BLADE TIP COOLING ARRANGEMENT

A turbine blade according to an example of the present disclosure includes, among other things, a platform extending from a root section, an airfoil section extending radially from the platform to an airfoil tip, a plurality of cooling passages defined in an external wall of the airfoil tip, the plurality of cooling passages extending radially between the airfoil tip and a cavity in the airfoil section bounded by the external wall, and each of the plurality of cooling passages defining an inlet port along the cavity and an exit port adjacent the airfoil tip, and at least one internal feature within each of the plurality of cooling passages that meter flow to the respective exit port. 1. A turbine blade for a gas turbine engine comprising:a platform extending from a root section;an airfoil section extending radially from the platform to an airfoil tip;a plurality of cooling passages defined in an external wall of the airfoil tip, the plurality of cooling passages extending radially between the airfoil tip and a cavity in the airfoil section bounded by the external wall, and each of the plurality of cooling passages defining an inlet port along the cavity and an exit port adjacent the airfoil tip; andat least one internal feature within each of the plurality of cooling passages that meter flow to the respective exit port.2. The turbine blade of claim 1 , wherein the at least one internal feature is spaced apart from the inlet and exit ports.3. The turbine blade of claim 1 , wherein at least one throat is defined between the at least one internal feature and a wall of a respective one of the plurality of the cooling passages.4. The turbine blade of claim 2 , wherein each of the cooling passages defines a passage axis that extends between the respective inlet and exit ports claim 2 , and a projection of the passage axis extends through a radially outermost boundary of the airfoil tip.5. The turbine blade of claim 4 , wherein the exit port is defined along a radially ...

Refractory metal core finishing technique

A refractory metal core (RMC) finishing method according to an exemplary aspect of the present disclosure includes, among other things, performing a plurality of finishing operations on a plurality of RMC samples, analyzing one or more properties of at least a portion of the plurality of RMC samples and selecting a combination of finishing operations for generating an RMC having desirable properties for manufacturing a part free from defects.

Ceramic Matrix Composite Component Cooling

Ceramic matrix composite (CMC) airfoils and methods for forming CMC airfoils are provided. In one embodiment, an airfoil is provided that includes opposite pressure and suction sides extending radially along a span and opposite leading and trailing edges extending radially along the span. The leading edge defines a forward end of the airfoil, and the trailing edge defines an aft end of the airfoil. A trailing edge portion is defined adjacent the trailing edge at the aft end, and a pocket is defined in and extends within the trailing edge portion. A heat pipe is received in the pocket. A method for forming an airfoil is provided that includes laying up a CMC material to form an airfoil preform assembly; processing the airfoil preform assembly; defining a pocket in a trailing edge portion of the airfoil; and inserting a heat pipe into the pocket. 120.-. (canceled)21. An airfoil for a gas turbine engine , comprising:opposite pressure and suction sides extending radially along a span;opposite leading and trailing edges extending radially along the span, the pressure and suction sides extending axially between the leading and trailing edges, the leading edge defining a forward end of the airfoil, the trailing edge defining an aft end of the airfoil;a trailing edge portion defined adjacent the trailing edge at the aft end;at least two pockets defined in the trailing edge portion, the at least two pockets extending within the trailing edge portion, the at least two pockets radially spaced apart from one another; anda heat pipe received in each pocket.22. The airfoil of claim 21 , wherein at least one pocket of the at least two pockets extends at an angle with respect to an axial direction.23. The airfoil of claim 22 , wherein the airfoil defines adjacent the leading edge a cavity for receipt of a flow of cooling fluid claim 22 , and wherein a condenser portion of the heat pipe is positioned near the cavity.24. The airfoil of claim 21 , wherein at least one pocket of the at ...

STAGGERED CORE PRINTOUT

A core for gas turbine engine component comprises a body extending between first and second ends to define a length, and extending between first and second edges to define a width. A plurality of core extensions are formed as part of the body. The plurality of core extensions are positioned to be staggered relative to each other such that at least two adjacent core extensions are variable relative to each other in at least one dimension. A gas turbine engine component is also disclosed. 1. A method of manufacturing a gas turbine engine component comprising:providing a body extending between first and second ends to define a length and extending between first and second edges to define a width;forming a plurality of openings within a wall surface of the body wherein the plurality of openings are positioned to be staggered relative to each other such that at least two adjacent openings are offset from each other in at least one direction;defining an arc segment along the length of the body, wherein the openings are spaced apart from each other along the arc segment; andforming the openings via one of a casting, EDM, laser, or additive manufacturing method.2. The method according to wherein the body comprises an airfoil with the first edge comprising a leading edge claim 1 , the second edge comprising a trailing edge claim 1 , the first end comprising a radially inner portion and the second end comprising a tip portion claim 1 , and including having a point along one of the first and second edges that defines a radius of curvature for the arc segment.3. The method according to claim 2 , wherein the point is located at a position that is nearer to the leading edge than the trailing edge.4. The method according to wherein the point is located at a position that is nearer to the trailing edge than the leading edge.5. The method according to wherein the point is located at a position that is nearer to a pressure side than a suction side.6. The method according to wherein ...

INTERCOOLED COOLING AIR WITH DUAL PASS HEAT EXCHANGER

A gas turbine engine includes a main compressor. A tap is fluidly connected downstream of the main compressor. A heat exchanger is fluidly connected downstream of the tap. An auxiliary compressor unit is fluidly connected downstream of the heat exchanger. The auxiliary compressor unit is configured to compress air cooled by the heat exchanger with an overall auxiliary compressor unit pressure ratio between 1.1 and 6.0. An intercooling system for a gas turbine engine is also disclosed. 1. A gas turbine engine comprising;a main compressor;a tap fluidly connected downstream of said main compressor;a heat exchanger fluidly connected downstream of said tap; andan auxiliary compressor unit fluidly connected downstream of said heat exchanger, wherein said auxiliary compressor unit is configured to compress air cooled by said heat exchanger with an overall auxiliary compressor unit pressure ratio between 1.1 and 6.0.2. The gas turbine engine as set forth in claim 1 , wherein a pressure ratio across said auxiliary compressor unit is greater than or equal to 4.1 and less than or equal to 6.0.3. The gas turbine engine as set forth in claim 1 , wherein said at least one of said more upstream locations is in a low pressure compressor4. The gas turbine engine as set forth in claim 2 , wherein a pressure ratio across said auxiliary compressor unit being greater than or equal to 1.1 and less than or equal to 4.0.5. The gas turbine engine as set forth in claim 4 , wherein said at least one of said more upstream locations is in a high pressure compressor.6. The gas turbine engine as set forth in claim 1 , wherein a main fan delivers bypass air into a bypass duct and into said main compressor section and said heat exchanger positioned within said bypass duct to be cooled by bypass air.7. The gas turbine engine as set forth in claim 1 , wherein said at least one of said more upstream locations is in a high pressure compressor.8. The gas turbine engine as set forth in claim 1 , wherein ...

HIGH TEMPERATURE DISK CONDITIONING SYSTEM

A gas-circulation system for conditioning a disk of an aircraft may comprise a first takeoff port configured to extract a combusted gas and a second takeoff port configured to extract an uncombusted gas. A first valve may comprise an inlet in fluid communication with the first and second takeoff ports and an outlet of the first valve in fluid communication with the disk. 1. A gas-circulation system for conditioning a compressor disk of an aircraft , the gas-circulation system comprising: commanding, by the controller, an actuator to open and close a valve that outputs a mixed gas, wherein the mixed gas comprises at least one of an uncombusted gas and a combusted gas;', 'determining, by the controller, an engine speed;', 'comparing, by the controller, the engine speed to closing criteria; and', 'concluding, by the controller, that a failure occurred in the valve if a state of the valve is open in response to the engine speed being within the closing criteria., 'a non-transitory memory communicating with a controller, the non-transitory memory having instructions stored thereon that, in response to execution by the controller, cause a processor to perform operations comprising2. The gas-circulation system of claim 1 , wherein the instructions further comprise:determining, by the controller, an operational state of the aircraft; andcommanding, by the controller, the actuator to open the valve in response to the operational state corresponding to a predetermined operational state.3. The gas-circulation system of claim 1 , wherein the instructions further comprise commanding claim 1 , by the controller claim 1 , the actuator to close the valve in response to the engine speed matching the closing criteria.4. The gas-circulation system of claim 1 , wherein the instructions further comprise determining the state of the valve in response to a measurement from a thermocoupling.5. The gas-circulation system of claim 1 , wherein the instructions further comprise responding ...

Abradable Material

A blade outer airseal comprising a body having: an inner diameter (ID) surface; an outer diameter (OD) surface; a leading end; a trailing end; a metallic substrate; and a coating system atop the substrate along at least a portion of the inner diameter surface. At least over a first area of the inner diameter surface, the coating system comprises an abradable layer comprising a metallic matrix and a filler. The filler forms at least 20% by volume of the abradable layer with agglomerates or aggregates of oxide particles, the oxide particles having a D50 size ≤200 nm. 117.-. (canceled)19. The method of wherein the D50 size is:10 nm to 50 nm.20. The method of wherein:the metallic matrix is sprayed from a source having particles of the matrix with a D50 size of 22-90 micrometers.21. The method of wherein:the filler is sprayed from the source having the aggregates.22. The method of wherein:the particles have a D50 size ≤200 nm form at least 50 weight percent of the aggregates.23. The method of further comprising forming the aggregates by calcining and sintering agglomerates.24. The method of further comprising forming the agglomerates by drying a slurry.25. The method of wherein the filler comprises:alumina, silica, titania, zirconia, hafnia, dysprosia, gadolinia, yttria, magnesia, nickel oxide, and/or chromia forming at least 40% by volume of the abradable layer.26. The method of wherein the filler comprises:alumina and magnesia forming aggregates that occupy at least 40% by volume of the abradable layer.27. The method of wherein:the magnesia forms 0.1% to 2% weight percent of the total of alumina and magnesia.28. The method of wherein the metallic matrix comprises an MCrAlY.29. The method of wherein the metallic matrix comprises an MCrAlY.30. The method of wherein the filler comprises:alumina, silica, titania, zirconia, hafnia, dysprosia, gadolinia, yttria, magnesia, nickel oxide, and/or chromia forming at least 40% by volume of the abradable layer.31. The method of ...

MANUFACTURING METHOD FOR POROUS THERMAL INSULATION COATING LAYER, POROUS THERMAL INSULATION COATING LAYER AND INTERNAL COMBUSTION ENGINE USING THE SAME

Disclosed are a manufacturing method for a porous thermal insulation coating layer, a porous thermal insulation coating layer with substantially reduced thermal conductivity and volumetric heat capacity and an internal combustion engine including the porous thermal insulation coating layer thereby having excellent durability. 1. A method of manufacturing a porous thermal insulation coating layer comprising:coating a reaction product obtainable from a reaction of metal alkoxide containing at least one selected from the group consisting of aluminum, zirconia, titanium and silicon with alcohol and water;drying the coated reaction product at a first temperature; andperforming a thermal treatment at a second temperature that is greater than the first temperature and less than about 300° C.2. The method of claim 1 , wherein the second temperature ranges from about 20° C. to about 220° C. and is greater than the first temperature.3. The method of claim 1 , whereinthe performing of the thermal treatment at the second temperature is conducted for about 12 hours to 48 hours.4. The method of claim 1 , wherein the first temperature ranges from about 30° C. to about 100° C.5. The method of claim 1 , wherein the second temperature ranges from about 100° C. to about 250° C.6. The method of claim 1 , wherein an amount of about 10 to 100 parts by weight of the alcohol relative to 100 parts by weight of the metal alkoxide of the metal is reacted.7. The method of claim 1 , wherein an amount of about 110 to 500 parts by weight of the water relative to 100 parts by weight of the metal alkoxide of the metal is reacted.8. The method of claim 1 , further comprising claim 1 ,before the drying of the coated reaction product at the first temperature, adding a solution including a silane-based compound to the product of the coating step.9. The method of claim 8 , wherein the silane-based compound comprises a silane compound substituted with at least one functional group selected from the group ...

AIRFOIL TIP POCKET WITH AUGMENTATION FEATURES

A component for a gas turbine engine includes, among other things, an airfoil that includes a pressure sidewall and a suction sidewall that meet together at both a leading edge and a trailing edge, the airfoil extending radially from a platform to a tip, a tip pocket formed in the tip and terminating prior to the trailing edge, and one or more heat transfer augmentation devices formed in the tip pocket. 1. A component for a gas turbine engine comprising:an airfoil that includes a pressure sidewall and a suction sidewall that meet together at both a leading edge and a trailing edge, the airfoil extending radially from a platform to a tip;a tip pocket formed in the tip and terminating prior to the trailing edge;wherein the tip pocket includes a suction side lip, a pressure side lip, a leading edge lip and a trailing edge lip that each extend radially outwardly from a floor;one or more heat transfer augmentation devices formed in the tip pocket, the one or more heat transfer augmentation devices including a first device extending from one of the suction side lip and the pressure side lip such that a wall of the first device is spaced apart from another one of the suction side lip and the pressure side lip; anda plurality of cooling holes defined in the floor that fluidly connect the tip pocket to at least one internal cooling cavity formed inside the airfoil.2. The component as recited in claim 1 , wherein the first device is spaced apart from the floor.3. The component as recited in claim 1 , wherein the first device axially extends across the one of the suction side lip and the pressure side lip.4. The component as recited in claim 3 , wherein the first device axially extends from the leading edge lip to the trailing edge lip.5. The component as recited in claim 3 , wherein the first device is a first rib radially spaced apart from the floor such that the rib radially bounds a portion of the tip pocket.6. The component as recited in claim 5 , wherein the one or more ...

BAFFLE FOR A COMPONENT OF A GAS TURBINE ENGINE

A method of repairing an airfoil according to an example of the present disclosure includes, among other things, providing an airfoil body, the airfoil body having external walls extending between a leading edge and a trailing edge, providing a baffle, the baffle including a baffle body defining an internal passage, and sidewalls of the baffle body defining a first contour, defining a cavity in the airfoil body, the cavity extending inwardly from the external walls to define a second contour complementary to the first contour, and inserting the baffle into the cavity. An airfoil arrangement is also disclosed. 1. A method of repairing an airfoil , comprising:providing an airfoil body, the airfoil body having external walls extending between a leading edge and a trailing edge;providing a baffle, the baffle including a baffle body defining an internal passage, and sidewalls of the baffle body defining a first contour;defining a cavity in the airfoil body, the cavity extending inwardly from the external walls to define a second contour complementary to the first contour; andinserting the baffle into the cavity.2. The method as recited in claim 1 , wherein:the step of defining the cavity includes removing material from the trailing edge to define an opening to the cavity; andthe sidewalls of the baffle body are spaced apart by an exit wall to define one or more exit ports situated adjacent to the opening.3. The method as recited in claim 2 , wherein the inserting step occurs subsequent to the removing material from the trailing edge to define the opening to the cavity.4. The method as recited in claim 1 , wherein the airfoil body is made of a first material claim 1 , and the baffle body is made of a second claim 1 , different material.5. The method as recited in claim 1 , wherein the sidewalls of the baffle body abut a majority of surfaces of the airfoil body that define the cavity.6. The method as recited in claim 5 , wherein the internal passage along the sidewalls is ...

METHOD OF MONITORING AIRCRAFT BRAKE PERFORMANCE AND APPARATUS FOR PERFORMING SUCH A METHOD

A brake performance monitoring system operates by an energy differential calculated from brake demand energy and energy absorbed during a braking operation. A significant differential would be reported as a possible problem with the braking system. 1. An airplane , comprising: a wheel;', 'a strut; and', 'a brake, wherein, 'landing gear includingthe airplane includes sensors configured to sense phenomenon associated with braking the airplane, and wherein the airplane is configured to calculate energy absorbed by the wheel brake.2. The airplane of claim 1 , wherein the airplane includes a brake pedal.3. The airplane of claim 1 , wherein the airplane includes a brake pedal configured to be operated by a pilot of the airplane.4. The airplane of claim 1 , further comprising a hydraulic master cylinder.5. The airplane of claim 1 , further comprising a servomechanism comprising a piston arranged to pressurize a working chamber of the airplane.6. The airplane of claim 1 , wherein the brake includes a brake caliper arranged to slow the wheel.7. The airplane of claim 1 , wherein the brake includes a brake caliper arranged to slow the wheel to decelerate the airplane.8. The airplane of claim 1 , further comprising a pressure transducer.9. The airplane of claim 1 , wherein the landing gear includes a speed sensor.10. The airplane of claim 1 , wherein the brake is configured to generate heat when braking.11. An aircraft landing gear braking system comprising: a wheel; and', 'a wheel brake for braking the wheel; and, 'an aircraft landing gear comprising a memory;', 'a first sensor arranged to measure wheel brake control input variable data over a braking event and send the wheel brake control input variable data to the memory;', 'a second sensor arranged to measure wheel brake performance output variable data over the braking event and send the wheel brake performance output variable data to the memory; and', (i) calculate a total pseudo-energy demand to be absorbed by the wheel ...

GAS ENGINE COMPONENT WITH COOLING PASSAGES IN WALL AND METHOD OF MAKING THE SAME

A method of manufacturing a gas turbine engine component and said resulting component configuration are disclosed. The method includes providing a ceramic structure including a sacrificial ceramic core. The ceramic structure is inserted into a casting mold. A component material is introduced into the casting mold. The ceramic structure is removed after the component material has solidified to define a casted component having a cooling cavity of a heat exchanger segment formed in the wall structure of the component. The ceramic body of the sacrificial ceramic core defines a plurality of apertures and includes a trailing first pin and a leading first pin. The trailing first pin and the leading first pin are extended away from an inner surface of the body, and the leading first pin is spaced from the trailing first pin and in closer proximity to a leading edge of the body than the trailing first pin. 1. A method of manufacturing a gas turbine engine component , comprising:providing a ceramic structure including a sacrificial ceramic core, the sacrificial ceramic core having a ceramic body corresponding to a cooling cavity of a heat exchanger segment, the ceramic body having an exterior surface comprising an outer surface and an inner surface interconnected to one another by a leading edge, a trailing edge, a tip facing edge and a base facing edge, the ceramic body defining a plurality of apertures extending between the outer surface and the inner surface, the plurality of apertures corresponding to a plurality of pedestals within the cooling cavity of the heat exchanger segment, the ceramic body including a trailing first pin and a leading first pin corresponding to a first inlet orifice and a second inlet orifice leading to the cooling cavity of the heat exchanger segment, the trailing first pin and the leading first pin extending away from the inner surface, the leading first pin spaced from the trailing first pin and in closer proximity to the leading edge than the ...

LOW TURN LOSS BAFFLE FLOW DIVERTER

An airfoil according to an example of the present disclosure includes, among other things, an airfoil body having an internal passage for conveying a fluid flow, the internal passage including first and second passage sections coupled at a turn section. A baffle includes an elongated body arranged in the second passage section to define a pair of opposed cooling flow paths that extend from the turn section along a common length of the second passage section, and a first wedge region extending from the elongated body into the first passage section such that the fluid flow is directed through the turn section between the first passage section and the pair of cooling flow paths. 1. An airfoil , comprising:an airfoil body having an internal passage for conveying a fluid flow, the internal passage including first and second passage sections coupled at a turn section; anda baffle including an elongated body arranged in the second passage section to define a pair of opposed cooling flow paths that extend from the turn section along a common length of the second passage section, and a first wedge region extending from the elongated body into the first passage section such that the fluid flow is directed through the turn section between the first passage section and the pair of cooling flow paths.2. The airfoil as recited in claim 1 , wherein the elongated body is dimensioned such that opposed walls of the elongated body abut against opposed walls of the internal passage to space apart the first and second passage sections.3. The airfoil as recited in claim 2 , wherein an apex of the first wedge region is oriented in a direction towards an outer wall of the turn section.4. The airfoil as recited in claim 2 , wherein the first wedge region includes first and second sloped sides that are joined at an apex claim 2 , the first sloped side defining a first reference plane intersecting surfaces of the turn section to define an acute angle.5. The airfoil as recited in claim 2 , ...

TRAILING EDGE PRESSURE AND FLOW REGULATOR

A gas turbine engine component comprises a body having a leading edge, a trailing edge, and a radial span. One internal channel in the body provides an upstream supply pressure. Another internal channel in body receives the upstream supply pressure and provides a downstream supply pressure. At least one axial rib separates an internal area adjacent to the trailing edge into a plurality of individual cavities. At least one pressure regulating feature is located at an entrance to at least one individual cavity entrance to control downstream supply pressure to the trailing edge. Exits formed in the trailing edge communicate with an exit pressure. The rib and pressure regulating features cooperate such that the downstream supply pressure mimics the exit pressure along the radial span. A method of manufacturing a gas turbine engine component and a method of controlling flow in a gas turbine engine component are also disclosed. 1. A gas turbine engine component comprising:a body having a leading edge extending to a trailing edge, and extending from a radially inner end to a radially outer end to define a radial span;a first internal channel formed within the body adjacent the leading edge;a second internal channel formed within the body downstream of the first internal channel to provide an upstream supply pressure;a third internal channel formed within the body adjacent the trailing edge and that receives the upstream supply pressure and provides a downstream supply pressure;a radial wall that separates the first internal channel from the second internal channel;at least one axial rib that separates the third internal channel into a plurality of individual cavities;at least one pressure regulating feature located at an entrance to at least one individual cavity to control the downstream supply pressure to the trailing edge; anda plurality of exits formed in the trailing edge that communicate with an exit pressure, and wherein the at least one axial rib and the at least ...

AIRFOIL FOR A TURBINE ENGINE

A method and apparatus for minimizing engine weight for a turbine engine by utilizing one or more discrete protuberances disposed on an engine component wall. The wall can have a nominal thickness to decrease engine weight while the protuberances can provide increased discrete thicknesses for providing one or more cooling holes. The increased thickness at the protuberances provides for an increased thickness to provide sufficient length to increase cooling hole effectiveness. 1. A component for a turbine engine , with the turbine engine generating a hot combustion gas flow and providing a cooling fluid flow , the component comprising:a wall having a nominal thickness separating the hot combustion gas flow from the cooling fluid flow having a hot surface facing the hot combustion gas flow and a cool surface facing the cooling fluid flow;at least one localized, protuberance extending from the cool surface; anda film hole extending through the protuberance and the wall, having a length greater than the nominal thickness of the wall.2. The component of wherein the film hole is angled relative to a local normal between the hot surface and the cool surface.3. The component of wherein the film hole is non-linear.4. The component of wherein the film hole includes an inlet and an outlet claim 1 , having a passage connecting the inlet to the outlet.5. The component of wherein at least one of the inlet claim 4 , outlet claim 4 , or passage is shaped.6. The component of wherein the protuberance includes an upstream side and a downstream side having the inlet disposed on the upstream side.7. The component of wherein the protuberance includes a height and the height is at least 50% of the nominal thickness.8. The component of wherein the height is at least 100% of the nominal thickness.9. The component of wherein the protuberance is symmetrical about an axis parallel to a direction of the cooling fluid flow within the component.10. The component of wherein the protuberance ...

Method And System For Dynamic Control Of Game Audio Based On Audio Analysis

A gaming device, during play of a particular game, receives a plurality of audio channels carrying game sounds and determines whether a first of the game sounds overwhelms a second of the game sounds. In response to the determination, the gaming device adjusts a level of one or more of the audio channels such that perceptibility of the second game sound is improved relative to perceptibility of the first game sound prior to the adjustment, wherein for the adjustment, a level of one or more of the audio channels carrying the first game sound is decreased while a level of one or more of the audio channels carrying the second game sound is maintained or increased. The audio channels include three or more audio channels and the adjustment of the level of the audio channels is performed while the three or more audio channels are combined into two stereo channels. 1. A method , comprising: receiving a plurality of surround sound channels carrying game sounds, said plurality of surround sound channels comprising three or more surround sound channels;', 'determining that a first of said game sounds has a higher loudness relative to a second of said game sounds; and', 'in response to said determining, adjusting a power level of one or more of said plurality of surround sound channels such that loudness of said second one of said game sounds relative to said first one of said game sounds is increased., 'in a gaming device during play of a particular game2. The method according to claim 1 , comprising:detecting when said first one of said game sounds ceases to be too loud relative to said second one of said game sounds; andin response to said detecting, adjusting said power level of said one or more of said plurality of surround sound channels.3. The method according to claim 1 , wherein said determining occurs dynamically based on acquired audio information for said particular game.4. The method according to claim 3 , comprising storing said audio information for said ...

COMPARTMENTALIZATION OF COOLING AIR FLOW IN A STRUCTURE COMPRISING A CMC COMPONENT

A structure in a gas turbine engine comprises a spar and a CMC component adjoining the spar and separated from the spar by a cavity supplied by cooling air. At least one rope seal is installed in the cavity within a groove made in the spar to thus compartmentalize the cavity and control the flow of cooling air. 120-. (canceled)21. A gas turbine engine system for compartmentalized airflow , the system comprising:a static metal component; anda ceramic matrix composite (CMC) component forming a cavity with the static metal component, the cavity being divided into sections with respective passages for receiving cooling air into the cavity through the static metal component and removing cooling air from the cavity through the CMC component.22. The system of claim 21 , wherein the static metal component includes at least one of a vane spar or a combustor liner.23. The system of claim 21 , further comprising an impingement hole in each section to receive airflow in the respective passages and an exit hole in each section to remove airflow from the respective passages.24. The system of claim 21 , further comprising a rope seal dividing the cavity into the respective passages.25. The system of claim 21 , further comprising a rope seal positioned in a moon-shaped groove formed in the static metal component.26. The system of claim 21 , further comprising a rope seal positioned in a groove formed between two raised landings with curved sidewalls extending from the static metal component.27. The system of claim 21 , wherein the respective passages are formed with a rope seal that includes at least one of aluminosilicate or aluminum oxide.28. A method of compartmentalizing airflow in a gas turbine engine system claim 21 , the method comprising:providing a static metal component;providing a ceramic matrix composite (CMC) component forming a cavity with the static metal component, the cavity being divided into a plurality of sections with respective passages to provide airflow ...

TURBINE AIRFOILS WITH MICRO COOLING FEATURES

A blade used in a gas turbine engine includes a pair of pedestals and an airfoil coupled between the pedestals. The airfoil includes cooling features to cool the airfoil. 1. An airfoil for use in a gas turbine engine and having a pressure side and a suction side , the airfoil comprisinga spar that extends radially relative to an axis and formed to define a cooling air plenum adapted to receive a flow of cooling air, anda skin coupled to an exterior surface of the spar and positioned to at least partially cover the spar along the pressure side and the suction side,wherein a first plurality of axially extending grooves is formed in the exterior surface of the spar on the suction side that defines a first plurality of cooling passageways between the spar and the skin, a second plurality of axially extending grooves is formed in the exterior surface of the spar on the pressure side that defines a second plurality of cooling passageways between the spar and the skin,wherein a first plurality of inlet ports is formed in the spar and in fluid communication with the cooling air plenum and the first plurality of cooling passageways to pass the flow of cooling air into the first plurality of cooling passageways from the cooling air plenum, a second plurality of inlet ports is formed in the spar and in fluid communication with the cooling air plenum and the second plurality of cooling passageways to pass the flow of cooling air into the second plurality of cooling passageways from the cooling air plenum, andwherein the first plurality of cooling passageways extending between the first plurality of inlet ports and a first plurality of outlet slots positioned axially aft of the first plurality of inlet ports, the second plurality of cooling passageways extending between the second plurality of inlet ports and a second plurality of outlet slots positioned axially aft of the second plurality of inlet ports, and the first and second plurality of outlet slots are configured to pass ...

VARIABLE VANE SEGMENT

A variable vane pack includes an inner platform, an outer platform, radially outward of the inner platform, a plurality of vanes connecting the inner platform to the outer platform, wherein the outer platform comprises a platform body and an impingement plate, the impingement plate having a radially inward impingement plate, a radially outward pressure distribution plate, and an impingement plenum defined between the radially inward impingement plate and the radially outward pressure distribution plate. 1. A method for cooling an engine component comprising:providing a cooling air feed to said engine components;passing said cooling air through a pressure distribution plate into an impingement plenum, thereby providing an even distribution of air pressure to said impingement plenum;feeding said cooling air through impingement openings in an impingement plate, thereby impinging cooling air on a radially outward platform of said engine component and cooling said radially outward platform.2. The method of claim 1 , wherein passing said cooling air through a pressure distribution plate into an impingement plenum comprises passing cooling air through a plurality of slots in the distribution plate.3. The method of claim 2 , wherein each of the slots overlaps at least one adjacent slot along an axis defined by the curvature of the engine component.4. The method of claim 2 , wherein each of the slots circumferentially overlaps at least one adjacent slot.5. The method of claim 1 , wherein cooling air exits said impingement plenum at least one of a joint between said engine component and an adjacent engine component claim 1 , and a plurality of openings in said radially outward platform.6. The method of claim 1 , wherein feeding said cooling air through a plurality of impingement openings in an impingement plate claim 1 , comprises providing a plurality of impingement airflows to said radially outward platform from said plurality of impingement openings and wherein each of ...