| Entries |
| Document | Title | Date |
|---|
| 20080295521 | METHOD AND APPARATUS FOR ASSEMBLING TURBINE ENGINES - A method of assembling a turbine engine includes defining a first chamber and defining a second chamber. The method also includes forming at least one venturi device oriented with a predetermined venturi step angle greater than approximately 48°. The method further includes coupling the first chamber in flow communication with the second chamber via the venturi device therebetween. | 12-04-2008 |
| 20090120103 | METHOD AND APPARATUS FOR SUB SEA POWER GENERATION - A method for generating power at sub sea level. An oxidizing fluid and oxidizer are fed separately for mixing at a sub sea station. The oxidizing fluid and oxidizer are chosen for chemical reaction in situ under the release of energy. The energy so produced is fed to drive means operated by at least one of heat, kinetic energy, pressure and electricity, and operative to drive sub sea processes or/and sub sea production equipment. An apparatus for generating power at sub sea level. The apparatus includes separate supplies of an oxidizing fluid and an oxidizer for mixing at a sub sea station. The oxidizing fluid and oxidizer are chosen for chemical reaction in situ under the release of energy. A feed for feeding the energy so produced to a drive operated by at least one of heat, kinetic energy, pressure and electricity. | 05-14-2009 |
| 20100095684 | MORPHABLE COMPOSITE STRUCTURE - A morphable composite structure is disclosed herein. The morphable composite structure includes at least first and second layers fixed relative to one another. Each of the first and second layers includes a plurality of structural fibers oriented substantially parallel to one another and a quantity of binder substantially fixing the plurality of structural fibers together. The plurality of fibers of the first layer is oriented asymmetrically to the plurality of fibers of the second layer. The morphable composite structure also includes at least one pattern of electrically-conductive particles connected with the first layer and spaced from the second layer. A current can be directed through the pattern to heat the plurality of fibers of the first layer and change the shape of the morphable composite structure. | 04-22-2010 |
| 20100175389 | Apparatus For Filtering Gas Turbine Inlet Air - An inlet air filtration system for a gas turbine includes, in an exemplary embodiment, an air plenum, and a plurality of filter elements mounted inside the air plenum, with each filter element including a support structure. The inlet air filtration system also includes a plurality of electrodes positioned proximate the plurality of filter elements, where the electrodes are coupled to a power source which supplies a voltage to the electrodes. The voltage is sufficient to establish an electrostatic field between the electrodes and the filter elements, and is sufficient to produce a corona discharge from the electrodes, wherein an amount of current applied to the filter elements is about 0.1 μA/ft | 07-15-2010 |
| 20110067414 | FLOW DISCOURAGING SYSTEMS AND GAS TURBINE ENGINES - A flow discouraging system includes a stator assembly, fins, and a rotor assembly. The stator assembly includes stationary components forming a side wall including an annular groove defined by an outer axially-extending surface, an inner axially-extending surface, and a radial surface. One or more outer axial fins disposed in the annular groove extend along the outer axially-extending surface of the side wall. One or more inner axial fins disposed in the annular groove extend along the inner axially-extending surface of the side wall. One or more radial fins disposed in the annular groove extend axially from the radial surface of the side wall. The rotor assembly is disposed adjacent to and is spaced apart from the stator assembly to form a portion of a cavity and includes an annular rim extending at least partially into the annular groove and disposed between the outer and inner axial fins. | 03-24-2011 |
| 20110072831 | SEALING APPARATUS WITH MULTISTAGE BRUSH SEAL - An apparatus is provided for sealing a channel fluidly connecting between a high pressure zone and a low pressure zone. The apparatus has at least two brush seal elements provided in the channel in series in a direction from the high pressure zone toward the low pressure zone to define an intermediate zone between the brush seal elements. The intermediate zone is fluidly connected to the low pressure zone through a bypass or passage, allowing a fluid in the intermediate zone to flow in part through the bypass to the low pressure zone. | 03-31-2011 |
| 20110219784 | COMPRESSOR SECTION WITH TIE SHAFT COUPLING AND CANTILEVER MOUNTED VANES - A compressor section to be mounted in a gas turbine engine has a plurality of compressor rotors arranged from an upstream location toward a downstream location. A tie shaft applies an axial force at one end of the compressor section to a downstream one of the compressor rotors, and biases the compressor rotors against a hub at the opposite end. Vane sections are mounted intermediate the compressor rotors. The vane sections include at least some variable vanes driven by actuators mounted at a radially outer position and at least some of the fixed vanes are cantilever mounted , such that they are spaced from a compressor rotor, but unsecured at a radially inner end. | 09-15-2011 |
| 20120102971 | TAPERED BEARINGS - A gear support assembly for a turbine engine includes an epicyclic gear arrangement and a first tapered bearing and a second tapered bearing spaced apart from the first tapered bearing. The first tapered bearing and the second tapered bearing are arranged axially forward of the epicyclic gear arrangement and support the epicyclic gear arrangement. | 05-03-2012 |
| 20120111024 | GAS TURBINE ENGINE AND HIGH SPEED ROLLING ELEMENT BEARING - One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique gas turbine engine high speed rolling element bearing system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and high speed rolling element bearing systems for gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith. | 05-10-2012 |
| 20120111025 | System For The Generation Of Mechanical And/Or Electrical Energy - A system for the generation of energy includes a further chain of units coupled with a gas turbine plant and at least one compressor consuming mechanical energy and/or at least one generator generating electrical energy. The further chain of units comprises a closed circuit having a work fluid and at least one heat exchanger, at least one expander for expanding the work fluid and for subsequently obtaining mechanical energy for the compressor and/or generator, at least one condenser for condensing the expanded work medium, and at least one pump for conveying the work fluid. The coupling of the gas turbine plant to the further chain of units is carried out by means of the heat exchanger which is fed with heat by means of the compressor air of the compressor and starts the closed circuit through the work fluid. | 05-10-2012 |
| 20120111026 | CASTABLE HIGH TEMPERATURE ALUMINUM ALLOY - A turbine engine includes an airfoil made of an aluminum-rare earth element alloy. | 05-10-2012 |
| 20120151937 | METHOD FOR BALANCING ROTATING ASSEMBLY OF GAS TURBINE ENGINE - A method for balancing a rotating assembly of a gas turbine engine includes removing a stator vane from a section of the gas turbine engine. Removing the stator vane provides access to a rotating assembly of the gas turbine engine. The method further includes at least one of adding, removing, and repositioning a weight with respect to the rotating assembly via access to the rotating assembly provided by removing the stator vane. | 06-21-2012 |
| 20120198858 | RING ELEMENT AND TURBOMACHINE HAVING SUCH A RING ELEMENT - A ring element for a turbomachine, in particular for an aircraft gas turbine, is disclosed. The ring element has a ring element main body that has two adjacently arranged ring ends, the ring ends being connected to one another in a form-locking manner with respect to an axial plane. Also disclosed is a turbomachine having at least one such ring element. | 08-09-2012 |
| 20120247124 | CONTINUOUS RING COMPOSITE TURBINE SHROUD - A composite annular shroud supported by a support assembly including at least two single piece full 360 degree rings and at least partially disposed within an innermost one of the rings. The shroud is biased against and in sealing engagement with an inner flange of the innermost ring. A three ring assembly includes the inner ring disposed radially inwardly of a middle ring disposed radially inwardly of an outer ring and the shroud at least partially disposed within the inner ring. At least three clocking pins extend radially inwardly from the middle ring through slots in the inner ring into notches in the shroud. | 10-04-2012 |
| 20120247125 | COMMUNICATING STRUCTURE BETWEEN COMBUSTOR AND TURBINE PORTION AND GAS TURBINE - In a communicating structure between combustors that generates combustion gas inside pipe pieces and a turbine portion that generates a rotational driving force by making the combustion gas sequentially pass through a turbine stage formed of turbine stator vanes and turbine rotor blades, at least some of the first-stage turbine stator vanes closest to the combustor among the turbine stator vanes are disposed downstream of sidewalls of one pipe piece and another pipe piece that are adjacent to each other, and the distance from leading edges of the first-stage turbine stator vanes disposed downstream of the sidewalls of the pipe pieces to end portions of the sidewalls closer to the turbine portion is equal to or less than a spacing between an internal surface of the sidewall of the one pipe piece and an internal surface of the sidewall of the other pipe piece. | 10-04-2012 |
| 20120272663 | CENTRIFUGAL COMPRESSOR ASSEMBLY WITH STATOR VANE ROW - A compressor assembly for a gas turbine engine includes rotatable impeller with forward and aft ends, including: an annular hub defining a generally concave-curved cross-sectional inner flowpath surface at its radially outer periphery, the inner flowpath surface extending between an inlet and an exit; an annular array of airfoil-shaped inducer blades extending radially outward from the inner flowpath surface near the forward end; and an annular array of exducer blades extending outward from the inner flowpath surface, the array of exducer blades axially spaced apart from the array of inducer blades; a non-rotating shroud assembly surrounding the impeller and including a convex-curved outer flowpath surface, wherein the inner and outer flowpath surfaces cooperate to define a primary flowpath past the blades and the stator vanes; and an annular array of airfoil-shaped stator vanes extending radially inward from the outer flowpath surface into a space between the inducer and exducer blades. | 11-01-2012 |
| 20120285177 | SYSTEM, TRANSITION CONDUIT, AND ARTICLE OF MANUFACTURE FOR DELIVERING A FLUID FLOW - Various systems and apparatuses are provided for a flow delivery system for an engine. In one example, a system includes a first turbine providing an exhaust flow and a second turbine having an inlet and being fluidically coupled to the first turbine. The second turbine further includes a plurality of nozzle vanes positioned within the inlet of the turbine. A transition conduit is curved about an axis and coupled to the inlet and to the first turbine. The transition conduit is configured to impart an angular momentum component to at least a portion of the exhaust flow, and includes a slot that delivers at least a portion of the exhaust flow to the plurality of nozzle vanes. | 11-15-2012 |
| 20120297793 | GAS TURBINE ENGINE - A combustor ( | 11-29-2012 |
| 20130000324 | INTEGRATED CASE AND STATOR - A gas turbine engine includes a compressor, a combustor section, and a turbine. The turbine includes an integrated case/stator segment that is comprised of a ceramic matrix composite material. | 01-03-2013 |
| 20130081406 | GAS TURBINE ENGINE ROTOR STACK ASSEMBLY - A rotor stack assembly for a gas turbine engine includes a first rotor assembly and a second rotor assembly axially downstream from the first rotor assembly. The first rotor assembly and the second rotor assembly include a rim, a bore and a web that extends between the rim and the bore. A tie shaft is positioned radially inward of the bores. The tie shaft maintains a compressive load on the first rotor assembly and the second rotor assembly. The compressive load is communicated through a first load path of the first rotor assembly and a second load path of the second rotor assembly. At least one of the first load path and the second load path is radially inboard of the rims. | 04-04-2013 |
| 20130081407 | AERO-DERIVATIVE GAS TURBINE ENGINE WITH AN ADVANCED TRANSITION DUCT COMBUSTION ASSEMBLY - An aero-derivative can annular gas turbine engine having: an aero gas turbine engine core including an aero high pressure compressor ( | 04-04-2013 |
| 20130091865 | EXHAUST GAS DIFFUSER - An exhaust gas diffuser is provided and includes a peripheral body, a center body, formed to define an interior and disposed within the peripheral body to define an annulus between the peripheral body and the center body through which a first fluid flows along a main flow direction, a plurality of first members, each of which is respectively coupled to the peripheral body and the center body, to support the center body within the peripheral body and a plurality of second members, each of which extends across the annulus from the peripheral body to the center body downstream from the plurality of the first members relative to the main flow direction, to transport a second fluid to the center body interior. The plurality of the second members is circumferentially clocked relative to the plurality of the first members. | 04-18-2013 |
| 20130104565 | TURBOMACHINE INCLUDING AN INNER-TO-OUTER TURBINE CASING SEAL ASSEMBLY AND METHOD | 05-02-2013 |
| 20130104566 | TURBINE OF A TURBOMACHINE | 05-02-2013 |
| 20130133337 | HYDROGEN ASSISTED OXY-FUEL COMBUSTION - System and methods for hydrogen assisted oxy-fuel combustion are provided. The system includes a combustor, an air separation unit, a fuel stream source, and a condenser. The combustor includes an oxygen input port, a fuel stream input port, a carbon dioxide input port, and an exhaust output port. The air separation unit is in fluid communication with the combustor via the oxygen input port of the combustor. The fuel stream source is in fluid communication with the combustor via the fuel stream input port and includes a fuel source and a hydrogen source. The condenser is disposed to receive an exhaust from the combustor via the exhaust output port and to return an output stream to the combustor via the carbon dioxide input port. The method includes combusting a fuel stream in a combustor in the presence of an oxidizer to generate an exhaust gas. | 05-30-2013 |
| 20130139523 | FULL HOOP CASING FOR MIDFRAME OF INDUSTRIAL GAS TURBINE ENGINE - A can annular industrial gas turbine engine, including: a single-piece rotor shaft spanning a compressor section ( | 06-06-2013 |
| 20130160466 | Nickel Based Forged Alloy, Gas Turbine Member Using Said Alloy and Gas Turbine Using Said Member - It is an objective of the invention to provide an Ni-based forged alloy having good large ingot formability and good hot formability as well as high mechanical strength at high temperature. There is provided an Ni-based forged alloy comprising: 0.001 to 0.1 mass % of C; 0.001 to 0.01 mass % of B; 16 to 22 mass % of Cr; 0.5 to 1.5 mass % of Al; 0.1 to 6.0 mass % of W; 3.5 to 5.5 mass % of Nb; 0.8 to 3.0 mass % of Ti; 16 to 20 mass % of Fe; 2.0 mass % or less of Mo; and the balance including Ni and unavoidable impurities, in which: a segregation parameter Ps defined by a formula of “Ps (mass %)=1.05[Al concentration (mass %)]+0.6[Ti concentration (mass %)]−0.8[Nb concentration (mass %)]−0.3[Mo concentration (mass %)]” satisfies a relationship of “Ps≧−3.0 mass %”; and total amount of W and Mo is 1.75 atomic % or less. | 06-27-2013 |
| 20130167555 | ALUMINUM FAN BLADE CONSTRUCTION WITH WELDED COVER - An airfoil includes, among other possible things, a main body extending between a leading edge and a trailing edge. Channels are formed into the main body, with a plurality of ribs extending intermediate the channels. A cover skin is attached to the main body. The cover skin is welded to the main body at outer edges. An adhesive is placed between inner surfaces of the cover skin and the main body. The adhesive is deposited inwardly of the outer edges of the cover skin. | 07-04-2013 |
| 20130186106 | PRECIPITATION HARDENING MARTENSITIC STAINLESS STEEL, AND STEAM TURBINE LONG BLADE, STEAM TURBINE, AND POWER PLANT USING THE SAME - The problem to be solved of the present invention is to provide a precipitation hardening martensitic stainless steel having excellent tissue stability, strength, toughness, and corrosion-resistance, requiring no sub-zero treatment, and having excellent productivity; and also a steam turbine long blade using the same. The problem is solved by providing a precipitation hardening martensitic stainless steel containing, by mass, 0.1% or less of C; 0.1% or less of N; 9.0% or more and 14.0% or less of Cr; 9.0% or more and 14.0% or less of Ni; 0.5% or more and 2.5% or less of Mo; 0.5% or less of Si; 1.0% or less of Mn; 0.25% or more and 1.75% or less of Ti; 0.25% or more and 1.75% or less of Al, and the rest is Fe and inevitable impurities; and a steam turbine long blade using the precipitation hardening martensitic stainless steel. | 07-25-2013 |
| 20130192260 | GAS TURBINE ENGINE SEAL CARRIER - A gas turbine engine includes a seal assembly that is supported by a member at a joint. The seal assembly includes a seal support having a radial flange secured to the joint. A first bend adjoins the radial flange to a first leg, which is oriented generally in an axial direction. A second bend adjoins the first leg to a second leg, which is conical in shape. A seal is supported by the second leg. | 08-01-2013 |
| 20130192261 | GAS TURBINE ENGINE MID TURBINE FRAME BEARING SUPPORT - A gas turbine engine includes high and low pressure turbines. A mid turbine frame is arranged axially between the high and low pressure turbines. A bearing is operatively supported by a support structure. An inner case is secured to the support structure and includes a first conical member and a bearing support to which the bearing is mounted. The bearing support includes a second conical member that is secured to the first conical member at a joint. The first and second conical members are arranged radially inward of the joint. | 08-01-2013 |
| 20130192262 | ANNULAR COMBUSTOR - An annular combustor includes an annular outer shell that includes a first flange that defines an inner diameter ID | 08-01-2013 |
| 20130192263 | GAS TURBINE ENGINE WITH HIGH SPEED LOW PRESSURE TURBINE SECTION - A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5. | 08-01-2013 |
| 20130192264 | TURBINE ENGINE GEARBOX - An example method of controlling performance of gearbox of a gas turbine engine includes establishing a gear characteristic of a plurality of double helical gears each disposed about a respective axis in a gearbox. Performance of the plurality of double helical gears is controlled by selecting a circumferential offset distance between a first plurality of gear teeth spaced apart from a second plurality of gear teeth on each of the plurality of double helical gears in response to the established gear characteristic. | 08-01-2013 |
| 20130192265 | GAS TURBINE ENGINE WITH HIGH SPEED LOW PRESSURE TURBINE SECTION AND BEARING SUPPORT FEATURES - A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5. The high pressure turbine is mounted on the low pressure turbine with an intermediate bearing. | 08-01-2013 |
| 20130192266 | GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE - A gas turbine engine includes a fan rotatable about an axis, a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section includes a fan drive turbine and a second turbine. The second turbine is disposed forward of the fan drive turbine. The fan drive turbine includes at least three rotors and at least one rotor having a bore radius (R) and a live rim radius (r), and a ratio of r/R is between about 2.00 and about 2.30. A speed change system is driven by the fan drive turbine for rotating the fan about the axis. | 08-01-2013 |
| 20130199207 | GAS TURBINE SYSTEM - A gas turbine system including a compressor section, a combustor section and a turbine section, wherein the compressor section includes a first wheel and a second wheel having a center bore extending axially therethrough, and wherein the first wheel and the second wheel are relatively adjacent each other. Also included is a gap disposed between the first wheel and the second wheel, wherein airflow is directed radially inward within the gap toward the center bore of the second wheel with or without the help of impellers extending radially. The compressor section further includes an airflow manipulation device disposed within the gap, comprising of at least one or multiple vanes which may start in radial direction but extend at least partially into the center bore. | 08-08-2013 |
| 20130205800 | VANE ASSEMBLIES FOR GAS TURBINE ENGINES - Vane assemblies for gas turbine engines and methods for assembling vane assemblies are disclosed. The vane assemblies may include at least one shroud having at least one vane-receiving portion, at least one vane having at least one end portion received in the vane-receiving portion, and at least one sealing member having an uncompressed cross-section that is substantially circular. The sealing member(s) are disposed between and in contact with the end portion of the vane and the vane-receiving portion of the shroud. | 08-15-2013 |
| 20130213057 | Gas Turbine Engines and Related Systems Involving Blade Outer Air Seals - Gas turbine engines and related systems involving blade outer air seals are provided. In this regard, a representative blade outer air seal segment for a set of rotatable blades includes: a blade arrival end; and a blade departure end; each of the blade arrival end and the blade departure end being angularly offset with respect to a longitudinal axis about which the blades rotate. | 08-22-2013 |
| 20130219922 | GEARED GAS TURBINE ENGINE WITH REDUCED FAN NOISE - A fan section for a gas turbine engine has a fan rotor with a plurality of fan blades. A plurality of exit guide vanes are positioned to be downstream of the fan rotor. The fan rotor is driven through a gear reduction relative to a turbine section. The exit guide vanes are desired to address resultant sound from interaction of wakes from the fan blades across exit guide vanes. A gas turbine engine incorporating a fan section is also disclosed. | 08-29-2013 |
| 20130239585 | TANGENTIAL FLOW DUCT WITH FULL ANNULAR EXIT COMPONENT - An arrangement ( | 09-19-2013 |
| 20130239586 | FAN BLADE AND METHOD OF MANUFACTURING SAME - An airfoil for a gas turbine engine includes a substrate and a sheath providing an edge. A cured adhesive secures the sheath to the substrate. The cured adhesive has a fillet that extends beyond the edge that includes a mechanically worked finished surface. A method of manufacturing the airfoil includes the steps of securing a sheath to a substrate with adhesive, curing the adhesive, and mechanically removing a portion of the adhesive extending beyond the sheath. | 09-19-2013 |
| 20130239587 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, a gear arrangement configured to drive the fan section, a compressor section, including both a low pressure compressor section and a high pressure compressor section. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio provided by the combination of a pressure ratio across the low pressure compressor section and a pressure ratio across the high pressure compressor section is greater than about 35. The pressure ratio across a first of the low and high pressure compressor sections is between about 3 and about 8. The pressure ratio across a second of the low and high pressure compressor sections is between about 7 and about 15. The fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct. | 09-19-2013 |
| 20130247586 | Blade Wedge Attachment - A rotor includes a disk that has slots circumferentially arranged around its periphery. Blades include respective roots that are mounted in respective ones of the slots. The roots are smaller than the slots such that there are circumferential gaps between the roots and circumferential sides of the slots. Wedges are respectively located within the circumferential gaps. The wedges are free floating with regard to the blades and the disk. | 09-26-2013 |
| 20130255277 | GAS TURBINE ENGINE NOSE CONE - A nose cone for a turbofan gas turbine engine includes a central tip, an outer perimeter and a substantially conical outer wall extending therebetween which encloses a cavity therewithin. The outer wall includes an inner substrate layer facing the cavity and an outer layer which overlies and at least partially encloses the inner substrate layer. The outer layer is composed entirely of a nanocrystalline metal forming an outer surface of the nose cone. | 10-03-2013 |
| 20130276455 | AIRFOIL WITH BREAK-WAY, FREE-FLOATING DAMPER MEMBER - An airfoil includes an airfoil body that has a leading edge and a trailing edge and a first sidewall and a second sidewall that is spaced apart from the first sidewall. The first sidewall and the second sidewall join the leading edge and the trailing edge and at least partially define a cavity in the airfoil body. A damper member is enclosed in the cavity and is free-floating within the cavity. | 10-24-2013 |
| 20130276456 | AIRFOIL INCLUDING MEMBER CONNECTED BY ARTICULATED JOINT - An airfoil includes a body that has a platform, an airfoil extending outwardly from a side of the platform and a root extending outwardly from another side of the platform. A member is connected in an articulated joint to the body. | 10-24-2013 |
| 20130276457 | AIRFOIL INCLUDING LOOSE DAMPER - An airfoil includes an airfoil body that has a leading edge and a trailing edge and a first sidewall and a second sidewall that is spaced apart from the first sidewall. The first sidewall and the second sidewall join the leading edge and the trailing edge and at least partially define a cavity in the airfoil body. A damper member is enclosed in the cavity and is loose within the cavity. | 10-24-2013 |
| 20130276458 | GAS TURBINE INLET SYSTEM AND METHOD - A gas turbomachine inlet system includes a duct member having an inlet portion fluidically coupled to an outlet portion through an intermediate portion. The inlet portion, outlet portion, and intermediate portion define a fluid flow zone. A throttling system is arranged in the duct member at one of the inlet portion, outlet portion and intermediate portion. The throttling system is configured and disposed to selectively establish a pressure drop through the fluid flow zone. A method of controlling inlet pressure drop through an inlet system for a gas turbomachine is also described herein. | 10-24-2013 |
| 20130276459 | RESISTIVE BAND FOR TURBOMACHINE BLADE - A turbomachine system includes a rotor that defines a longitudinal axis of the turbomachine system. A first blade is coupled to the rotor, and the first blade has first and second laminated plies. A first band is coupled to the first blade and is configured to resist separation of the first and second laminated plies. | 10-24-2013 |
| 20130283820 | HOLLOW FAN BLADED WITH BRAIDED FABRIC TUBES - A fan blade having a body with a dovetail and an airfoil extending radially outwardly. The airfoil includes a pair of skins spaced to form an internal core, which define a pressure side and a suction side, and extending from a radially inner end to a radially outer tip. The core receives a plurality of braided tubes, with the tubes extending with at least a component in a radially outward direction. A fan and an engine are also described. | 10-31-2013 |
| 20130283821 | GAS TURBINE ENGINE AND NACELLE NOISE ATTENUATION STRUCTURE - A nacelle structure for a gas turbine engine assembly includes a fan case, an inlet, and a noise attenuation device. The fan case is configured to be disposed about a fan section of the gas turbine engine, which fan section has a diameter D. The inlet is attached to the fan case and extends axially forward of the fan case. A hilite of the inlet is spaced a distance L from a region in which the fan section is configured to be disposed. A ratio L/D is less than about 0.6. The noise attenuation structure covers a portion of an inner surface of the fan case and the inlet. | 10-31-2013 |
| 20130319004 | SHIELD SLOT ON SIDE OF LOAD SLOT IN GAS TURBINE ENGINE ROTOR - A rotor body rotates about an axis of rotation. A ledge provides a holding structure for holding blades. A plurality of blades are positioned beneath the ledge. load slot is sized to allow a mount portion of the blades to be moved radially inwardly of the ledge. The blades are moved circumferentially to have the mounted structure radially inwardly of the ledge. A lock slot is positioned on one circumferential side of the load slot. The lock slot is formed to receive a lock, and the lock is partially received within a portion of at least one of the blades, to lock the blades within the rotor, and a shield slot on a second circumferential side of the load slot. The shield slot is sized to be different from the lock slot such that a lock cannot be inadvertently positioned within the shield slot. | 12-05-2013 |
| 20130319005 | FLOATING SEGMENTED SEAL - A gas turbine engine rotor section includes a rotor body with a ledge extending axially from a location on the rotor body. The ledge defines a radially inner surface radially inwardly of the ledge, and a hub extends axially from the rotor, and beyond the ledge. The hub has a radially outer surface spaced from the ledge radially inner surface. A first distance is defined between the radially inner surface of the ledge and the radially outer surface of the hub. A knife edge seal has at least one pointed knife seal portion at a radially outer end. A radially inwardly extending arm portion extends from the seal, and an axially inwardly extending portion extends axially inwardly from the radially inwardly extending portion. The axially inwardly extending portion has a radially outer face and a radially inner face, which are spaced by a second distance. The second distance is less than the first distance. The axially inwardly extending portion is received between the radially inner face of the rotor and the radially outer face of the hub, such that the knife edge seal is free floating between the ledge and the hub. | 12-05-2013 |
| 20130319006 | DIRECT FEED AUXILIARY OIL SYSTEM FOR GEARED TURBOFAN ENGINE - A lubrication system for a fan drive gear system includes a main lubrication system and an auxiliary lubrication system. The auxiliary lubrication system including a collection channel disposed about the fan drive gear system for collecting expelled lubricant and an auxiliary pump including an inlet receiving lubricant from the collection channel and an outlet in communication with an auxiliary passage. The auxiliary pump supplies lubricant through the auxiliary passages to a bearing passage for communicating lubricant to a bearing. The auxiliary passages include a reservoir after the outlet of the auxiliary pump and before the bearing passage for storing lubricant. | 12-05-2013 |
| 20130319007 | LINER HANGER CABLE - A liner for a gas turbine engine includes a liner defining an inner surface exposed to exhaust gases and a duct spaced radially outward of the liner. A plurality of hanger assemblies is disposed within the radial space between the liner and the duct for supporting the liner relative to the duct. Each of the hanger assemblies includes a cable having a first end attached to the duct and a second end attached to the liner. | 12-05-2013 |
| 20130319008 | TURBINE BLADE SUPPORT - A turbine blade is disclosed. The turbine blade includes a platform, an airfoil extending from one side of the platform, a root extending radially from another side of the platform, and a pocket located beneath the platform. The pocket is defined by a plurality of walls, and a pad is disposed in a corner of the pocket. The pad includes three pad corners and three sides connecting the three pad corners, wherein each side extends along a different one of the plurality of walls. | 12-05-2013 |
| 20130319009 | PROTECTING OPERATING MARGIN OF A GAS TURBINE ENGINE - A method of protecting operating margin of the gas turbine engine includes calculating an aerodynamic distortion of air entering an inlet of a gas turbine engine that has a compressor section with variable vanes that are movable subject to a control parameter. The control parameter is selectively modified in response to the aerodynamic distortion. | 12-05-2013 |
| 20130319010 | AIRFOIL COVER SYSTEM - An example airfoil for a gas turbine engine includes a body having a first surface extending from a first edge to a second edge and a cavity disposed in the body. A first cover is at least partially disposed within the cavity. The first cover includes a first portion cooperates with a corresponding second portion. A second cover covers the first cover and forms at least a portion of the first surface with the body. The first cover is disposed between the body and the second cover. The first cover and the second cover have a different coefficient of thermal expansion than the body. | 12-05-2013 |
| 20130319011 | GEARED ARCHITECTURE CARRIER TORQUE FRAME ASSEMBLY - A fan drive gear system for a gas turbine engine includes a carrier supporting circumferentially arranged gears within gear mount sections spaced circumferentially about a periphery. A shelf is disposed between each of the gear mounting sections. A torque frame is attached to the carrier and includes circumferentially arranged finger sections. Each of the finger sections includes first and second ends that are spaced apart to receive the shelf. A pin extends through openings in the shelf and first and second ends to secure the torque frame to the carrier. | 12-05-2013 |
| 20130327060 | SINGLE TURBINE DRIVING DUAL COMPRESSORS - A gas turbine engine is provided, which includes a first compressor rotor coupled to a first shaft, a second compressor rotor downstream of the first compressor and coupled to a second shaft. A combustor is disposed downstream of the compressor rotors. A turbine is disposed downstream of the combustor and coupled to the second shaft. A gear is configured to be driven by the second shaft. The second shaft is configured for driving the first shaft through the gear, whereby the turbine drives the first compressor rotor at a different rotational speed than the turbine. | 12-12-2013 |
| 20130333392 | TURBINE COMPRESSOR BLADE TIP RESISTANT TO METAL TRANSFER - A gas turbine engine having an engine casing extending circumferentially about an engine centerline axis; and a compressor section, a combustor section, and a turbine section within said engine casing. At least one of said compressor section and said turbine section includes at least one airfoil and at least one seal member adjacent to the at least one airfoil, wherein a tip of the at least one airfoil is metal having a thin film ceramic coating and the at least one seal member is coated with an abrasive. | 12-19-2013 |
| 20140000280 | GAS TURBINE ENGINE TURBINE BLADE AIRFOIL PROFILE | 01-02-2014 |
| 20140000281 | GAS TURBINE ENGINE TURBINE BLADE AIRFOIL PROFILE | 01-02-2014 |
| 20140007591 | ADVANCED TIP-TIMING MEASUREMENT BLADE MODE IDENTIFICATION - A disclosed airfoil health monitoring system and method obtains a signal comprising a waveform indicative of an airfoil path with a sensor. Features of the waveform are determined and compared waveform characteristics indicative of a vibrational mode. A vibrational mode of the airfoil may then be determined based on the comparison between the predetermined waveform characteristics and the obtained waveform indicative of the airfoil path. | 01-09-2014 |
| 20140013772 | JOINT BETWEEN AIRFOIL AND SHROUD - A stator joint for a gas turbine engine has a center axis, and a shroud having a radial wall facing substantially radially with respect to the center axis. A slot wall defines in-part a slot in the shroud. A relief wall defines a relief area of the slot. The relief wall extends between the radial wall and the slot wall. A vane has an airfoil and a lug extending into the slot. A flowable attachment material is disposed in the relief area for engagement of the vane to the shroud. A vane assembly and a gas turbine engine are also disclosed. | 01-16-2014 |
| 20140020403 | SEALING DEVICE, AXIAL TURBINE AND POWER PLANT - In one embodiment, a sealing device includes seal fins provided on an inner circumferential surface of a stationary body or an outer circumferential surface of a rotating body so as to be adjacent to each other in an axial direction of the rotating body in a gap between the outer circumferential surface of the rotating body and the inner circumferential surface of the stationary body. The device further includes at least one opening member provided on the inner circumferential surface of the stationary body, the opening member being provided at a position between seal fins adjacent to each other in the axial direction, and having holes opened on a side of the inner circumferential surface of the stationary body. | 01-23-2014 |
| 20140020404 | FUNDAMENTAL GEAR SYSTEM ARCHITECTURE - A fan drive gear system for a gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan and a mount flexibly supporting portions of the gear system. A lubrication system supporting the fan drive gear system provides lubricant to the gear system and removes thermal energy produced by the gear system. The lubrication system includes a capacity for removing energy equal to less than about 2% of energy input into the gear system. | 01-23-2014 |
| 20140041395 | GAS TURBINE - A gas turbine burns the air compressed in a compressor with supplying fuel in a combustor so as to obtain rotary power by supplying the generated combustion gas to a turbine. The turbine includes turbine vane elements and turbine blade elements that are alternately positioned in a direction in which the combustion gas fluidizes in a turbine cylinder having a cylindrical shape, and a flue gas diffuser having a cylindrical shape and connected to a rear portion of the turbine cylinder. The turbine blade element includes a plurality of turbine blades positioned at equal intervals in the circumference direction. The turbine blades have a throat width on a longitudinal end side made larger than a throat width on a longitudinally intermediate portion side. This efficiently restores the pressure of the flue gas. This improves the efficiency of the turbine so that the performance can be improved. | 02-13-2014 |
| 20140060080 | GAS TURBINE ENGINE AFT BEARING ARRANGEMENT - An example gas turbine engine includes a turbine and first and second spools coaxial with one another. The first spool is arranged within the second spool and extends between forward and aft ends. The aft end extends axially beyond the second spool and supports the turbine. A housing is arranged downstream from the turbine. First and second bearings are mounted to the aft end of the first spool and supported by the housing portion. | 03-06-2014 |
| 20140060081 | SINGLET VANE CLUSTER ASSEMBLY - A vane cluster for a gas turbine engine includes multiple singlet vanes and a forward wear liner connecting a forward edge of each singlet vane, thereby allowing the vane cluster to be manipulated as a single component. | 03-06-2014 |
| 20140060082 | COMBUSTOR SHELL AIR RECIRCULATION SYSTEM IN A GAS TURBINE ENGINE - A shell air recirculation system for use in a gas turbine engine includes one or more outlet ports located at a bottom wall section of an engine casing wall and one or more inlet ports located at a top wall section of the engine casing wall. The system further includes a piping system that provides fluid communication between the outlet port(s) and the inlet port(s), a blower for extracting air from a combustor shell through the outlet port(s) and for conveying the extracted air to the inlet port(s), and a valve system for selectively allowing and preventing air from passing through the piping system. The system operates during less than full load operation of the engine to circulate air within the combustor shell but is not operational during full load operation of the engine. | 03-06-2014 |
| 20140060083 | GAS TURBINE ENGINE AFT BEARING ARRANGEMENT - An example gas turbine engine includes a turbine and first and second spools coaxial with one another. The first spool is arranged within the second spool and extends between forward and aft ends. The aft end extends axially beyond the second spool and supports the turbine. A housing is arranged downstream from the turbine. First and second bearings are mounted to the aft end of the first spool and supported by the housing portion. | 03-06-2014 |
| 20140069108 | BUCKET ASSEMBLY FOR TURBOMACHINE - Bucket assemblies are provided. The bucket assembly includes a shank, and an airfoil positioned radially outward of the shank. The bucket assembly further includes a main cooling circuit defined in the airfoil and the shank, the main cooling circuit comprising seven passages, each of the seven passages fluidly connected with an adjacent one of the seven passages. A maximum rotation number in each of the seven passages is less than or equal to approximately 0.4. | 03-13-2014 |
| 20140069109 | ELECTRICAL GROUNDING FOR BLADES - A rotor in a gas turbine engine has a rotor body with at least one slot receiving a blade. The blade has an outer surface formed of a first material and an airfoil extending from a dovetail. The dovetail is received in the slot. A grounding element is in contact with a portion of the dovetail formed of a second material that is more electrically conductive than the first material. The grounding element is in contact with a rotating element that rotates with the rotor and is formed of a third material. The first material is less electrically conductive than the third material. The grounding element and rotating element together form a ground path from the portion of the dovetail into the rotor. | 03-13-2014 |
| 20140069110 | TURBINE BUCKET INTERNAL CORE PROFILE - Turbine bucket nominal internal core profiles and core insert external profiles are provided. In one embodiment, a turbine bucket includes an airfoil, platform, shank and dovetail. The bucket has a nominal internal core profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table 1 wherein the Z values are non-dimensional values from 0 to 1 convertible to Z distances in inches by multiplying the Z values by a height of the bucket in inches, and wherein X and Y are non-dimensional values which, when connected by smooth continuing arcs, define internal core profile sections at each distance Z along the bucket, the profile sections at the Z distances being joined smoothly with one another to form said bucket internal core profile. | 03-13-2014 |
| 20140083113 | FLOW CONTROL TAB FOR TURBINE SECTION FLOW CAVITY - Flow control tabs for turbine sections are provided. The turbine section includes a stator assembly and a rotor assembly. The rotor assembly includes a bucket, the bucket having a shank. The rotor assembly and the stator assembly are spaced apart along a longitudinal axis and defining a flow cavity therebetween such that a cooling fluid flows therein. The flow control tab includes a mount surface configured for mounting to the stator assembly, a radially outer surface, a radially inner surface, and a leading edge. The leading edge connects the outer surface and the inner surface such that cooling fluid interacting with the leading edge is divided into an outer purge flow and an inner recirculation flow. | 03-27-2014 |
| 20140090401 | Endwall Controuring - An airfoil array is disclosed. The airfoil array may include an endwall, and a plurality of airfoils radially projecting from the endwall. Each airfoil may have a first side and an opposite second side extending axially in chord between a leading edge and a trailing edge. The airfoil array may further include a convex profiled region extending from the endwall adjacent the first side of at least one of said plurality of airfoils and near the leading edge of the at least one of said plurality of airfoils. The airfoil array may further include a concave profiled region in the endwall adjacent the first side of said at least one of said plurality of airfoils and aft of the convex profiled region. | 04-03-2014 |
| 20140102115 | TURBO-ENGINE, PARTICULARLY INTERNAL COMBUSTION ENGINE - The invention is directed to a turbo-engine, particularly internal combustion engine, comprising a housing and therein a bladeless turbine section ( | 04-17-2014 |
| 20140116065 | TURBINE NOZZLE GUIDE IN A TURBINE ENGING - A sectorized nozzle for a turbine engine turbine including an inner sectorized annular platform and an outer sectorized annular platform connected together by radial airfoils, at least one of the platforms including a plurality of orifices for passing air in a neighborhood of its upstream end, the orifices being distributed over the circumference of the platform and opening out at their ends remote from the airfoils into a circumferential annular cavity of the sector of the platform, which cavity is closed by a metal sheet fastened to the platform sector and pierced by orifices for feeding cooling air. | 05-01-2014 |
| 20140123680 | MODULAR DROP-IN COMBUSTOR ASSEMBLY FOR INDUSTIAL GAS TURBINE AND METHOD FOR INSTALLATION - A preassembled modular drop-in combustor having internal components in conformity with assembly and function specifications prior to and after insertion into an industrial gas turbine access port and internal combustor transition. The combustor assembly maintains conformity with those specifications after insertion into the combustor case if it does not inadvertently impact other turbine components during its installation. Inadvertent impact is avoided by having a combustor service zone proximal the combustor case, enabling slidable insertion of each combustor assembly into its corresponding access port and transition along its corresponding insertion path without contacting other turbine system components. A multi-axis motion combustor handling tool in the combustor service zone, preferably under automatic control, is coupled to each combustor and facilitates precise alignment along the insertion path. Automatic control facilitates consistent repetitive combustor installation and removal by executing a sequence of stored pre-determined manipulation steps. | 05-08-2014 |
| 20140130513 | SYSTEM AND METHOD FOR IMPROVING GAS TURBINE PERFORMANCE AT PART-LOAD OPERATION - A compressor section of a gas turbine generally includes a stage of inlet guide vanes positioned adjacent to an inlet of the compressor section and a stage of rotor blades disposed downstream from the inlet guide vanes. A stage of stator vanes is positioned downstream from the stage of rotor blades. The stage of stator blades includes a row of leading guide vanes having a trailing edge. A row of trailing guide vanes coupled to an actuator is disposed between two corresponding adjacent leading guide vanes. Each of the trailing guide vanes includes a trailing edge. The leading edge of each trailing guide vane is disposed upstream of the trailing edge of a corresponding leading guide vane when the trailing guide vane is in an open position and downstream from the trailing edge of the corresponding leading guide vane when the trailing guide vane is in a closed position. | 05-15-2014 |
| 20140144157 | DOVETAIL ATTACHMENT SEAL FOR A TURBOMACHINE - A dovetail attachment seal for a turbomachine includes an outer seal member having a first end that extends to a second end through an intermediate portion, and at least one articulating element encapsulated, at least in part, by the outer seal member at one of the first and second ends. The at least one articulating element is slidingly disposed within the outer seal member. | 05-29-2014 |
| 20140144158 | TURBOMACHINE COMPONENT INCLUDING A SEAL MEMBER - A turbomachine component includes a first housing portion having a first end that extends to a second end. At least one of the first and second ends includes a first sealing surface having a seal receiving passage. A second housing portion includes a first end that extends to a second end. At least one of the first and second ends includes a second sealing surface configured and disposed to register with the first sealing surface. A seal member is provided in the seal receiving passage. The seal member includes a plurality of bristles that extend toward the second sealing surface. | 05-29-2014 |
| 20140150454 | TURBINE BLADE APPARATUS - A turbine blade is disclosed. The turbine blade may include a platform, an airfoil extending from one side of the platform, and a neck extending from another side of the platform, wherein the neck includes a forward buttress and an aft buttress. The turbine blade may further include a root extending from the neck, a pocket defined by a plurality of walls and located between the forward. buttress and the aft buttress, and a variable radius fillet. The variable radius fillet may be disposed within the pocket and extend between the forward buttress and the aft buttress, wherein a radius of the variable radius fillet increases from the forward buttress to the aft buttress. | 06-05-2014 |
| 20140165591 | TURBINE UNDER PLATFORM AIR SEAL STRIP - A gas turbine engine stage includes a rotor having a slot. A blade has a root received in the slot, and a shank that extends radially outward from the root to a platform that supports an airfoil. A seal is supported on the shank and in engagement circumferentially between and with the shank and the rotor within the slot. A gas turbine engine includes compressor and turbine sections. The turbine section has a rotor with a slot. A combustor is provided axially between the compressor and turbine sections. A turbine blade in the turbine section includes a root received in the slot and a shank that extends radially outward from the root to a platform that supports an airfoil. | 06-19-2014 |
| 20140165592 | COMPRESSOR BLADE FOR GAS TURBINE ENGINE - A compressor blade for a gas turbine engine is disclosed. The compressor blade may have a root configured to engage a hub of the gas turbine engine, and an airfoil radially extending a distance from the root to a tip. The airfoil may have a suction side, a pressure side, a leading edge connecting the suction and pressure sides, and a trailing edge connecting the suction and pressure sides opposite the leading edge. The distance that the airfoil extends from the root to the tip may be divided into a plurality of radially adjacent regions. At least one, but not all, of the plurality of radially adjacent regions away from the base and the tip may have a substantially constant thickness. | 06-19-2014 |
| 20140174098 | TURBINE DISC WITH REDUCED NECK STRESS CONCENTRATION - A disc with two sides includes a hub having a bore and a bore radius, a neck, and a rim. The neck is connected to and radially outward of the hub and has an inner wedge with a curved section on one side of the disc, an outer wedge with a curved section on that same side of the disc, and a center section between the wedges with a flat side on that same side of the disc. The rim is connected to and radially outward of the neck, the rim having a radius that is no more than seven times greater than the bore radius. | 06-26-2014 |
| 20140182309 | GEARED GAS TURBINE ENGINE EXHAUST NOZZLE WITH CHEVRONS - A gas turbine engine includes a core engine having a compressor section fluidly connected to a combustor section that is also fluidly connected to a turbine section. A core nacelle surrounds the core engine and includes a core nozzle. A fan is connected to the compressor section and is arranged upstream from the core engine. A gear train interconnects the turbine section to the fan. A fan nacelle at least partially surrounds the core nacelle and includes a fan nozzle. The fan is disposed in the fan nacelle. At least one of the fan nozzle and core nozzle includes circumferentially fixed chevrons that provide a fixed exit area throughout engine operation. | 07-03-2014 |
| 20140190181 | ROTOR COVER PLATE - A cover plate for a rotor disk in a gas turbine machine includes a cylindrical body having multiple outward facing snaps and multiple inward facing snaps. | 07-10-2014 |
| 20140196472 | GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE - A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine. | 07-17-2014 |
| 20140196473 | GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE - A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine. | 07-17-2014 |
| 20140202170 | Composite Articles and Methods - An article has a polymeric substrate and a coating system. The coating system includes a metallic plating and a polymeric coating atop the metallic plating. The metallic plating has a thickness of at least 0.05 mm. | 07-24-2014 |
| 20140238043 | GAS TURBINE ENGINE SYSTEM AND SUPERSONIC EXHAUST NOZZLE - One embodiment of the present invention is a unique gas turbine engine system. Another embodiment is a unique exhaust nozzle system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engine systems and exhaust nozzle systems for gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. | 08-28-2014 |
| 20140245752 | SYSTEM AND METHOD FOR ATTACHING A ROTATING BLADE IN A TURBINE - A system for attaching a rotating blade in a turbine includes a bush having an axial slot and a radial slot that intersects with the axial slot. A radial retention member fits within the radial slot, and an axial retention member fits within the axial slot and engages with the radial retention member. A method for attaching a rotating blade in a turbine includes inserting a bush into an axial passage in a rotor wheel and inserting a radial retention member into a radial passage in the rotor wheel and through at least a portion of the bush. The method further includes inserting the rotating blade in a slot in the rotor wheel, inserting the radial retention member into a retention slot in the rotating blade, and inserting an axial retention member into the bush. | 09-04-2014 |
| 20140245753 | GAS TURBINE ENGINE ROTOR BLADE - A rotor blade for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil extending in span between a root region and a tip region and a tip portion extending at an angle from the tip region of the airfoil. | 09-04-2014 |
| 20140260324 | TURBO-MACHINERY ROTORS WITH ROUNDED TIP EDGE - A rotor for a gas turbine engine includes a plurality of radially extending blades, each having a remote blade tip defining an outer tip surface, and a leading edge defined between opposed pressure and suction side airfoil surfaces. A shroud circumferentially surrounds the rotor, and a radial distance between an inner surface of the shroud and the outer tip surface of the blades defines a radial tip clearance gap therebetween. The tip of each of the blades has a pressure side edge formed at the intersection between the outer tip surface and the pressure side airfoil surface, and a suction side edge formed at the intersection between the outer tip surface and the pressure side airfoil surface. The suction side edge has a larger radius of curvature than the pressure side edge, thereby reducing the amount of tip leakage flow through the radial tip clearance gap. | 09-18-2014 |
| 20140260325 | GAS TURBINE ENGINE EXHAUST FLUID PASSAGE DUCT - A fluid plenum including a body defining an internal cavity having an inlet and an outlet. The fluid plenum further includes at least one wall positioned in the internal cavity that divides the internal cavity into first and second passageways, and which also divides the inlet into first and second inlet portions, and divides the outlet into first and second outlet portions. The first passageway receives fluid through the first inlet portion and directs fluid to the first outlet portion, and the second passageway receives fluid through the second inlet portion and directs fluid to the second outlet portion. The first and second passageways extend first and second lengths that are different from one another, and also generate a substantially common back-pressure at the first and second inlet portions during flow of a fluid stream through the inlet, including a first sub-stream of the fluid stream through the first passageway and a second sub-stream of the fluid stream through the second passageway. | 09-18-2014 |
| 20140260326 | GEARED TURBOFAN ENGINE WITH HIGH COMPRESSOR EXIT TEMPERATURE - A gas turbine engine includes a fan with a plurality of fan blades rotatable about an axis, and a compressor section that includes at least first and second compressor sections. An average exit temperature of the compressor section is between about 1000° F. and about 1500° F. The engine also includes a combustor that is in fluid communication with the compressor section, and a turbine section that is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis. | 09-18-2014 |
| 20140318151 | VOLUTE WITH TWO CHAMBERS FOR A GAS TURBINE | 10-30-2014 |
| 20140345295 | FLUID SUPERCHARGING DEVICE AND TURBINE ENGINE - A fluid supercharging device ( | 11-27-2014 |
| 20150020531 | GAS TURBINE-HEATED HIGH-TEMPERATURE BATTERY - A power plant system is provided having a high temperature battery, supplied with fluid via at least one supply line, for storing and releasing electrical energy, a gas turbine for generating electrical energy, and a heat exchanger which is designed to extract thermal energy from the exhaust stream of the gas turbine and transfer said thermal energy to the fluid, which fluid can be supplied after heat transfer to the high temperature battery via the at least one supply line. | 01-22-2015 |
| 20150027131 | Axial Compressor, Gas Turbine with Axial Compressor, and its Remodeling Method - An axial compressor includes: a rotor as a rotational shaft; a plurality of rotor blades mounted on the rotor; a compressor casing that covers the rotor and the rotor blades; a plurality of stator vanes mounted on the compressor casing. The rotor blades and the stator vanes are each disposed in a circumferential direction of the rotational shaft to form a rotor-blade cascade and a stator-vane cascade, respectively. The rotor-blade cascade and the stator-vane cascade are arranged in plural rows, respectively, in an axial direction of the rotational shaft. The stator vane has a dovetail as a base for supporting a vane section. The compressor casing has a dovetail groove formed therein. The dovetail groove extends in the circumferential direction of the rotational shaft to receive the dovetail inserted therein to fix the stator vanes. Two or more stator vanes, the stator vanes belonging to stator-vane cascades different from each other, are fixed in the dovetail groove. | 01-29-2015 |
| 20150075178 | GAS TURBINE ENGINES WITH TURBINE ROTOR BLADES HAVING IMPROVED PLATFORM EDGES - A turbine rotor blade is provided. The turbine rotor blade includes a root, a platform coupled to the root, and an airfoil extending from the platform. The platform has a leading edge, a trailing edge, a suction side edge, and a pressure side edge. The pressure side edge includes a first concave portion. | 03-19-2015 |
| 20150075179 | Systems and Methods for Modifying a Pressure Side on an Airfoil About a Trailing Edge - An airfoil is disclosed herein. The airfoil may include a leading edge, a trailing edge, a suction side defined between the leading edge and the trailing edge, and a pressure side defined between the leading edge and the trailing edge opposite the suction side. The pressure side may include a concave profile about the trailing edge that varies from a profile of a remainder of the pressure side. | 03-19-2015 |
| 20150089958 | GAS TURBINE ENGINE WITH HIGH SPEED LOW PRESSURE TURBINE SECTION - A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5. | 04-02-2015 |
| 20150089959 | GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION - A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supporting the shaft relative to the inlet case. The second bearing is arranged radially outward from the shaft. | 04-02-2015 |
| 20150096305 | METHOD AND SYSTEM FOR PROVIDING COOLING FOR TURBINE COMPONENTS - A method and system for providing cooling of a turbine component that includes a region to be cooled is provided. A recess is defined within the region to be cooled, and includes an inner face. At least one support projection extends from the inner face. The at least one support projection includes a free end. A cover is coupled to the region to be cooled, such that an inner surface of the cover is coupled to the free end of the at least one support projection, such that at least one cooling fluid passage is defined within the region to be cooled. | 04-09-2015 |
| 20150101346 | LOCKING SPACER ASSEMBLY - A locking spacer assembly for securing adjacent rotor blades includes a first end piece having a platform portion and a root portion that define a first inner surface of the first end piece. The root portion defines a first projection and an opposing second projection of the first end piece. The first projection has an outer profile adapted to project into a first lateral recess of the attachment slot. The second projection has an outer profile adapted to project into a second lateral recess of the attachment slot. A second end piece fits between the first inner surface of the first end piece and a sidewall portion of the attachment slot and includes a platform portion and a root portion. A borehole extends continuously through the first end piece and the second end piece. A fastener configured to engage with a sidewall portion of the attachment slot extends through the borehole. | 04-16-2015 |
| 20150101347 | LOCKING SPACER ASSEMBLY - A locking spacer assembly for securing adjacent rotor blades includes a first end piece having a platform portion and a root portion that define an angled first inner surface of the first end piece. The root portion defines a first projection adapted to project into a recess portion of the attachment slot. A second end piece fits between the first inner surface and a sidewall portion of the attachment slot and includes a platform portion and a root portion that define a second projection adapted to project into a recess portion of the attachment slot. The platform portion and the root portion define an angled second inner surface and that is configured to mate with the first inner surface. A borehole extends through the platform portion of the first end piece and the root portion of the second end piece and a fastener extends through the borehole. | 04-16-2015 |
| 20150101348 | LOCKING SPACER ASSEMBLY - Locking spacer assemblies, rotor assemblies and turbomachines are provided. In one embodiment, a locking spacer assembly includes a first end piece and a second end piece each configured to fit into a space between platforms of adjacent rotor blades, the first end piece and second end piece each comprising an outer surface and an inner surface, the outer surface having a profile adapted to project into an attachment slot, wherein the inner surfaces of the first and second end pieces generally face each other. The locking spacer assembly further includes an actuator movable between the inner surfaces, the actuator comprising a projection configured to engage the inner surface, the actuator further comprising a plurality of locating protrusions extending from the projection, the locating protrusions configured to fit within locating channels defined in the first end piece and the second end piece. | 04-16-2015 |
| 20150101349 | LOCKING SPACER ASSEMBLY - Locking spacer assemblies, rotor assemblies and turbomachines are provided. In one embodiment, a locking spacer assembly includes a first end piece and a second end piece each configured to fit into a space between platforms of adjacent rotor blades, the first end piece and second end piece each comprising an outer surface and an inner surface, the outer surface having a profile adapted to project into an attachment slot, wherein the inner surfaces of the first and second end pieces generally face each other. The locking spacer assembly further includes an actuator movable between the inner surfaces, the actuator comprising a projection, the projection comprising a first surface and a second surface formed on the projection and configured to engage the inner surfaces, the first and second surfaces generally perpendicular to radial. | 04-16-2015 |
| 20150101350 | LOCKING SPACER ASSEMBLY - A locking spacer assembly for securing adjacent rotor blades includes a first end piece configured to fit into a space between the platforms of adjacent rotor blades. The first end piece comprises an outer surface and an inner surface. The outer surface has a profile that is adapted to project into the attachment slot. A second end piece is configured to fit into a space between the platforms. The second end piece comprises an outer surface and an inner surface. The outer surface has a profile that is adapted to project into the attachment slot. The inner surfaces of the first and second end pieces generally face each other. The first end piece and the second end piece are bonded together either directly or via a spacer block that is inserted between the inner surfaces. | 04-16-2015 |
| 20150101351 | LOCKING SPACER ASSEMBLY - Locking spacer assemblies and turbomachines are provided. In one embodiment, a locking spacer assembly includes a spacer, the spacer including a platform and a plurality of legs extending generally radially inward from the platform. The locking spacer assembly further includes a clamp configured to contact and cause elastic deformation of each of the plurality of legs in a generally axial direction towards each other. The locking spacer assembly further includes a locking lug configured to contact and impart a force against each of the plurality of legs in an opposite generally axial direction. | 04-16-2015 |
| 20150107265 | TURBINE BUCKET WITH ENDWALL CONTOUR AND AIRFOIL PROFILE - Turbine frequency tuning, fluid dynamic efficiency, and performance can be improved using an airfoil profile and/or an endwall contour including at least one of a pressure side bump, a pressure side leading edge bump, or a suction side trough. In particular, by including two endwall bumps on the pressure side and a trough on the suction side combined with a particular airfoil profile, performance can be further improved. | 04-23-2015 |
| 20150107266 | TURBINE BUCKET PROFILE YIELDING IMPROVED THROAT - Turbine frequency tuning, fluid dynamic efficiency, and performance can be improved using a particular airfoil profile, which can be used to determine a throat between adjacent airfoils. By shaping the throat according to the particular profile, the total pressure at an endwall can be energized, improving performance of the turbine. | 04-23-2015 |
| 20150128611 | STACKED WHEEL ASSEMBLY FOR A ROTOR OF A ROTARY MACHINE - A stacked wheel assembly for a rotor of a rotary machine includes a plurality of stacked wheels for rotation about a common axis and forming a portion of the rotor. Also included is a tie bolt passing through aligned bolt holes of the plurality of stacked wheels for retaining the plurality of stacked wheels in axially stacked relation, the tie bolt extending out of a forward end of a forward wheel of the plurality of stacked wheels and out of an aft end of an aft wheel of the plurality of stacked wheels. Further included is a rotor component disposed adjacent the aft end of the aft wheel. Yet further included is a nut mounted within a forward face of the rotor component, the nut configured to be in threaded engagement with the tie bolt to exert a clamping force on the plurality of stacked wheels. | 05-14-2015 |
| 20150128612 | SYSTEMS AND METHODS FOR VARYING A THROAT AREA BETWEEN ADJACENT BUCKETS IN A TURBINE FOR IMPROVED PART LOAD PERFORMANCE - A gas or steam turbine is disclosed herein. The turbine may include a throat area formed between adjacent buckets. The turbine also may include a variable throat device associated with at least one of the adjacent buckets. The variable throat device may be configured to vary the throat area between the adjacent buckets for improved part load performance. | 05-14-2015 |
| 20150128613 | Tapered Thread and Gas Turbine - The object of the present invention is to reduce galling that occurs at a tapered thread according to generation of excessive pressure on a tip end side of the tapered thread. In order to attain the above-mentioned object, a tapered thread according to the present invention is characterized in that it is a tapered thread comprising a taper-shaped external thread threadedly engaged with a taper-shaped internal thread formed in a member that is to be fastened, a tapered thread center hole is formed in a bottom surface of the external thread, and a depth of the tapered thread center hole is not more than a half of an axial length of the external thread. | 05-14-2015 |
| 20150292347 | FORWARD STEP HONEYCOMB SEAL FOR TURBINE SHROUD - The present application provides a stage of a gas turbine engine. The stage may include a bucket extending radially about a longitudinal axis of the gas turbine engine, a shroud facing the bucket, the shroud including a fore end portion including a radially inner surface spaced a first distance from the longitudinal axis, and a forward step honeycomb seal positioned on the shroud downstream of the fore end portion and facing the bucket. The forward step honeycomb seal may include a first linear portion including a radially inner surface spaced a second distance from the longitudinal axis, and a forward step portion positioned adjacent to and downstream of the first linear portion, the forward step portion including a radially inner surface spaced a third distance from the longitudinal axis, wherein the second distance is greater than the third distance, and wherein the third distance is greater than the first distance. | 10-15-2015 |
| 20150292411 | ANGLED CORE GAS TURBINE ENGINE MOUNTING - A propulsion system for an aircraft includes first and second turbine engines mounted within a fuselage of the aircraft. The first turbine engine includes a first engine core that drives a first propulsor disposed about a first propulsor axis. The second turbine engine includes a second engine core and a second propulsor disposed about a second propulsor axis parallel to the first propulsor axis. The first engine core and the second engine core are mounted at an angle relative to corresponding ones of the first and second propulsor axes. | 10-15-2015 |
| 20150300248 | COMPRESSOR ARRANGEMENT AND TURBOSHAFT ENGINE WITH A COMPRESSOR ARRANGEMENT - A compressor arrangement with centrifugal compressors arranged in sequence and each centrifugal compressor includes a rotor and a drive shaft, such that air flows through an inlet and out an outlet to a subsequent one of the centrifugal compressors, wherein drive shafts at two or more of the centrifugal compressors are arranged at an angle of between 0 degrees and 90 degrees with respect to one another. | 10-22-2015 |
| 20150322819 | MULTI-STAGE RADIAL FLOW TURBINE - Various multi-stage radial turbine configurations that provide highly efficient momentum transfer between a fluid and the mechanical interface in both power producing and power consuming undertakings. | 11-12-2015 |
| 20150322854 | GAS TURBINE ENGINE OIL TANK - A gas turbine engine includes an engine static structure housing that includes a compressor section and a turbine section. A combustor section is arranged axially between the compressor section and the turbine section. A core nacelle encloses the engine static structure to provide a core compartment. An oil tank is arranged in the core compartment and is axially aligned with the compressor section. A heat exchanger is secured to the oil tank and arranged in the core compartment. | 11-12-2015 |
| 20150322855 | REVERSE FLOW GAS TURBINE ENGINE CORE - A gas turbine engine has a fan rotor for delivering air into a bypass duct and into a core airflow duct. Air in the core flow duct passes axially downstream from the fan and past a reverse core engine including a turbine section, a combustor section, and a compressor section. The core airflow duct reaches a turning duct which turns the airflow radially inwardly to communicate with an inlet for the compressor section. Air in the compressor section passes to the combustor section. Products of the combustion pass downstream across a turbine rotor. An exhaust turning duct communicates products of the combustion from a full cylindrical portion downstream of the turbine rotor through a plurality of circumferentially separated mixing lobe outlets to mix with the bypass air in the bypass duct. The bypass duct extends past the mixing lobe outlets, and is defined circumferentially intermediate the mixing lobe outlets. | 11-12-2015 |
| 20150337675 | ACTIVE CLEARANCE CONTROL FOR GAS TURBINE ENGINE - A gas turbine engine includes a fluid intake. A turbine section includes a turbine case. A firewall is located upstream of the turbine section. A conduit is configured to direct a fluid from the fluid intake to the turbine case. A valve is located on a first side of the firewall opposite from the turbine section and is configured to regulate a flow of the fluid through the conduit. | 11-26-2015 |
| 20150344127 | AEROELASTICALLY TAILORED PROPELLERS FOR NOISE REDUCTION AND IMPROVED EFFICIENCY IN A TURBOMACHINE - Aeroelastically tailored propellers for noise reduction and improved efficiency in a turbomachine are provided including one or more upstream blades and one or more downstream blades disposed downstream relative to the one or more upstream blades. Each of the one or more upstream blades and the one or more downstream blades are aeroelastically tailored such that the one or more downstream blades include a greater degree of effective clipping during a second condition than at a first condition. Each blade among the one or more upstream blades comprises one or more geometric parameters. Each blade among the one or more downstream blades comprises one or more geometric parameters. In addition, an open rotor aircraft gas turbine engine assembly including the aeroelastically tailored propellers and a method of decreasing noise and improving efficiency in a turbomachine are disclosed. | 12-03-2015 |
| 20150345296 | TURBINE BUCKET ASSEMBLY AND TURBINE SYSTEM - A turbine bucket assembly and turbine system are disclosed. The turbine bucket assembly includes a single-lobe joint having an integral platform, the joint having a first axial length; a segmented airfoil having a root segment extending radially outward from the platform and a tip segment coupled to the root segment, the tip segment having a second axial length, which is less than the first axial length; and a turbine wheel defining a receptacle with a geometry corresponding to the single-lobe joint and being coupled to the single-lobe joint. The tip segment includes a tip segment material, the root segment includes a root segment material, and the turbine wheel includes a turbine wheel material, the root segment material and the turbine wheel material having a lower heat resistance and a higher thermal expansion than the tip segment material. | 12-03-2015 |
| 20150345307 | TURBINE BUCKET ASSEMBLY AND TURBINE SYSTEM - A turbine bucket assembly and turbine system are disclosed. The assembly includes a multi-lobe joint having an integral platform, the joint having a first axial length; a segmented airfoil having a root segment extending radially outward from the platform and a tip segment coupled to the root segment, the tip segment having a second axial length less than the first axial length; and a turbine wheel defining a receptacle with a geometry corresponding to the multi-lobe joint and being coupled to the multi-lobe joint. A tip segment material, a root segment material, and a turbine wheel material are selected, such that the turbine wheel material and the root segment material have a lower heat resistance and a higher thermal expansion than the tip segment material. | 12-03-2015 |
| 20150345309 | TURBINE BUCKET ASSEMBLY AND TURBINE SYSTEM - A turbine bucket assembly and turbine system are disclosed. The assembly includes a single-lobe joint having an integral platform, the joint having a first axial length; a segmented airfoil having a root segment extending radially outward from the platform and a tip segment coupled to the root segment, the tip segment having a second axial length being less than the first axial length; and a turbine wheel having a receptacle with a geometry corresponding to the single-lobe joint and being coupled to the single-lobe joint. The tip segment includes a tip segment material, the root segment includes a root segment material, and the turbine wheel includes a turbine wheel material having a lower heat resistance and a higher thermal expansion than the root segment material and the tip segment material. | 12-03-2015 |
| 20150345310 | TURBINE BUCKET ASSEMBLY AND TURBINE SYSTEM - A turbine bucket assembly and turbine system are disclosed. The assembly includes a single-lobe joint having an integral platform, the joint having a first axial length; a non-segmented airfoil extending radially outward from the integral platform and having a tip end with a second axial length, the second axial length being less than the first axial length; and a turbine wheel having a receptacle with a geometry corresponding to the single-lobe joint and being coupled to the single-lobe joint. The joint and the non-segmented airfoil include a turbine bucket material, and the turbine wheel includes a turbine wheel material, the turbine wheel material having a lower heat resistance and a higher thermal expansion than the turbine bucket material. | 12-03-2015 |
| 20150345314 | TURBINE BUCKET ASSEMBLY AND TURBINE SYSTEM - A turbine bucket assembly and turbine system are disclosed. The turbine bucket assembly includes a single-lobe joint having an integral platform, the joint having a first axial length; a non-segmented airfoil having a root section and a tip section integral with the root section, the tip section having a tip end with a second axial length, the second axial length being less than the first axial length; and a turbine wheel having a receptacle with a geometry corresponding to the single-lobe joint and being coupled to the single-lobe joint. The turbine wheel includes a turbine wheel material and the single-lobe joint and the non-segmented airfoil include a turbine bucket material, the turbine bucket material having a higher heat resistance and a lower thermal expansion than the turbine wheel material. | 12-03-2015 |
| 20150345325 | ROTATING MACHINERY MONITORING SYSTEM - A method for monitoring rotating component includes receiving a continuous waveform sensor signal from a sensor apparatus, retaining the continuous waveform in a memory, and isolating at least one characteristic and/or at least one period of the waveform. The isolated characteristic and/or period is analyzed thereby determining the presence of a waveform anomaly. | 12-03-2015 |
| 20150354362 | GAS TURBINE ENGINE AIRFOIL - An airfoil for a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a total chord length and a span position and corresponds to a curve that has an increasing total chord length from the 0% span position to a first peak. The first peak occurs in the range of 45-65% span position, and the curve either remains generally constant or has a decreasing total chord length from the first peak to the 100% span position. The total chord length is at the 0% span position in the range of 8.2-10.5 inches (20.8-26.7 cm). | 12-10-2015 |
| 20150354363 | GAS TURBINE ENGINE AIRFOIL - An airfoil for a turbine engine includes an airfoil having pressure and suction sides extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a curve corresponding to a relationship between a trailing edge sweep angle and a span position. The trailing edge sweep angle is in a range of 10° to 20° in a range of 40-70% span position. The trailing edge sweep angle is positive from 0% span to at least 95% span. | 12-10-2015 |
| 20150354367 | GAS TURBINE ENGINE AIRFOIL - An airfoil for a turbine engine includes an airfoil having pressure and suction sides extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a leading edge dihedral and a span position. The leading edge dihedral is negative from the 0% span position to the 100% span position. A positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning. | 12-10-2015 |
| 20150354372 | GAS TURBINE ENGINE COMPONENT WITH ANGLED APERTURE IMPINGEMENT - A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a body portion that includes a first wall spaced apart from a second wall. At least one rib extends between the first wall and the second wall and at least one aperture extends through the at least one rib. The at least one aperture is angled relative to a radial axis of the at least one rib. | 12-10-2015 |
| 20150354375 | Hybrid Fan Blade Biscuit Construction - An airfoil for a gas turbine engine is disclosed. The airfoil may include a first portion including a first slot, a second portion including a second slot, and a biscuit disposed within the first slot and the second slot. The first portion and the second portion may be joined by the biscuit. A method for constructing an airfoil is also disclosed. The method may include making a first slot on a sheath, the first slot sized to fit a first part of a biscuit; making a second slot on a body, the second slot sized to fit a second part of the biscuit; and joining the sheath and the body together through a biscuit joint, the biscuit disposed within the first slot and the second slot. | 12-10-2015 |
| 20150354376 | ENHANCED PROTECTION FOR ALUMINUM FAN BLADE VIA SACRIFICIAL LAYER - A component is described which may comprise a structure formed from a material selected from the group consisting of aluminum and an aluminum alloy. The component may further comprise a sacrificial layer in electrical contact with at least a portion of a surface of the structure. The sacrificial layer may protect the surface from localized corrosion and may comprise an alloy that is more anodic than the material forming the structure. The alloy may be selected from the group consisting of an aluminum alloy and a zinc alloy. | 12-10-2015 |
| 20150354403 | OFF-LINE WASH SYSTEMS AND METHODS FOR A GAS TURBINE ENGINE - The present application and the resultant patent provide a wash system for a gas turbine engine. The wash system may include a water source containing a volume of water therein, and a surface filming agent source containing a volume of a surface filming agent therein. The wash system also may include a mixing chamber in fluid communication with the water source and the surface filming agent source, wherein the mixing chamber is configured to mix the water and the surface filming agent therein to produce a film-forming mixture. The film-forming mixture may be a liquid-gas mixture of the surface filming agent in a liquid phase and the water in a gaseous phase. The wash system further may include a number of supply lines in fluid communication with the mixing chamber, wherein the supply lines are configured to direct the film-forming mixture into the gas turbine engine. | 12-10-2015 |
| 20150354449 | Below Wing Reverse Core Gas Turbine Engine With Thrust Reverser - A core engine includes a compressor section, a combustor and a turbine section, with the turbine section being closest to a fan, the combustor section and then the compressor section being positioned further away from the fan relative to the turbine section. A downstream end of a nozzle has at least one pivoting shell and an actuator pivots the shell between a stowed position and a deployed position. A mount bracket is mounted at one circumferential location of the engine. The shell moves in a direction having at least a component perpendicular to a vertical direction defined perpendicular to a top surface of the mount bracket. | 12-10-2015 |
| 20150361796 | TURBINE PLATFORM REPAIR USING LASER CLAD - A method of restoring a gas turbine engine component includes removing a defect section of a turbine engine component according to a template. The template is produced based upon common defect sections from other turbine engine components. A laser cladding is used to build a replacement section in place of the defect section. Thus, the turbine engine component is restored to near its original shape. | 12-17-2015 |
| 20150361797 | GAS TURBINE ENGINE AIRFOIL - An airfoil for a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a stacking offset and a span position that is at least a third order polynomial curve that includes at least one positive and negative slope. The positive slope leans aftward and the negative slope leans forward relative to an engine axis. The positive slope crosses an initial axial stacking offset corresponding to the 0% span position at a zero-crossing position. A first axial stacking offset X | 12-17-2015 |
| 20150361882 | GAS TURBINE ENGINE WITH DISTRIBUTED FANS AND BYPASS AIR MIXER - A gas turbine engine comprises a plurality of distributed fan rotors. A gas generator has a core fan rotor. at least one compressor rotor. at least one gas generator turbine rotor, and a combustion section. A fan drive turbine is downstream of the at least one gas generator turbine rotor. A shaft is configured to be driven by the fan drive turbine, the shaft engaging gears to drive the plurality of distributed fan rotors. The core fan rotor delivers a portion of air into the at least one compressor rotor, and a portion of bypass air into a bypass duct which bypasses the gas generator. The bypass air mixes with products of combustion downstream of the at least one gas generator turbine rotor. | 12-17-2015 |
| 20150361898 | Heat Shield for Gas Turbine Engine Gearbox - A heat shield ( | 12-17-2015 |
| 20150361900 | GAS TURBINE ENGINE WITH PAIRED DISTRIBUTED FAN SETS - A gas turbine engine comprises a gas generator rotating along a first axis of rotation, with at least one compressor rotor, at least one gas generator turbine rotor and a combustion section. A fan drive turbine rotates along a second axis of rotation, downstream of at least one gas generator turbine rotor. The fan drive turbine drives a pair of shaft portions extending in opposed directions beyond the axis of rotation of the gas generator. | 12-17-2015 |
| 20150361901 | GAS TURBINE ENGINE COUPLING STACK - The present disclosure relates generally a system for preventing relative rotation between three components. The three components include tabs and slots such that at least one first tab and at least one second tab on one component is disposed within at least one slot formed in the other two components to prevent relative rotation between any two of the three components. | 12-17-2015 |
| 20150367274 | System and Method for Producing Carbon Dioxide - A system for producing carbon dioxide including a collection subsystem configured to collect a process gas, the process gas including a hydrocarbon, a combustion subsystem configured to combust the hydrocarbon in the process gas and output a gaseous combustion effluent, wherein the gaseous combustion effluent includes carbon dioxide and water, and a separation subsystem configured to separate the carbon dioxide from the gaseous combustion effluent. | 12-24-2015 |
| 20150369045 | REDUCED VIBRATORY RESPONSE ROTOR FOR A GAS POWERED TURBINE - A rotor for a turbomachine includes a rim defining a base of a rotor, an airfoil shaped blade extending from said rim and defining a chord line and a bore extending from said rim opposite said airfoil shaped blade. The rim further includes at least one rail extending away from said airfoil shaped blade. | 12-24-2015 |
| 20150369046 | Low Speed Fan for Gas Turbine Engines - In accordance with one aspect of the disclosure, a fan section for a gas turbine engine is disclosed. The fan may include a rotor disk and a plurality of airfoils fixedly attached to and supported by the rotor disk as a single unitary piece. The airfoils may extend radially outward from the rotor disk with respect to an engine axis. The rotor disk may be made of metal and the airfoils may each be made at least partially of an organic matrix composite. | 12-24-2015 |
| 20150369061 | DOUBLE SNAPPED COVER PLATE FOR ROTOR DISK - According to an exemplary embodiment of this disclosure, among other possible things a gas turbine engine includes a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor section, a plurality of turbine disks in at least one of the compressor section and the turbine sections, at least one cover plate corresponding to at least one of the turbine disks, each of the cover plates includes, at least two snaps connected via a webbing portion, and a bore region radially inward of the at least two snaps and connected to at least one of the at least two snaps via the webbing portion. | 12-24-2015 |
| 20150377040 | REMOVABLE FILM FOR AIRFOIL SURFACES - A fan section to be incorporated into a gas turbine engine has a rotor and a plurality of fan blades. The fan blades deliver air into a bypass duct defined inwardly of the nacelle and into a core engine. There are static vanes inward of the nacelle. A surface of the fan section is provided with a removable film material. A gas turbine engine and a method of refurbishing a surface are also disclosed. | 12-31-2015 |
| 20150377041 | Lock for Retaining Minidisks with Rotors of a Gas Turbine Engine - In accordance with one aspect of the disclosure, a rotor is disclosed. The rotor may include a disk having a central axis, an airfoil radially extending from the disk, a bayonet tab extending radially from the disk, and a lock. The lock may further include a short tab and a long tab, both extending radially from the disk and in an axial direction with respect to the central axis. The long tab may have a greater axial length than the short tab. | 12-31-2015 |
| 20150377048 | STATOR VANE ASSEMBLY AND METHOD THEREFORE - A stator vane assembly includes a plurality of distinct vane segments that each respectively include a first platform, a second platform and at least one vane airfoil connected at opposed ends thereof to the first platform and the second platform. The first platforms meet at distinct first joints with each other, and the second platforms meet at distinct second joints with each other such that the plurality of vane segments forms an annular structure. | 12-31-2015 |
| 20150377071 | GAS TURBINE ENGINE HYDRAULICALLY OPERATED NACELLE LATCH - A method of actuating a hydraulic latch and a fan duct includes providing a pressurized fluid for actuating a latch, providing a valve to control flow of pressurized fluid to the latch and the fan duct, and selectively opening the valve, whereby the pressurized fluid opens the fan duct. A gas turbine engine includes a fan duct with an inner structure surrounding an engine core, a fan case surrounding a fan, and at least one latch. The at least one latch secures a first portion of the fan duct inner structure to a core engine fame or to a second portion of the fan duct inner structure. The at least one latch is also configured to secure the second portion of the fan duct inner structure to the core engine fame or to the first portion of the fan duct inner structure. | 12-31-2015 |
| 20150377073 | TITANIUM ALUMINIDE TURBINE EXHAUST STRUCTURE - A turbine exhaust case for a gas turbine engine includes a multiple of CMC turbine exhaust case struts between a CMC core nacelle aft portion and a CMC tail cone. | 12-31-2015 |
| 20150377123 | TURBINE SECTION OF HIGH BYPASS TURBOFAN - A turbofan engine comprises a fan having fan blades. A compressor is in communication with the fan section. The fan is configured to communicate a portion of air into a bypass path defining a bypass area outwardly of the compressor and a portion into the compressor. A bypass ratio is defined as air communicated through the bypass path relative to air communicated to the compressor being greater than about 6.0. A combustor is in fluid communication with the compressor. A turbine is in communication with the combustor. The turbine has a first turbine section that includes two or more stages and a second turbine section that includes at least two stages. A ratio of airfoils in the first turbine section to the bypass ratio is less than about 170. The first turbine section includes a maximum gas path radius. A ratio of the maximum gas path radius to a maximum radius of the fan blades is less than about 0.50. A speed reduction mechanism is coupled to the fan and rotatable by the turbine. | 12-31-2015 |
| 20150377124 | TURBINE SECTION OF HIGH BYPASS TURBOFAN - A turbofan engine includes an engine case, a gaspath through the engine case, a fan having a circumferential array of fan blades, a compressor in fluid communication with the fan, a combustor in fluid communication with the compressor, and a turbine in fluid communication with the combustor. The turbine has a fan drive turbine section having 3 to 6 blade stages. A speed reduction mechanism couples the fan drive turbine section to the fan. A bypass area ratio is between about 8.0 and about 20.0. A ratio of maximum gaspath radius along the fan drive turbine section to maximum radius of the fan is less than about 0.50. A ratio of a turbine section airfoil count to the bypass area ratio is between about 10 and about 170. The fan drive turbine section airfoil count being the total number of blade airfoils and vane airfoils of the fan drive turbine section. | 12-31-2015 |
| 20160001356 | CAST COMPONENT HAVING CORNER RADIUS TO REDUCE RECRYSTALLIZATION - A cast component includes a cast body that has a single crystal microstructure and an internal corner bounding an internal cavity. The single crystal microstructure defines a critical internal residual stress with respect to investment casting of the cast body using a refractory metal core beyond which the single crystal microstructure recrystallizes under a predetermined condition. The internal corner has a corner radius that is greater than a critical corner radius below which an amount of internal residual stress in the single crystal microstructure exceeds the critical internal residual stress. The internal cavity includes a cross section less than about 20 mils near the corner radius. | 01-07-2016 |
| 20160003048 | Airfoil with Thickened Root and Fan and Engine Incorporating Same - In accordance with one aspect of the disclosure, an airfoil is disclosed. The airfoil may include a platform and a blade extending from the platform. The blade may have a root proximate the platform and a tip radially outward from the platform. The root may have a greater thickness than a cross-section at about a quarter-span of the blade or greater. | 01-07-2016 |
| 20160003061 | Hollow Fan Blade with Extended Wing Sheath - A fan blade for a turbomachinery fan and methods for fabricating a fan blade for a turbomachinery fan are disclosed. The fan blade for a turbomachinery fan includes an airfoil having a leading edge and a trailing edge in a chordwise direction, a tip and a root in a spanwise direction, a suction side and a pressure side. The fan blade includes a sheath including a solid portion that covers the leading edge of the airfoil, a first wing attached to the suction side of the airfoil, and a second wing attached to the pressure side of the airfoil. Construction of the fan blade includes one or more hollow cavities between the suction side and the pressure side of the airfoil. | 01-07-2016 |
| 20160003070 | GAS TURBINE ENGINE STATOR VANE PLATFORM REINFORCEMENT - A stator vane for a gas turbine engine includes a first platform and a second platform radially spaced apart from one another. The first and second airfoils are circumferentially spaced from one another and interconnect the first and second platforms. The first platform has a gas path side facing the airfoils and a non-gas path side opposite the gas path side. A circumferentially extending rail provided on the first platform extends radially outward from the gas path side to the non-gas path side to form a pocket on the non-gas path side between the first platform and the rail. A reinforcement is arranged in the pocket and joins the first platform and the rail. The reinforcement includes a variable thickness in the circumferential direction and is arranged generally centrally between the first and second airfoils. | 01-07-2016 |
| 20160003083 | ABRADABLE SEAL INCLUDING AN ABRADABILITY CHARACTERISTIC THAT VARIES BY LOCALITY - An abradable seal for a gas turbine engine includes a seal body that has a seal side and a non-seal side. The seal body includes an abradability characteristic that varies by locality. | 01-07-2016 |
| 20160003094 | CMC CORE COWL AND METHOD OF FABRICATING - A CMC core cowl for an aircraft gas turbine engine. The ceramic core cowl comprises an interlaced fiber structure having fibers oriented in substantially transverse directions, and a ceramic matrix surrounding the ceramic fiber structure. The core cowl further comprises several panels. The ceramic fiber and matrix are formed into a substantially cylindrical shape extending from a fore end at the fan outlet guide vanes to an aft end at the low pressure turbine outlet guide vanes. The CMC core cowl includes a means for mechanical attachment circumferentially oriented around the fore end and the aft end with mating parts. The CMC core cowl further includes additional plies oriented in a third preselected direction, thereby providing additional strength for mechanical attachment. | 01-07-2016 |
| 20160003194 | Asymmetric Fan Nozzle in High-BPR Separate-Flow Nacelle - A fan nozzle for an aircraft gas turbine engine is comprised of a core engine cowl that is disposed within a fan cowl so that an air flow area is defined therebetween. The core engine cowl and fan cowl are disposed around a horizontal central plane. The fan cowl has a substantially circular shape and is formed of an upper substantially semi-circular portion having a first radius and a lower substantially semi-circular portion having a second radius. The core engine cowl has a substantially circular shape and is formed of an upper substantially semi-circular portion having a third radius and a lower substantially semi-circular portion having a third radius. The upper substantially semi-circular portion of the core engine cowl includes a left arcuate member and a right arcuate member. The second radius is less than the first radius and the third radius is less than the fourth radius. | 01-07-2016 |
| 20160010470 | Frangible Sheath for a Fan Blade of a Gas Turbine Engine | 01-14-2016 |
| 20160010475 | CANTILEVER STATOR WITH VORTEX INITIATION FEATURE | 01-14-2016 |
| 20160010549 | OIL BAFFLES IN CARRIER FOR A FAN DRIVE GEAR SYSTEM | 01-14-2016 |
| 20160010557 | GAS TURBINE SILENCER, AND GAS TURBINE PROVIDED WITH SAME | 01-14-2016 |
| 20160010565 | GAS TURBINE ENGINE WITH FAN VARIABLE AREA NOZZLE FOR LOW FAN PRESSURE RATIO | 01-14-2016 |
| 20160010590 | NOZZLE ARRANGEMENT FOR A GAS TURBINE ENGINE | 01-14-2016 |
| 20160017732 | Off-Cambered Vanes for Gas Turbine Engines - An off-cambered vane for a guide vane assembly in a gas turbine engine is described. The guide vane assembly may comprise a nominal vane having a tip portion, a mid-span portion, and a hub portion. The mid-span portion of the nominal vane may adopt a nominal geometry and the hub portion of the nominal vane may adopt a common geometry. The guide vane assembly may further comprise an off-cambered vane having a tip portion, a mid-span portion, and a hub portion. The mid-span portion of the off-cambered vane may deviate variably with respect to the nominal geometry and at least one of the hub portion and the tip portion may adopt the common geometry. | 01-21-2016 |
| 20160017754 | MID-TURBINE FRAME ROD AND TURBINE CASE FLANGE - A turbine section of a gas turbine engine includes a first turbine supported for rotation about an axis, a second turbine spaced axially aft of the for first turbine section for rotation about the axis, and a mid-turbine frame disposed between the first turbine and the second turbine defining a passage between the first turbine and the second turbine. A first case surrounds the first turbine and a second case surrounding the second turbine and attached to the first case. The mid-turbine frame is disposed between the first turbine section and the second turbine section and includes at least one support structure extending through an interface between the first turbine case and the second turbine case. | 01-21-2016 |
| 20160017755 | COMMON JOINT FOR A COMBUSTOR, DIFFUSER, AND TOBI OF A GAS TURBINE ENGINE - An assembly for a gas turbine engine is disclosed. The gas turbine engine may have a combustor, a diffuser, and a tangential onboard injector. The assembly may include a common joint between the combustor, the diffuser, and the tangential onboard injector. | 01-21-2016 |
| 20160017811 | GEARBOX MOUNTING ASSEMBLY - An accessory gearbox mounting assembly includes a gearbox bracket engaged to a case bracket mounted on an engine case. The case bracket includes a first boss receivable within a first cavity of the engine case and a locator opening through the first boss. The gearbox bracket is mountable to a gearbox case and includes a pin receivable within the locator opening for transferring load from the gearbox bracket to the case bracket. | 01-21-2016 |
| 20160024934 | ALUMINUM FAN BLADE CONSTRUCTION WITH WELDED COVER - An airfoil includes, among other possible things, a main body extending between a leading edge and a trailing edge. Channels are formed into the main body, with a plurality of ribs extending intermediate the channels. A cover skin is attached to the main body. The cover skin is welded to the main body with a weld at outer edges. An adhesive is placed between inner surfaces of the cover skin and the main body. The adhesive is deposited inwardly of the outer edges of the cover skin. A method of constructing an airfoil is also disclosed as is a gas turbine engine. | 01-28-2016 |
| 20160024944 | TRANSIENT LIQUID PAHSE BONDED TURBINE ROTOR ASSEMBLY - A rotor assembly for a turbine engine includes a rotor disk constructed of a first material. Multiple rotor blades constructed of a second material are connected to the rotor disk via a diffusion material. | 01-28-2016 |
| 20160024949 | MID-TURBINE FRAME AND GAS TURBINE ENGINE INCLUDING SAME - A mid-turbine frame for a gas turbine engine ducts gases between a high pressure turbine and a low pressure turbine. The mid-turbine frame may include an outer flowpath ring, an inner flowpath ring, and a plurality of vanes extending therebetween. The outer flowpath ring comprises a unitary structure, while the inner flowpath ring and the plurality of vanes comprises a plurality of segments. | 01-28-2016 |
| 20160025004 | LOW NOISE TURBINE FOR GEARED TURBOFAN ENGINE - A gas turbine engine comprises a fan, a compressor section including a compressor having a low pressure portion and a high pressure portion, a combustor section, and a turbine having a downstream portion. A gear reduction effects a reduction in the speed of the fan relative to an input speed to the fan. The downstream portion of the turbine has a number of turbine blades in each of a plurality of rows of the downstream turbine portion. The turbine blades operate at least some of the time at a rotational speed. The number of blades and the rotational speed are such that the following formula holds true for at least one of the blade rows of the downstream turbine: (number of blades×speed)/60≧5500. The rotational speed is an approach speed in revolutions per minute. A method of designing a gas turbine engine and a turbine module are also disclosed. | 01-28-2016 |
| 20160025013 | TURBINE ENGINE FACE SEAL ARRANGEMENT INCLUDING ANTI-ROTATION FEATURES - A turbine engine includes a main shaft bearing compartment seal. The seal includes at least an approximately circular seal portion and a seal carrier disposed about the approximately circular seal portion. A plurality of anti-rotation pins maintain the seal carrier in position relative to a housing and are received in an anti-rotation slot of the seal carrier. | 01-28-2016 |
| 20160032728 | GAS TURBINE ENGINE AIRFOIL - An airfoil of a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a tangential stacking offset and a span position that is at least a third order polynomial curve that includes at least one positive and negative slope. The positive slope leans toward the suction side and the negative slope leans toward the pressure side. An initial slope starts at the 0% span position is either zero or positive. The first critical point is less than 15% span. | 02-04-2016 |
| 20160032729 | Composite Fan Blade - A composite fan blade for a gas turbine engine is disclosed. The fan blade may include a core being made of a first material and a shell enclosing the core. The shell may be made of a second material and the second material may have less plasticity than the first material. | 02-04-2016 |
| 20160032739 | AXIAL FLOW COMPRESSOR AND GAS TURBINE EQUIPPED WITH AXIAL FLOW COMPRESSOR - An axial flow compressor comprising: a rotor blade and a stator blade to intake air from an atmosphere and compress the air; a wheel to secure the rotor blade; a casing to secure the stator blade; wherein a wheel dovetail which is machined on an outer circumferential surface of the wheel to have a certain angle with respect to a rotational axis of the axial flow compressor; and a rotor blade dovetail which is machined on an inner circumferential side of the rotor blade to secure by being fitted to the wheel dovetail, characterized in that: a cut surface is formed on at least one of an upstream side and a downstream side of the rotor blade dovetail by cutting a part of the rotor blade dovetail in a surface form. | 02-04-2016 |
| 20160032740 | VARIABLE-PITCH ROTOR WITH REMOTE COUNTERWEIGHTS - A pitch control mechanism includes: a rotor structure configured for rotation about a longitudinal axis; a row of blades carried by the rotor structure, each blade having an airfoil and a trunnion mounted for pivoting movement relative to the rotor structure, about a trunnion axis which is perpendicular to the longitudinal axis; a unison ring interconnecting the blades; an actuator connected to the unison ring and the rotor structure, operable to move the unison ring relative to the rotor structure; at least one moveable counterweight carried by the rotor structure, remote from the blades; and an interconnection between the blades and the counterweight, such that movement of the counterweight causes a change in the pitch angle of the blades. | 02-04-2016 |
| 20160032754 | ADJUSTABLE BLADE OUTER AIR SEAL APPARATUS - An adjustable blade outer air seal apparatus includes a case that extends circumferentially around an axis, a support ring non-rigidly mounted to the case on spring connections radially inwards of the case, whereby the support ring floats with respect to the case, and at least one blade outer air seal segment that is radially adjustable relative to the support ring. | 02-04-2016 |
| 20160032779 | CASTELLATED LATCH MECHANISM FOR A GAS TURBINE ENGINE - A gas turbine engine includes a fan duct including a fan duct inner structure that surrounds a core engine, a fan case that surrounds a fan, a core engine frame, and at least one mechanism configured to secure a portion of the fan duct inner structure to a portion of the core engine frame. The at least one mechanism includes a castellated arcuate portion mounted to one of the fan duct inner structure and the core engine frame and an inwardly projecting retaining feature mounted to the other of the fan duct inner structure and the core engine frame. The castellated arcuate portion is rotatable about an engine central longitudinal axis to position a feature of the castellated arcuate portion proximate to a portion of the inwardly projecting retaining feature to latch the fan duct inner structure. | 02-04-2016 |
| 20160032831 | Spinner Aft-Extended Forward Return Flange - A spinner for a fan assembly of a gas turbine engine, a method of fabricating a spinner for a fan assembly of a gas turbine engine, and a gas turbine engine are disclosed. The fan section may include a nosecap. The spinner may include a forward end, an aft end, and a return flange associated with the forward end. The return flange may include a forward flange extending forward towards the nosecap and an aft flange extending aft towards the aft end. | 02-04-2016 |
| 20160040538 | ALUMINUM FAN BLADE TIP WITH THERMAL BARRIER - A fan blade for a gas turbine engine is described. The fan blade may comprise a body portion formed from a metallic material, and it may include a suction side, a pressure side, a leading edge, a trailing edge, and a tip. A coating may be applied to the tip, and the coating may have a thermal conductivity of no more than about 10 watt per meter kelvin. The coating may be a thermal barrier coating comprising yttria-stabilized zirconia. | 02-11-2016 |
| 20160040539 | ENGINE COMPONENT HAVING SUPPORT WITH INTERMEDIATE LAYER - Disclosed is a gas turbine engine component, and a method for forming the component. The component includes a first portion, a second portion formed separately from the first portion, and an intermediate layer provided between the first portion and the second portion. | 02-11-2016 |
| 20160040547 | BLADE OUTER AIR SEAL WITH SECONDARY AIR SEALING - A blade outer air seal (BOAS) for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion. A retention flange extends from one of the leading edge portion and the trailing edge portion and a seal contacts the retention flange. | 02-11-2016 |
| 20160040555 | COMPRESSOR CASE SNAP ASSEMBLY - A turbine engine case spacer includes an air seal and a flange. The air seal is snap fit to the flange, and each flange includes at least two distinct material layers. | 02-11-2016 |
| 20160040594 | COUNTER-ROTATING LOW PRESSURE TURBINE WITHOUT TURBINE EXHAUST CASE - A gas turbine engine comprises a compressor section, a combustor section downstream of the compressor section, and a turbine section downstream of the combustor section. A mid-turbine frame includes an outer case portion and is configured to support the turbine section. At least one shaft defines an axis of rotation, and the turbine section comprises an inner rotor directly driving the shaft. The inner rotor includes an inner set of blades. An outer rotor is positioned immediately adjacent to the outer case portion and has an outer set of blades interspersed with the inner set of blades. The outer rotor is configured to rotate in an opposite direction about the axis of rotation from the inner rotor. A gear system is positioned downstream of the combustor section, is mounted to the mid-turbine frame, and is coupled to the outer rotor to drive the at least one shaft. | 02-11-2016 |
| 20160040600 | MULTI-AXIAL BRUSH SEAL - A seal assembly, includes a first brush supported between first and second plates, a second brush supported on the first and second plates transverse to the first brush seal, and a third plate attached to the second brush. | 02-11-2016 |
| 20160047247 | Fan Blade Grounding Tab - An airfoil for a gas turbine engine is disclosed. The airfoil may include a leading edge, a sheath on the leading edge; and a grounding element connected to the sheath. The grounding element may have a radially extending tab, and may be configured for connection to a component of the gas turbine engine to form a ground path from the sheath to the component. | 02-18-2016 |
| 20160047255 | GAS TURBINE ENGINE BLADE CONTAINMENT SYSTEM - A gas turbine engine blade containment system is disclosed. The blade containment system may include a generally cylindrical casing being made of a first material, and a generally cylindrical ring being made of a second material coaxially surrounding the casing, at least some portion of the ring metallurgically bonded to the casing. | 02-18-2016 |
| 20160047313 | BUSHING FOR JOINING TURBOMACHINE COMPONENTS - A bushing includes a body extending from a first end to a second end through an intermediate portion. The body includes a passage having a first centerline extending from the first end to the second end. A first alignment member is formed on a first section of the intermediate portion at the first end. The first alignment member includes an outer surface having a plurality of splines configured to be received by a first component to be joined. A second alignment member is formed on a second section of the intermediate portion. The second alignment member includes a second centerline that is off-set relative to the first centerline and is configured to be received by a second component to be joined with the first component. | 02-18-2016 |
| 20160053617 | ROTORS WITH MODULUS MISTUNED AIRFOILS - A rotor assembly for a gas turbine engine includes a rotor defining an outer periphery; and a plurality of blades attached to the outer periphery. The plurality of blades includes a material property different than the other of the plurality of blades to provide mistuning of the rotor. | 02-25-2016 |
| 20160069187 | GAS TURBINE ENGINE AIRFOIL - An airfoil of a turbine engine according to an example of the present disclosure includes, among other things, pressure and suction sides extending in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between a gap/chord ratio and span position that defines a curve with a gap/chord ratio having a portion with a negative slope. | 03-10-2016 |
| 20160069258 | TURBINE SYSTEM - A turbine system including a compressor ( | 03-10-2016 |
| 20160069270 | TURBOFAN ENGINE FRONT SECTION - A turbofan engine includes a geared architecture for driving a fan about an axis. The geared architecture includes a sun gear rotatable about an axis, a plurality of planet gears driven by the sun gear and a ring gear circumscribing the plurality of planet gears. A carrier supports the plurality of planet gears. The geared architecture includes a power transfer parameter (PTP) defined as power transferred through the geared architecture divided by gear volume multiplied by a gear reduction ratio and is between about 219 and 328. | 03-10-2016 |
| 20160076380 | INCIDENCE-TOLERANT, HIGH-TURNING FAN EXIT STATOR - A gas turbine engine component is described. The gas turbine engine component includes an inner diameter edge, an outer diameter edge, a trailing edge and a leading edge. The leading edge has a positive (aft) aerodynamic sweep across substantially an entire span of the leading edge. The gas turbine engine component has a camber angle greater than 50 degrees across substantially an entire span of the component. The gas turbine engine component may have asymmetrical tangential stacking of the component in the radial direction. | 03-17-2016 |
| 20160076386 | Tangential Blade Root Neck Conic - A root extending from a platform of an airfoil is disclosed. The root may include a first portion having a generally cylindrical shape, and a second portion extending from the first portion to the platform. The second portion may have a circumference larger than a circumference of the first portion. | 03-17-2016 |
| 20160076392 | Endwall Contouring for Airfoil Rows With Varying Airfoil Geometries - An airfoil array for a gas turbine engine comprises a plurality of airfoils spaced circumferentially apart from each other about an engine center axis. Each airfoil is associated with a platform. An endwall extends circumferentially about the engine center axis, and is defined by adjacent platforms. Each pair of adjacent platforms are separated from each other by a gap. An endwall contour shape extends from a first location on one side of the gap to a second location on an opposite of the gap. The endwall contour shape is the same for all adjacent platforms. A gas turbine engine and a method of forming an airfoil array for a gas turbine engine are also disclosed. | 03-17-2016 |
| 20160076454 | SEALING ARRANGEMENT AT THE INTERFACE BETWEEN A COMBUSTOR AND A TURBINE OF A GAS TURBINE AND GAS TURBINE WITH SUCH A SEALING ARRANGEMENT - A sealing arrangement is provided at the interface between a combustor and a turbine of a gas turbine. The turbine includes guiding vanes at its inlet. The guiding vanes are each mounted within the turbine at their outer diameter by means of rear outer diameter vane hook and are each at their inner diameter in sealing engagement by means of a front inner diameter vane tooth with a honeycomb seal arranged at the corresponding inner diameter part of the outlet of said combustor. The rear outer diameter vane hook allows a relative movement of the guiding vane in form of a rotation around the rear outer diameter vane hook. The sealing adapts the front inner diameter vane tooth and the corresponding honeycomb seal in their configuration to the rotating relative movement of the guiding vane, such that the compression of said honeycomb seal through transients of the gas turbine is reduced. | 03-17-2016 |
| 20160076460 | NACELLE AND COMPRESSOR INLET ARRANGEMENTS - A gas turbine engine includes a nacelle defining a centerline axis and an annular splitter radially inward from the nacelle. A spinner is radially inward of the nacelle forward of a compressor section. A fan blade extends from a fan blade platform. A distance X is the axial distance from a first point to a second point, wherein the first point is defined on a leading edge of the annular splitter and the second point is defined on a leading edge of the fan blade where the fan blade meets the fan blade platform. A distance H is the radial distance from the first point to the second point. The relative position of the first point and the second point is governed by the ratio of X/H≧1.5 for reducing foreign object debris (FOD) intake into the compressor section. | 03-17-2016 |
| 20160084096 | CLAMPED VANE ARC SEGMENT HAVING LOAD-TRANSMITTING FEATURES - A vane arc segment includes a radially inner and outer platforms and an airfoil mechanically clamped between the platforms. The airfoil has an airfoil section that extends radially between radially inner and outer fairing platforms. At least one of the fairing platforms includes forward and aft sides, circumferential sides, and a gas path side and an opposed radial side. The radial side includes a plurality of protrusions that have faces that are oriented substantially normal to, respectively, radial, tangential, and axial load transmission directions of the airfoil such that the faces, respectively, primarily bear radial, tangential, and axial load transmissions of the airfoil. | 03-24-2016 |
| 20160084100 | SYSTEM AND APPARATUS FOR SEAL RETENTION AND PROTECTION - A sheath and seal assembly for protecting, containing and insulating a seal is provided. The seal may be installable within the sheath forming a seal-sheath assembly. The assembly may be capable of being installed in the hot section of a gas turbine. The sheath may be a woven, braided, and/or chain link structure. The sheath may be capable of allowing pressure to be conducted to a portion of the seal to load the seal against one or more portions of a housing. | 03-24-2016 |
| 20160090849 | FAN BLADE WITH STATIC DISSIPATIVE COATING - The present disclosure relates generally to a fan blade for use in a gas turbine engine, wherein the fan blade includes a metallic fan blade body, including a fan blade body leading edge and body outer surface, extending radially outward from a root, and a static dissipative coating material disposed on the body outer surface. | 03-31-2016 |
| 20160102567 | DIFFUSER FOR A GAS TURBINE - A diffuser for a gas turbine including a support structure and a liner connected by holders. The liner includes a plurality of adjacent segments with overlapping borders. The overlapping borders are clamped to one another and slidingly rest one above the other. | 04-14-2016 |
| 20160102580 | POWER TURBINE INLET DUCT LIP - A power turbine section for a gas turbine engine includes an inlet duct upstream of a first power turbine vane array, the inlet duct including a lip. | 04-14-2016 |
| 20160108746 | VANE ASSEMBLY INCLUDING TWO- AND THREE-DIMENSIONAL ARRANGEMENTS OF CONTINUOUS FIBERS - A vane assembly includes a plurality of airfoils that extend from a first end to a second, opposed end. A first platform is at the first end and is joined to the plurality of airfoils. A second platform is at the second end and is joined to the plurality of airfoils. At least one of the first platform, the second platform and the plurality of airfoils include a three-dimensional arrangement of continuous fibers and at least one different one of the first platform, the second platform and the plurality of airfoils include a two-dimensional arrangement of continuous fibers. | 04-21-2016 |
| 20160108807 | TURBOFAN ENGINE FRONT SECTION - A turbofan engine includes a geared architecture for driving a fan about an axis. The geared architecture includes a sun gear rotatable about an axis, a plurality of planet gears driven by the sun gear and a ring gear circumscribing the plurality of planet gears. A carrier supports the plurality of planet gears. The geared architecture includes a power transfer parameter (PTP) defined as power transferred through the geared architecture divided by gear volume multiplied by a gear reduction ratio and is between about 430 and 645. | 04-21-2016 |
| 20160108808 | Turbofan Engine Bearing and Gearbox Arrangement - A turbofan engine ( | 04-21-2016 |
| 20160108812 | CONDUIT FOR GUIDING LOW PRESSURE COMPRESSOR INNER DIAMETER SHROUD MOTION - Conduits for guiding the motion of an inner diameter shroud of a low pressure compressor of a gas turbine engine are disclosed. The inner diameter shroud has at least three slots formed in one or more radially inwardly extending flanges. Each of the conduits are configured to assemble with a respective one of the at least three slots. Each conduit comprises a bushing having a first panel, and the first panel is capable of being inserted in a respective one of the slots of the inner diameter shroud. The conduit further comprises a bracket capable of being attached to a bearing support of a fan intermediate case of the gas turbine engine. The bushing is capable of being attached to the bracket. A contact between the first panel and the at least one slot of the inner diameter shroud restricts a circumferential rotation of the inner diameter shroud with respect to a central axis of the gas turbine engine when the first panel is inserted in the at least one slot, but allows a radial motion of the inner diameter shroud with respect to the central axis. | 04-21-2016 |
| 20160108821 | RADIALLY FASTENED FIXED-VARIABLE VANE SYSTEM - A split case assembly for a gas turbine engine includes an outer diameter case defining a partial case structure for a gas turbine engine and multiple fixed-variable vanes attached to an inner diameter surface of the outer diameter case. Each of the fixed-variable vanes protrudes radially inward from the outer diameter case. Each of the fixed-variable vanes in the plurality of fixed-variable vanes is interfaced with one of a plurality of inner diameter boxes at a radially inward end of the fixed-variable vane, such that the inner diameter boxes define an inner diameter of a flow path and the outer diameter case defines an outer diameter flow path. Each of the fixed-variable vanes are interfaced with the one of the plurality of inner diameter boxes through at least one inner diameter shoe in a plurality of inner diameter shoes. | 04-21-2016 |
| 20160115801 | MULTI-PIECE TURBINE AIRFOIL - A gas turbine engine includes a compressor, a combustor fluidly connected to the compressor via a flow path, and a turbine fluidly connected to the combustor via the flow path. At least one multi-piece structure is disposed within the flow path such that the multi-piece structure at least partially radially spans the flow path. The at least one multi-piece structure includes a fore portion defining a leading edge, a consumable aft portion defining a trailing edge, a pressure surface at least partially defined by a first surface of the fore portion and a first surface of the consumable aft portion, and a suction surface at least partially defined by a second surface of the fore portion and a second surface of the consumable aft portion. | 04-28-2016 |
| 20160122552 | Abrasive Rotor Coating With Rub Force Limiting Features - The present disclosure relates to an abrasive coating forming a seal material on components of gas turbine engines and a process for forming the abrasive coating. The abrasive coating may be applied to a structure in proximity to at least one section of the gas turbine engine having a plurality of airfoils. The abrasive coating in a first mode of operation of the gas turbine engine is capable of causing wearing of tips of the airfoils that come into contact with the abrasive coating and in a second mode of operation of the gas turbine engine has an interparticle strength sufficient to allow for fracture of the abrasive coating. | 05-05-2016 |
| 20160123176 | HIGH PRESSURE COMPRESSOR ROTOR THERMAL CONDITIONING USING OUTER DIAMETER GAS EXTRACTION - A compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A radially outer housing surrounds an outer diameter of the blades. A lower pressure tap and a higher pressure tap tap air from two distinct locations within the compressor and radially outwardly through the outer housing. A valve selectively delivers at least one of the lower pressure tap and the higher pressure tap to the bore of the disc. A control for the valve is programmed to move the valve to a position delivering the higher pressure tap at a point prior to take-off when the compressor is mounted in a gas turbine engine on an aircraft. A gas turbine engine and a method of operating a gas turbine engine are also disclosed. | 05-05-2016 |
| 20160123177 | VANE ARM WITH INCLINED RETENTION SLOT - The present disclosure includes vane assemblies having a vane lever arm with an inclined profile. The inclined profile may engage with a complementary end of a vane shaft to locate, retain, and prevent rotation of the vane shaft and vane. | 05-05-2016 |
| 20160123234 | HIGH PRESSURE COMPRESSOR ROTOR THERMAL CONDITIONING USING DISCHARGE PRESSURE AIR - A compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A high pressure air tap includes a lower temperature tapped path and a higher temperature tapped path and a valve for selectively delivering one of the lower temperature tapped path and the higher temperature tapped path into the bore of the disc. The valve is operable to selectively block flow of either of the lower pressure and higher pressure tapped paths to the bore of the disc, with the disc including holes to allow air from compressor chambers to communicate with the bore of the disc. A gas turbine engine and a method of operating a gas turbine engine are also disclosed. | 05-05-2016 |
| 20160123236 | COMPRESSOR AND GAS TURBINE - A compressor includes: a plurality of vanes at a vane stage provided to a rotor casing demarcating the primary duct; an air bleed chamber casing that demarcates an air bleed chamber interconnecting with the primary duct; and an air bleed tubing connected to the air bleed chamber casing. Of the plurality of vanes, when a plurality of vanes positioned at a region including the position in the peripheral direction corresponding to the air bleed tubing are a first vane group and a plurality of vanes other than the first vane group are a second vane group, the spacing between the ends at the outside in the radial direction of the vanes that are adjacent in the first vane group is closer than the spacing between the ends at the outside in the radial direction of the vanes that are adjacent in the second vane group. | 05-05-2016 |
| 20160130958 | Gas Turbine Engine Structural Guide Vanes - A structural guide vane is disclosed. The structural guide vane may include a first root, a second root positioned radially outward from the first root and a truss extending between the first root and second root. The truss may be made of a first material. The guide vane may further include an overmold made of a second material that surrounds the truss to form an airfoil. | 05-12-2016 |
| 20160130963 | GAS TURBINE ENGINE AND SEAL ASSEMBLY THEREFORE - The present disclosure relates generally to a hydrostatic advanced low leakage seal having a shoe supported by at least two beams. An anti-vibration beam spacer is disposed in contact with two adjacent beams and operative to mitigate an externally induced vibratory response of the beams. | 05-12-2016 |
| 20160130977 | TURBINE ROTOR SEGMENTED SIDEPLATES WITH ANTI-ROTATION - A segmented sideplate for use in a gas turbine engine is described. The segmented sideplate includes a first plate having a first circumferential edge configured to interface with a complementary circumferential edge. The segmented sideplate also includes a second plate having a second circumferential edge configured to interface with the first circumferential edge. | 05-12-2016 |
| 20160131028 | GAS TURBINE - A gas turbine, in particular an aircraft engine, including a core flow channel (K), in which a first compressor ( | 05-12-2016 |
| 20160131044 | GAS TURBINE ENGINE DRIVING MULTIPLE FANS - A gas turbine engine includes a core engine with a compressor section, a combustor and a turbine. The turbine drives an output shaft, and the output shaft drives at least four gears. Each of the at least four gears extends through a drive shaft to drive an associated fan rotor. | 05-12-2016 |
| 20160131146 | PRESSURE SENSOR SYSTEM FOR CALCULATING COMPRESSOR MASS FLOW RATE USING SENSORS AT PLENUM AND COMPRESSOR ENTRANCE PLANE - A pressure sensor system for a compressor including an inlet bellmouth is disclosed. The system includes a first static pressure sensor positioned within a plane of a plenum that is upstream of the inlet bellmouth; and a second static pressure sensor positioned at an entrance plane of the compressor. A mass flow rate calculator may calculate a mass flow rate based on a pressure differential between the plane of the plenum and the entrance plane of the compressor. | 05-12-2016 |
| 20160138399 | Reinforced Gas Turbine Engine Rotor Disk - A structurally-reinforced rotor disk for a gas turbine engine is disclosed. The rotor disk may comprise a body including a rim configured to support airfoils (which may be separate or integral with the airfoils), an axially-extending bore disposed radially inward of the rim, and a radially-extending web connecting the rim and the bore. The bore may include an axial outer edge and at least one circumferentially-extending annular recess formed axially between the outer edge and the web. The rotor disk may further comprise an annular ring retained in the annular recess, and the annular ring may be formed from a different material than the body of the rotor disk so as to increase a self-sustaining radius of the rotor disk. | 05-19-2016 |
| 20160138416 | VARIABLE FAN NOZZLE USING SHAPE MEMORY MATERIAL - A gas turbine engine includes a fan, a nacelle arranged about the fan, and an engine core at least partially within the nacelle. A fan bypass passage downstream of the fan between the nacelle and the gas turbine engine conveys a bypass airflow from the fan. A nozzle associated with the fan bypass passage is operative to control the bypass airflow. The nozzle includes a shape memory material having a first solid state phase that corresponds to a first nozzle position and a second solid state phase that corresponds to a second nozzle position. | 05-19-2016 |
| 20160146025 | VARIABLE PITCH FAN FOR GAS TURBINE ENGINE AND METHOD OF ASSEMBLING THE SAME - A gas turbine engine is provided. The gas turbine engine includes a core and a variable pitch fan arranged in flow communication with the core. The variable pitch fan has a disk and at least nine fan blades coupled to the disk for rotation together with the disk. The gas turbine engine further includes a rotatable nacelle covering the disk such that the engine has a fan hub radius ratio of between about 0.1 and about 0.4. | 05-26-2016 |
| 20160146040 | Alternating Vane Asymmetry - A vane assembly for a gas turbine engine may include a plurality of vanes being arranged in vane groupings symmetrically spaced circumferentially from each other. Each vane grouping may include at least a first and a second vane. The at least first and second vanes may be spaced from each other at a first pitch. Each vane grouping may be spaced from each other at a second pitch. The first pitch may be dissimilar from the second pitch. | 05-26-2016 |
| 20160153106 | METHOD FOR PRODUCING A METAL UNDERCOAT MADE FROM PLATINUM ON A METAL SUBSTRATE | 06-02-2016 |
| 20160153286 | TURBINE CLEARANCE CONTROL UTILIZING LOW ALPHA MATERIAL | 06-02-2016 |
| 20160153360 | COMPRESSOR END-WALL TREATMENT WITH MULTIPLE FLOW AXES | 06-02-2016 |
| 20160153463 | Fiber Reinforced Spacer for a Gas Turbine Engine | 06-02-2016 |
| 20160153464 | ALUMINUM AIRFOIL | 06-02-2016 |
| 20160153465 | AXIAL COMPRESSOR ENDWALL TREATMENT FOR CONTROLLING LEAKAGE FLOW THEREIN | 06-02-2016 |
| 20160160666 | Pre-Diffuser with Multiple Radii - A pre-diffuser may include a plurality of struts, and each strut may have a leading edge. An upper contour of the leading edge may have a forward end and an aft end, and may include a first radius and a second radius. The first and second radii may be associated with the forward end and the aft end, respectively. The first radius may also be located farther forward than the second radius, and may be larger than the second radius. | 06-09-2016 |
| 20160160667 | DISCOURAGER SEAL FOR A TURBINE ENGINE - A sealing assembly for a turbine engine includes a first stationary component. Also included is a second stationary component, wherein the first stationary component and the second stationary component define a gap therebetween. Further included is a discourager seal in contact with at least one of the first stationary component and the second stationary component, the discourager seal having a lip portion disposed within the gap to reduce a fluid flow through the gap. | 06-09-2016 |
| 20160160867 | ELECTRICALLY COUPLED COUNTER-ROTATION FOR GAS TURBINE COMPRESSORS - A system and method for implementing stage-by-stage counter rotation in a multi-stage axial compressor of a gas turbine engine. The system includes an electrical power generator and an electric motor. A turbine-driven shaft connected to an armature of the electrical generator drives a first plurality of compressor blades. The electrical generator armature induces changing magnetic flux in the stator coils of the electrical generator which generates electrical power that is sent to a power control module. The power control module controls the electrical motor and excites the coils in the electric motor stator which drives the electric motor armature. The electric motor armature drives a second shaft which drives a second plurality of compressor blades in an opposite direction to the first plurality of compressor blades. | 06-09-2016 |
| 20160169017 | CIRCUMFERENTIALLY VARYING AXIAL COMPRESSOR ENDWALL TREATMENT FOR CONTROLLING LEAKAGE FLOW THEREIN | 06-16-2016 |
| 20160169241 | COMPRESSOR AND GAS TURBINE | 06-16-2016 |
| 20160177732 | HOLLOW FAN BLADE FOR A GAS TURBINE ENGINE | 06-23-2016 |
| 20160177751 | BLADE AND GAS TURBINE PROVIDED WITH THE SAME | 06-23-2016 |
| 20160177762 | System and Method Including a Circumferential Seal Assembly to Facilitate Sealing in a Turbine | 06-23-2016 |
| 20160177763 | GAS TURBINE ENGINE SEALING ARRANGEMENT | 06-23-2016 |
| 20160177819 | SIMPLIFIED ENGINE BLEED SUPPLY WITH LOW PRESSURE ENVIRONMENTAL CONTROL SYSTEM FOR AIRCRAFT | 06-23-2016 |
| 20160177820 | LOW PRESSURE ENVIRONMENTAL CONTROL SYSTEM WITH SAFE PYLON TRANSIT | 06-23-2016 |
| 20160186599 | TURBINE ENGINE WITH GUIDE VANES FORWARD OF ITS FAN BLADES - A turbine engine such as a pusher fan engine is provided. This turbine engine includes a nacelle with a bypass flowpath. A fan rotor is configured to propel air out of the bypass flowpath. A plurality of guide vanes are configured to direct the air to the fan rotor. | 06-30-2016 |
| 20160186608 | INTEGRAL GUTTER AND FRONT CENTER BODY - A fan drive gear system for a turbofan engine is disclosed and includes a gear assembly and a front center body. The front center body is an annular case that supports the gear assembly. The front center body includes a passage portion that defines a portion of a core flow path, a forward flange configured for attachment to a first case structure forward of the front center body, and gutter portion disposed on a radially inner side of the front center body for collecting lubricant exhausted from the geared assembly. | 06-30-2016 |
| 20160195010 | VANELESS COUNTERROTATING TURBINE | 07-07-2016 |
| 20160201470 | INTEGRALLY BLADED ROTOR HAVING AXIAL ARM AND POCKET | 07-14-2016 |
| 20160201515 | CONNECTION FOR A FAIRING IN A MID-TURBINE FRAME OF A GAS TURBINE ENGINE | 07-14-2016 |
| 20160376919 | TRUNNION RETENTION FOR A TURBINE ENGINE - A fan for a gas turbine engine is provided. The fan includes a plurality of fan blades, a disk, and a trunnion mechanism for attaching the fan blades to the disk. The disk can be formed of a plurality of individual disk segments, with the trunnion mechanism attaching one of the plurality of fan blades to a respective disk segment. A retention member is also provided. The retention member includes a means for catching a portion of the trunnion mechanism should a primary attachment system of the trunnion mechanism fail. | 12-29-2016 |
| 20160376984 | GEARED TURBOFAN WITH INDEPENDENT FLEXIBLE RING GEARS AND OIL COLLECTORS - A geared turbofan engine includes a fan rotatable about an engine axis. A compressor section compresses air and delivers the compressed air to a combustor where the compressed air is mixed with fuel and ignited to drive a turbine section that in turn drives the fan and the compressor section. A gear system is driven by the turbine section for driving the fan at a speed different than the turbine section. The gear system includes a carrier attached to a fan shaft. A plurality of planet gears are supported within the carrier. Each of the plurality of planet gears includes a first row of gear teeth and a second row of gear teeth supported within the carrier. A sun gear is driven by a turbine section. The sun gear is in driving engagement with the plurality of planet gears. At least two separate ring gears circumscribe the plurality of planet gears. Each of the at least two ring gears are supported by a respective flexible ring gear mount that enables movement relative to an engine static structure. A fan drive gear system for a gas turbine engine is also disclosed. | 12-29-2016 |
| 20170234161 | Flowpath Contouring | 08-17-2017 |
| 20180023396 | ROTOR BLADE FOR A GAS TURBINE ENGINE | 01-25-2018 |
| 20180023418 | ENGINE BEARING DAMPER WITH INTERRUPTED OIL FILM | 01-25-2018 |
| 20190145321 | SYSTEM AND METHOD FOR LIMITING MOVEMENT OF A RETAINING RING | 05-16-2019 |
