| Class / Patent application number | Description | Number of patent applications / Date published |
|---|
| 060805000 | Having turbine | 72 |
| 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 |
| 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 |
| 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 |
| 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 |
| 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 |
| 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 |
| 20120111026 | CASTABLE HIGH TEMPERATURE ALUMINUM ALLOY - A turbine engine includes an airfoil made of an aluminum-rare earth element alloy. | 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 |
| 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 |
| 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 |
| 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 |
| 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 |
| 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 |
| 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 |
| 20120297793 | GAS TURBINE ENGINE - A combustor ( | 11-29-2012 |
| 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 |
| 060806000 | And cooling | 51 |
| 20100115967 | ECCENTRIC CHAMFER AT INLET OF BRANCHES IN A FLOW CHANNEL - A flow channel with a branch channel perpendicular to a main channel including an edge defining an inlet opening of the branch channel is provided. A chamfer is disposed at the upstream edge of the inlet opening and a straight edge is disposed at the downstream edge of the inlet opening, perpendicular to the main channel. | 05-13-2010 |
| 20130081408 | METHOD AND APPARATUS FOR COOLING GAS TURBINE ROTOR BLADES - An airfoil for a gas turbine engine includes a first sidewall and a second sidewall coupled together at a leading edge and a trailing edge, such that a cavity is defined therebetween. A plurality of cooling circuits are defined within the cavity. Each cooling circuit channels cooling fluid through at least one cooling chamber to facilitate cooling the airfoil. More specifically, a cascade impingement circuit, a down pass circuit, a flag tip circuit, and a trailing edge circuit are provided. The cascade impingement circuit includes a central chamber and a plurality of impingement chambers. | 04-04-2013 |
| 20130031914 | TWO STAGE SERIAL IMPINGEMENT COOLING FOR ISOGRID STRUCTURES - A system for cooling a wall ( | 02-07-2013 |
| 20100101233 | Cooling Temperature Ladder and Applications Thereof - A reverse flow heat exchanger is combined with a thermal energy sink to generate a temperature ladder. This system is used to cool a fluid more efficiently and/or to a lower temperature than would be possible without the reverse flow heat exchanger. The cooled fluid is optionally used to cool a sensor, a superconductor, a circuit, a cooling surface, or the like. The cooled fluid is optionally combined with a catalyst to remove unwanted constituents. | 04-29-2010 |
| 20090199568 | TRANSITION SCROLLS FOR USE IN TURBINE ENGINE ASSEMBLIES - An engine assembly includes a combustor having a combustion chamber in which an air and fuel mixture is combusted to produce combustion gases. The engine assembly further includes a transition scroll coupled to the combustor for receiving the combustion gases. The transition scroll includes an interior surface, an exterior surface, and effusion cooling holes for providing cooling air to the interior surface. The engine assembly further includes a turbine coupled to the transition scroll for receiving and extracting energy from the combustion gases. | 08-13-2009 |
| 20130067932 | COMBUSTION SECTIONS OF GAS TURBINE ENGINES WITH CONVECTION SHIELD ASSEMBLIES - A combustion section is provided for a gas turbine engine. The combustion section includes a first liner; a second liner forming a combustion chamber with the first liner, the combustion chamber configured to receive an air-fuel mixture for combustion therein; a first case circumscribing the first liner and forming a first plenum with the first liner; and a convection shield assembly positioned between the first liner and the first case. | 03-21-2013 |
| 20130067933 | GAS TURBINE - A gas turbine is provided and includes a compressor, which via an air intake inducts and compresses air; a combustion chamber, in which a fuel is combusted using the compressed air, producing a hot gas; and a turbine, equipped with turbine blades, in which the hot gas is expanded, performing work. A first device is provided in order to cool turbine blades with compressed cooling air. The first device includes at least one separate compressor stage which produces compressed cooling air independently of the compressor. | 03-21-2013 |
| 20090272124 | Cooling channel for cooling a hot gas guiding component - The invention relates to a cooling channel for a component conveying hot gas for the purposes of conveying a coolant along a direction of flow with a dowrnstream and an upstream side, with a plurality of inlet apertures for a coolant, with a number of inlet apertures that vary their configuration at least partly among themselves is arranged at least in one section of the cooling channel. As a result, the heat-transfer coefficient is substantially increased at points particularly requiring cooling and therefore the cooling is substantially improved. The cooling channel is characterized by a particularly low pressure loss. Furthermore, a combustion chamber with a cooling channel of this type is specified | 11-05-2009 |
| 20100122538 | METHODS, APPARATUS AND SYSTEMS CONCERNING THE CIRCUMFERENTIAL CLOCKING OF TURBINE AIRFOILS IN RELATION TO COMBUSTOR CANS AND THE FLOW OF COOLING AIR THROUGH THE TURBINE HOT GAS FLOWPATH - A method of operating a turbine engine, wherein the turbine engine includes a compressor, a combustor, a turbine, a plurality of successive axially stacked stages that include a row of circumferentially spaced stator blades and circumferentially spaced rotor blades, and a plurality of circumferentially spaced injection ports disposed upstream of a first row of stator blades in the turbine; the injection ports comprising a port through which cooling air is injected into the hot-gas path of the turbine, the method comprising: configuring the stator blades in the first row of stator blades such that the circumferential position of a leading edge of one of the stator blades is located within +/−15% pitch of the first row of stator blades of the circumferential location of the injection port midpoint of at least a plurality of the injection ports. | 05-20-2010 |
| 20090235671 | PARTIAL OXIDATION GAS TURBINE COOLING - A power generation system and method in which a fuel gas is introduced into a combustor and at least a portion of the fuel gas is combusted in the combustor, producing an exhaust gas having no appreciable available oxygen. The exhaust gas is introduced as a working fluid into a gas turbine, thereby generating power. Cooling of the power generation system is accomplished using a cooling fluid selected from the group consisting of synthesis gas, natural gas, natural gas/steam mixture, flue gas, flue gas/steam mixture, and mixtures thereof. | 09-24-2009 |
| 20110283713 | System For Cooling A Heat Exchanger On Board An Aircraft | 11-24-2011 |
| 20110296848 | FLUID TRANSFER ARRANGEMENT - A fluid transfer arrangement comprising a duct having a first end and a second end, a pulse generation mechanism located at the first end of the duct to direct fluid pulses towards the second end of the duct in use, and a baffle located at the second end of the duct that defines an aperture having sharp edges. The sharp edges generate ring vortex fluid flow from the aperture in use. Applications include impingement heating and cooling. | 12-08-2011 |
| 20110271689 | GAS TURBINE COOLING - In one embodiment, a compressor discharge casing of a gas turbine engine is designed to receive discharge air from a compressor and to direct a first portion of the discharge air into a combustor of the gas turbine engine and a second portion of the discharge air into a nozzle assembly of a gas turbine to cool components of the gas turbine. A heat transfer device is configured to receive a cooling fluid and to cool the second portion of the discharge air with the cooling fluid. | 11-10-2011 |
| 20100101234 | Evaporative Cooler and Use Thereof and Gas Turbine System Featuring an Evaporative Cooler - An evaporative cooler for cooling a gas stream, in particular an air stream, including a number of cooling elements located in a flow channel, is provided. A liquid, preferably water, is supplied by a feed device and will be vaporized or evaporated. In one aspect, the surface of at least one of the cooling elements has hydrophilic properties, at least in one sub-region designed to form a liquid film. | 04-29-2010 |
| 20110113790 | THERMAL MACHINE - A thermal machine including a wall defining a hot gas duct for transferring a hot gas stream and a cooling jacket disposed at a distance from the wall on an outside of the hot gas duct so as to define a cooling duct with an inlet and an outlet. The cooling duct is configured to conduct a cooling medium along an external face of the wall from the inlet to an outlet in a direction counter to a flow of hot gas in the hot gas duct. An impingement cooling plate is disposed at the inlet of the cooling duct and includes cooling baffle holes configured such that cooling medium entering the cooling duct through the cooling baffle holes flows in a direction perpendicular to the wall. The impingement cooling plate is positioned such that an inflow-side edge sealingly abuts the wall of the hot gas duct so as to reduce a transverse flow of the cooling medium through the cooling duct. | 05-19-2011 |
| 20110203294 | FUEL INJECTION SYSTEM FOR A COMBUSTION CHAMBER OF A TURBOMACHINE - A fuel injection system for an annular combustion chamber of a turbomachine, the system comprising support means for supporting and centering a fuel injector head, and a bowl arranged downstream from the support means and including at its downstream end an annular collar that extends radially outwards and that is cooled by air impacting against its upstream radial surface, and including on said surface means for disturbing the flow of cooling air and for increasing the heat exchange area between the air and the collar. | 08-25-2011 |
| 20090007570 | METHODS AND SYSTEMS FOR COOLING FLUID IN A TURBINE ENGINE - A method of assembling a turbine engine is provided. The method includes providing a heat exchanger having a curvilinear body. The method also includes coupling the heat exchanger to at least one of a fan casing and an engine casing of the turbine engine. The curvilinear body facilitates reducing pressure losses in airflow channeled into the heat exchanger. | 01-08-2009 |
| 20080250792 | APPARATUS AND METHOD OF OPERATING A GAS TURBINE ENGINE AT START-UP - A method of operating a gas turbine engine having a turbine and a compressor connected via a shaft, a main fuel supply line for supplying fuel to a combustor that is positioned to release expanding hot gases to the turbine, the engine further including a starter/generator connected to the shaft via a gearbox assembly, the method is characterised the step of during engine start up fuel is circulated in a re-circulating fuel circuit positioned on the main fuel supply line and which a first fuel/oil heat exchanger, for cooling the oil, and a fuel accumulator. | 10-16-2008 |
| 20100146987 | GAS TURBINE ENGINE - A combustor ( | 06-17-2010 |
| 20100146985 | High Temperature-Resistant Sealing Assembly, Especially for Gas Turbines - The invention relates to a high temperature-resistant sealing assembly comprising a sealing segment and a component border which is connected to the sealing segment. A flexible sealing element is provided in order to joining the sealing segment and the component border. The invention is characterized in that the flexible sealing element compensates both thermal expansions and relative movements of the component border and the components resting against the sealing segment. The flexible sealing element and the sealing segment are subject to little wear and have a long service life. The invention further relates to a combustion chamber that is equipped with a high temperature-resistant sealing assembly as well as a gas turbine encompassing such a combustion chamber. | 06-17-2010 |
| 20100186419 | HEAT TRANSFER AUGMENTATION IN A COMPACT HEAT EXCHANGER PEDESTAL ARRAY - A compact heat exchanger pedestal array for augmenting heat transfer in a machine is disclosed. The compact heat exchanger pedestal array includes a wall having first and second surfaces. The first surface faces a heated flow path and the second surface partially forms a flow path for cooling fluid. A plurality of pedestals extend from the second surface of the wall. At least one turbulator strip extends between adjacent pedestals. The turbulator strips and pedestals are operable for mixing the cooling fluid to increase heat transfer from the wall to the cooling fluid. | 07-29-2010 |
| 20090077979 | RESIDUAL STEAM REMOVAL MECHANISM AND RESIDUAL STEAM REMOVAL METHOD FOR STEAM COOLING PIPING OF GAS TURBINE - A combustor | 03-26-2009 |
| 20120192572 | GAS TURBINE ENGINE - One embodiment of the present invention is a unique gas turbine engine. Another embodiment of the present invention is a unique gas turbine engine bearing system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine engine bearing systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. | 08-02-2012 |
| 20100218509 | Combustion chamber - A combustion chamber is provided, including an inner casing with a sliding surface and an outer casing with a sliding wall portion. The sliding surface and the sliding wall portion are slidably attached to each other. A cooling hole is located in the sliding wall portion. The cooling hole is at least partially located in the sliding wall portion such that it opens due to a sliding movement of the sliding surface relative to the sliding wall portion when the inner casing thermally expands and/or closes due to a sliding movement of the sliding surface relative to the sliding wall portion when the inner casing thermally contracts. Moreover, a gas turbine including an inventive combustion chamber is disclosed. | 09-02-2010 |
| 20100071382 | Gas Turbine Transition Duct - A transition member between a combustion section and a turbine section in a gas turbine engine. The transition member includes a casing inner wall and a plurality of spanning members. The spanning members extend radially outwardly from a radially outer surface of the casing inner wall. Each of the spanning members included a slot formed therein. Each slot is in communication with a first aperture formed in the radially inner surface of the casing inner wall and a plurality of second apertures formed in an aft side of the spanning member for effecting a passage of the cooling fluid from a first cooling fluid channel to an inner volume defined within the radially inner surface of the casing inner wall. The slots include a component in the radial direction and a component in the axial direction such that the first aperture is not radially aligned with the second apertures. | 03-25-2010 |
| 20110232299 | IMPINGEMENT STRUCTURES FOR COOLING SYSTEMS - An impingement structure | 09-29-2011 |
| 20100146986 | Cavity ventilation - A ventilation arrangement for an annular cavity located about the core engine of a gas turbine, the annular cavity having an axis and an inner wall located about the axis and an outer wall radially spaced there-from, the outer wall being interposed between the cavity and an engine bypass duct. The ventilation arrangement comprises a plurality of ventilation ducts passing through said outer wall and providing for fluid communication between the bypass duct and the cavity and a flow diverter arranged at an oblique angle to a radial alignment with said axis so as to induce in the ventilation flow from said ducts a vortical flow component about the inner wall. The ducts may be shaped to provide the flow diverter. The internal geometry of the ducts may provide for a sudden expansion to control the exit profile of the flow into the cavity. | 06-17-2010 |
| 20100146988 | Gas turbine system - A gas turbine system, in particular for a utility power plant, includes a combustion chamber ( | 06-17-2010 |
| 20110079021 | APPARATUS AND METHOD FOR REMOVING HEAT FROM A GAS TURBINE - An apparatus for removing heat from a turbine includes a stator having a cavity and a first plenum and a second plenum inside the cavity. The second plenum is connected to the first plenum and surrounds the first plenum inside the cavity. A refrigerant flows through the first plenum and the second plenum to remove heat from the stator. A method for cooling a turbine includes forming a cavity in a component to be cooled, installing a first plenum inside the cavity, and installing a second plenum inside the cavity. The method further includes connecting the second plenum to the first plenum, surrounding the first plenum with the second plenum inside the cavity, and flowing a refrigerant through the first plenum and the second plenum to cool the component. | 04-07-2011 |
| 20090314006 | Gas turbine engine and integrated heat exchange system - One embodiment of a heat exchange system is disclosed having a heat exchange compressor and multiple evaporators capable of operating at different heat transfer requirements. The heat exchange compressor may be a single-stage or multi-stage compressor. In one form the evaporators return working fluid in separate streams to the heat exchange compressor. The heat exchange compressor may be integrated with a gas turbine engine and includes a number of inlets that correspond to a number of separate evaporators. Each inlet can be configured to receive working fluid at different locations within a pressure and velocity flow field created in the compressor. The heat exchange compressor may be driven by a shaft of the gas turbine engine and may be positioned at a variety of locations. | 12-24-2009 |
| 20110100020 | APPARATUS AND METHOD FOR TURBINE ENGINE COOLING - A turbine engine comprises a turbine housing, a turbine disposed in the turbine housing that is configured to receive hot combustion gas, a turbine component subject to thermal energy from the hot combustion gas and a cooling system disposed externally of the turbine housing and having a cooling medium disposed therein. A heat pipe has a high temperature end in communication with the turbine component and a low temperature end extending out of the turbine housing in communication with the cooling medium in the cooling system for transferring the thermal energy from the component to the cooling medium. | 05-05-2011 |
| 20110072832 | VENTILATION FOR A TURBINE WHEEL IN A TURBINE ENGINE - Turbine engine, including a final centrifugal compressor stage associated with a diffuser for supplying air to a combustion chamber, and ventilation means for ventilating a high-pressure turbine wheel, including injection means for injecting air onto the wheel and take-up means for taking up a flow for cooling the impeller of the compressor, wherein these take-up means comprise a labyrinth seal mounted at the outlet of the injection means and the air outlet orifices installed between the injection means and said labyrinth seal and leading upstream of the turbine wheel. | 03-31-2011 |
| 20110247346 | COOLING FLUID METERING STRUCTURE IN A GAS TURBINE ENGINE - A gas turbine engine includes a supply of cooling fluid, a rotatable shaft, structure defining at least one bypass passage in fluid communication with the supply of cooling fluid for supplying cooling fluid from the supply of cooling fluid, and metering structure located at an outlet of the at least one bypass passage. The metering structure includes at least one flow passageway extending therethrough at an angle to a central axis of the engine for permitting cooling fluid in the bypass passage to pass into a turbine rim cavity. The cooling fluid flowing out of the flow passageway has a velocity component in a direction tangential to the circumferential direction in the same direction as a rotation direction of the shaft. | 10-13-2011 |
| 20110247345 | COOLING FLUID PRE-SWIRL ASSEMBLY FOR A GAS TURBINE ENGINE - A gas turbine engine includes a pre-swirl structure. Inner and outer wall structures of the pre-swirl structure define a flow passage in which swirl members are located. The swirl members include a leading edge and a circumferentially offset trailing edge. Cooling fluid exits the flow passage with a velocity component in a direction tangential to the circumferential direction, wherein a swirl ratio defined as the velocity component in the direction tangential to the circumferential direction of the cooling fluid to a velocity component of a rotating shaft in the direction tangential to the circumferential direction is greater than one as the cooling fluid exits the flow passage outlet, and the swirl ratio is about one as the cooling fluid enters at least one bore formed in a blade disc structure. An annular cavity extends between the flow passage and the at least one bore formed in the blade disc structure. | 10-13-2011 |
| 20110048030 | IMPINGEMENT COOLED TRANSITION PIECE AFT FRAME - An aft frame of a turbine engine transition piece body is provided and includes an annular body disposed within a first annular space defined between an impingement sleeve and a compressor discharge casing and aft of a second annular space defined between the transition piece body and the impingement sleeve and including a main portion with a first surface facing the first annular space and a second surface facing the forward annular space. The main portion has an impingement hole extending therethrough from an inlet at the first surface of the annular body to an outlet at the second surface of the annular body to define a fluid path along which the first and second annular spaces communicate with one another. | 03-03-2011 |
| 20110247347 | PARTICLE SEPARATOR IN A GAS TURBINE ENGINE - A gas turbine engine includes a supply of cooling fluid, a rotatable shaft, blade disc structure coupled to the shaft and having at least one bore for receiving cooling fluid, and a particle separator. The particle separator includes particle deflecting structure upstream from the blade disc structure, and a particle collection chamber. The particle deflecting structure deflects solid particles from the cooling fluid prior to the cooling fluid entering the at least one bore in the blade disc structure. The particle collection chamber is upstream from the particle deflecting structure and receives the solid particles deflected from the cooling fluid by the particle deflecting structure. The solid particles deflected by the particle deflecting structure flow upstream from the particle deflecting structure to the particle collection chamber. | 10-13-2011 |
| 20100263388 | Vapor cooled static turbine hardware - A cooling system for a gas turbine engine includes a non-rotating component extending into an engine flowpath, a vapor cooling assembly configured to transport thermal energy from a vaporization section to a condenser section through cyclical evaporation and condensation of a working medium sealed within the vapor cooling assembly, wherein the vaporization section is located at least partially within the non-rotating component, and wherein the condenser section is located outside the non-rotating component and away from the engine flowpath. | 10-21-2010 |
| 20080314047 | COOLING SYSTEMS FOR USE ON AIRCRAFT - Cooling systems for an aircraft are provided. In an embodiment, a system includes an engine nacelle, an engine, a bypass duct, and a heat exchanger. The engine nacelle includes an airflow inlet. The engine is housed in the engine nacelle in flow communication with the airflow inlet. The bypass duct extends between the engine nacelle and the engine is in flow communication with the airflow inlet. The bypass duct includes an outer wall and an opening formed therein. The heat exchanger is integrated with the engine and is disposed over the opening of the bypass duct outer wall between the bypass duct outer wall and the engine nacelle. | 12-25-2008 |
| 20110162387 | TURBINE COOLING SYSTEM - A cooling system is provided for cooling a turbine of a gas turbine engine. The system has first and second flow paths for guiding cooling air received from the compressor of the engine. The routes of both flow paths bypass the combustor of the engine. The system also has a preswirler for receiving the cooling air at the ends of the two flow paths, swirling the cooling air tangentially to the engine axis, and delivering the swirled cooling air to a rotor of the turbine. The first flow path is routed through a heat exchanger which cools the cooling air guided by the first flow path relative to the cooling air guided by the second flow path. | 07-07-2011 |
| 20110016884 | COOLING PASSAGE COVER, MANUFACTURING METHOD OF THE COVER, AND GAS TURBINE - To provide a cover of a cooling passage that forms a cooling passage for supplying cooling air to a turbine rotor blade at the last stage via inside of a disk of a turbine, and the cover comprises: a cylindrical cover portion that covers a cavity provided in a annular pattern in an outer circumference of the disk in a mode where a first passage opened from inside of the disk to the cavity and a second passage opened from a cooling passage of the turbine rotor blade at the last stage to the cavity are connected to each other; and a flexible portion that is formed integrally with the cover portion and allows flexure in an axial direction of the turbine. | 01-27-2011 |
| 20110259017 | HOT GAS PATH COMPONENT COOLING SYSTEM - A cooling system for a hot gas path component is disclosed. The cooling system may include a component layer and a cover layer. The component layer may include a first inner surface and a second outer surface. The second outer surface may define a plurality of channels. The component layer may further define a plurality of passages extending generally between the first inner surface and the second outer surface. Each of the plurality of channels may be fluidly connected to at least one of the plurality of passages. The cover layer may be situated adjacent the second outer surface of the component layer. The plurality of passages may be configured to flow a cooling medium to the plurality of channels and provide impingement cooling to the cover layer. The plurality of channels may be configured to flow cooling medium therethrough, cooling the cover layer. | 10-27-2011 |
| 20100018219 | DEVICE FOR OPTIMIZING COOLING IN GAS TURBINES | 01-28-2010 |
| 20120144843 | GAS TURBINE ENGINE AND COOLING SYSTEM - One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique cooling system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for cooling one or more objects of cooling. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith. | 06-14-2012 |
| 20120000205 | ADAPTIVE POWER AND THERMAL MANAGEMENT SYSTEM - An aircraft adaptive power thermal management system for cooling one or more aircraft components includes an air cycle system, a vapor cycle system, and a fuel recirculation loop operably disposed therebetween. An air cycle system heat exchanger is between the air cycle system and the fuel recirculation loop, a vapor cycle system heat exchanger is between the vapor cycle system and the fuel recirculation loop, and one or more aircraft fuel tanks are in the fuel recirculation loop. An intercooler including a duct heat exchanger in an aircraft gas turbine engine FLADE duct may be in the air cycle system. The system is operable for providing on-demand cooling for one or more of the aircraft components by increasing heat sink capacity of the fuel tanks. | 01-05-2012 |
| 20120060511 | APPARATUS AND METHOD FOR COOLING A COMBUSTOR CAP - A combustor includes an end cap having a perforated downstream plate and a combustion chamber downstream of the downstream plate. A plenum is in fluid communication with the downstream plate and supplies a cooling medium to the combustion chamber through the perforations in the downstream plate. A method for cooling a combustor includes flowing a cooling medium into a combustor end cap and impinging the cooling medium on a downstream plate in the combustor end cap. The method further includes flowing the cooling medium into a combustion chamber through perforations in the downstream plate. | 03-15-2012 |
| 20120011857 | High-Flow-Capacity Centrifugal Hydrogen Gas Compression Systems, Methods and Components Therefor - Hydrogen gas compression systems that each include a multistage centrifugal compressor in which each stage has an inlet-to-outlet pressure rise ratio of about 1.20 or greater. In one embodiment, the multistage compressor includes six high-speed centrifugal compressors driven at a speed of about 60,000 rpm. The compressor has an output of more than 200,000 kg/day at a pressure of more than 1,000 psig. The compressors for the compression stages are distributed on both sides of a common gear-box, which has gearing that allows axial thrusts from the compressors to be handled effectively. Each stage's compressor has a unique impeller, which is secured to a support shaft using a tension-rod-based attachment system. In another embodiment, the multistage compressor is driven by a combustion turbine and one or more intercoolers are provided between compression stages. Each intercooler is cooled by coolant from an absorption chiller utilizing exhaust gas from the combustion turbine. | 01-19-2012 |
| 20120017605 | HEAT TRANSFER AUGMENTED FLUID FLOW SURFACES - A heat transfer augmented channel wall includes a bulk portion, a wall surface and a plurality of multi-portion indented features extending from the wall surface into the bulk portion. The multi-portion indented features include a first indented portion and a second indented portion that are divided by a ridge which disrupts fluid flow between first and second indented portions. The ridge has a height that is less than a maximum depth of the multi-portion indented features. | 01-26-2012 |
| 20120167595 | GAS TURBINE ENGINE WITH SECONDARY AIR FLOW CIRCUIT - One embodiment of the present invention is a unique gas turbine engine. Another embodiment of the present invention is a gas turbine engine having a unique secondary air flow circuit. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and secondary air flow circuits. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. | 07-05-2012 |
| 20120255311 | COOLING STRUCTURE, GAS TURBINE COMBUSTOR AND MANUFACTURING METHOD OF COOLING STRUCTURE - A cooling structure, a gas turbine combustor, and a method of manufacturing the cooling structure attain a high cooling efficiency without increasing manufacturing cost. The cooling structure includes a first member as a cooling object having a first plane, and a second member arranged above the first plane and having an opposing second plane such that a passage is formed between the first plane and the second plane for a cooling medium to flow. The first member has a plurality of prominences each of which extends upwardly from the first plane, and extends to be inclined along a direction in which the cooling medium flows. A clearance between the second plane and a tip of each prominence is set such that a heat transfer rate between the cooling medium and the first member becomes larger than that when each prominence extends vertically upward from the first plane. | 10-11-2012 |
| 20100229570 | BURNER FOR A GAS TURBINE AND METHOD FOR LOCALLY COOLING A HOT GASES FLOW PASSING THROUGH A BURNER - The burner ( | 09-16-2010 |
| 20120085104 | TURBINE ENGINE INCLUDING AN IMPROVED MEANS FOR ADJUSTING THE FLOW RATE OF A COOLING AIR FLOW SAMPLED AT THE OUTPUT OF A HIGH-PRESSURE COMPRESSOR - A turbine engine including a channel for injecting a cooling air flow of a high-pressure turbine disk, opening into a cavity that is substantially isolated, upstream, from a cavity in which an air flow sampled at the output of a high-pressure compressor circulates, by a first labyrinth seal, and downstream, from a cavity communicating with the primary flow of the turbine engine, by a second labyrinth seal. The turbine engine includes channels communicating with the injection channel and opening through a static part of the first labyrinth seal between two lips of that seal, so as to allow an air flow coming from the injection channel to be injected between the lips. | 04-12-2012 |
