Entries |
Document | Title | Date |
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 |
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 |
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 |
20090077979 | RESIDUAL STEAM REMOVAL MECHANISM AND RESIDUAL STEAM REMOVAL METHOD FOR STEAM COOLING PIPING OF GAS TURBINE - A combustor | 03-26-2009 |
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 |
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 |
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 |
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 |
20100018219 | DEVICE FOR OPTIMIZING COOLING IN GAS TURBINES | 01-28-2010 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
20100146987 | GAS TURBINE ENGINE - A combustor ( | 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 |
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 |
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 |
20100229570 | BURNER FOR A GAS TURBINE AND METHOD FOR LOCALLY COOLING A HOT GASES FLOW PASSING THROUGH A BURNER - The burner ( | 09-16-2010 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
20110232299 | IMPINGEMENT STRUCTURES FOR COOLING SYSTEMS - An impingement structure | 09-29-2011 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
20130031914 | TWO STAGE SERIAL IMPINGEMENT COOLING FOR ISOGRID STRUCTURES - A system for cooling a wall ( | 02-07-2013 |
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 |
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 |
20130104567 | METHOD AND APPARATUS FOR COOLING GAS TURBINE ROTOR BLADES | 05-02-2013 |
20130192267 | INTERNALLY COOLED SPOKE - A turbine engine includes a compressor section, a combustor section in fluid communication with the compressor section, a high pressure turbine in fluid communication with the combustor, a low pressure turbine in fluid communication with the high pressure turbine, and a mid turbine frame located axially between the high pressure turbine and the low pressure turbine. The mid turbine frame includes an outer frame case, an inner frame case, and a plurality of hollow spokes that distribute loads from the inner frame case to the outer frame case. The spokes are hollow to allow cooling airflow to be supplied through the spokes to the inner frame case. | 08-01-2013 |
20130192268 | INTERNALLY COOLED SPOKE - A turbine engine includes a compressor section, a combustor section in fluid communication with the compressor section, a high pressure turbine in fluid communication with the combustor, a low pressure turbine in fluid communication with the high pressure turbine, and a mid turbine frame located axially between the high pressure turbine and the low pressure turbine. The mid turbine frame includes an outer frame case, an inner frame case, and a plurality of hollow spokes that distribute loads from the inner frame case to the outer frame case. The spokes are hollow to allow cooling airflow to be supplied through the spokes to the inner frame case. | 08-01-2013 |
20130205801 | MULTI-LOBED COOLING HOLE AND METHOD OF MANUFACTURE - A gas turbine engine component includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet at the first wall surface, an outlet at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally and laterally from the metering section, a second lobe adjacent the first lobe and diverging longitudinally from the metering section, a third lobe adjacent the second lobe and diverging longitudinally and laterally from the metering section, and a transition region having an end adjacent the outlet and a portion that extends between the lobes and the outlet. The first and third lobes each include a curved outer portion. | 08-15-2013 |
20130205802 | COOLING HOLE WITH CRENELLATION FEATURES - A gas turbine engine component includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface and a diffusing section in communication with the inlet and extending to the outlet. The diffusing section includes a plurality of crenellation features that encourage lateral spreading of cooling air flowing through the cooling hole. | 08-15-2013 |
20130205803 | MULTI-LOBED COOLING HOLE ARRAY - A gas turbine engine component includes a wall having first and second wall surfaces and first and second cooling holes extending through the wall. The first and second cooling holes each include an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. Each diffusing section includes first and second lobes, each lobe diverging longitudinally and laterally from the metering section. The outlets of each cooling hole include first and second lateral ends and a trailing edge. One of the lateral ends of the outlet of the first cooling hole and one of the lateral ends of the outlet of the second cooling hole meet upstream of the trailing edge of the first cooling hole and the trailing edge of the second cooling hole. | 08-15-2013 |
20130227964 | TRANSITION PIECE AFT FRAME ASSEMBLY HAVING A HEAT SHIELD - A transition piece aft frame assembly is provided, and includes a transition piece aft frame and a heat shield. The transition piece aft frame has an aft face. At least a portion of the aft face is exposed to an exhaust gas stream. The heat shield is connected to the transition piece aft frame. The heat shield is oriented to generally deflect the exhaust gas stream away from the aft face of the transition piece aft frame. | 09-05-2013 |
20130232991 | AIRFOIL WITH IMPROVED INTERNAL COOLING CHANNEL PEDESTALS - An airfoil for a turbine engine, the airfoil including a first side wall, a second side wall spaced apart from the first side wall, and an internal cooling channel formed between the first side wall and the second side wall. The internal cooling channel includes at least one pedestal having a first pedestal end connected to the first side wall and a second pedestal end connected to the second side wall. The internal cooling channel also includes a first fillet disposed around the periphery of the first pedestal end between the first side wall and the first pedestal end; and a second fillet disposed around the periphery of the second pedestal end between the second side wall and the second pedestal end. At least one of the first fillet and the second fillet includes a profile that is non-uniform around the periphery of the corresponding pedestal end. | 09-12-2013 |
20130239588 | PUMP SYSTEM FOR TMS AOC REDUCTION - An engine includes a duct containing a flow of cool air and a pump system having an impeller with an inlet for receiving air from the duct and an outlet for discharging air into a discharge manifold. The discharge manifold containing at least one heat exchanger which forms part of a thermal management system. | 09-19-2013 |
20130247587 | BI-METALLIC ACTUATOR FOR SELECTIVELY CONTROLLING AIR FLOW BETWEEN PLENA IN A GAS TURBINE ENGINE - A system for selectively supplying air between separate plena of a gas turbine engine includes a gas turbine engine, a door, and a bi-metallic door actuator. The gas turbine engine comprises at least a first plenum and a second plenum, and has an opening between the first plenum and the second plenum. The is door mounted in the gas turbine engine and is movable between a closed position, in which air is prevented from flowing through the opening, and an open position, in which air may flow though the opening. The bi-metallic door actuator is coupled to the door and is configured to selectively move the door between the closed position and the open position. | 09-26-2013 |
20130255278 | EFFUSION COOLED SHROUD SEGMENT WITH AN ABRADABLE SYSTEM - A turbine casing assembly, comprising an annular seal segment assembly for surrounding the turbine adjacent to the turbine blades. An abradable coating is provided on the inboard surface of the seal segments of the assembly and one or more coolant ducts extend from the outboard surface of the seal segment assembly through respective seal segments and the abradable coating, for carrying a coolant towards the blade tips. One or more annular grooves are formed in the inboard surface of the abradable coating, the or each coolant duct opening into one of the one or more annular grooves. | 10-03-2013 |
20130276460 | AIRFOIL HAVING MINIMUM DISTANCE RIBS - An airfoil includes an airfoil body that defines a longitudinal axis. The airfoil body includes a leading edge and a trailing edge and a first side wall and a second side wall that is spaced apart from the first side wall to define a camber line there between. The first side wall and the second side wall join the leading edge and the trailing edge and at least partially define a cavity in the airfoil body. Multiple ribs extend longitudinally in the cavity and are laterally spaced apart from each other relative to the longitudinal axis. In at least one plane that is perpendicular to the longitudinal axis, each of the ribs connects the first side wall and the second side wall along respective minimum distance directions that are perpendicular to the camber line. At least two of the respective minimum distance directions are non-parallel. | 10-24-2013 |
20130276461 | AIRFOIL HAVING INTERNAL LATTICE NETWORK - An airfoil includes an airfoil body that defines a longitudinal axis. The airfoil body includes a leading edge and a trailing edge and a first side wall and a second side wall that is spaced apart from the first side wall. The first side wall and the second side wall join the leading edge and the trailing edge and at least partially define a cavity in the airfoil body. A lattice network connects the first side and the second side. The lattice network includes at least one enlarged node spaced apart from the first side wall and the second side wall and ribs that extend from the at least one enlarged node. Each of the ribs connects to one of the first side wall and the second side wall. | 10-24-2013 |
20130305740 | METHOD AND APPARATUS FOR GAS TURBINE ENGINE TEMPERATURE MANAGEMENT - A turbine engine includes a turbine, a compressor for compressing air and a combustor for receiving the compressed air through an inlet passage and operable to burn fuel therewith to deliver hot exhaust gas to the turbine. Also included is a wheel space defined proximate to the combustor. Further included is a cooling air passage extending between the compressor and the wheel space. Yet further included is a valve assembly having a valve member disposed in the cooling air passage and operable to admit a cooling air to the wheel space in response to a condition therein. | 11-21-2013 |
20130312425 | PASSIVE THERMOSTATIC VALVE - An airfoil includes an airfoil body having an internal cavity. A thermostatic valve is located at least partially within the internal cavity. The thermostatic valve is configured to passively control fluid flow into the internal cavity in response to a temperature within the internal cavity. | 11-28-2013 |
20130327061 | TURBOMACHINE BUCKET ASSEMBLY AND METHOD OF COOLING A TURBOMACHINE BUCKET ASSEMBLY - A turbomachine bucket assembly includes a rotor member including a body having a center portion and an outer edge portion joined by a web. The rotor member includes one or more cooling fluid conduit having a dimension, and an inlet arranged at the outer edge. A plurality of blades are provided on the rotor member and mechanically linked to the outer edge. Each of the plurality of blades includes an internal cooling passage that is fluidly connected to the one or more cooling fluid conduits. A cooling fluid control element is provided at each of the one or more cooling fluid conduits. The cooling fluid control element is configured and disposed to adjust the dimension of the one or more cooling fluid conduits to alter fluid flow into the plurality of blades. | 12-12-2013 |
20130333393 | GAS TURBINE CONTROL SYSTEMS AND METHODS - Methods and systems for controlling a gas turbine system are provided herein. In one embodiment, a method includes the steps of receiving at least one parameter of turbine inlet air and determining, based on the at least one parameter, an expected condensation level at an intercooler disposed downstream of an inlet air chilling system and in-line between a low pressure compressor and a high pressure compressor. The method further includes determining a desired temperature of the turbine inlet air corresponding to substantially no expected condensation at the intercooler and controlling the inlet air chilling system to chill the turbine inlet air to the desired temperature. | 12-19-2013 |
20140000282 | TURBINE BLADE PLATFORM WITH U-CHANNEL COOLING HOLES | 01-02-2014 |
20140000283 | COVER PLATE FOR A COMPONENT OF A GAS TURBINE ENGINE | 01-02-2014 |
20140000284 | COOLING APPARATUS FOR A MID-TURBINE FRAME | 01-02-2014 |
20140000285 | GAS TURBINE ENGINE TURBINE VANE PLATFORM CORE | 01-02-2014 |
20140000286 | GAS TURBINE ENGINE TURBINE VANE AIRFOIL PROFILE | 01-02-2014 |
20140000287 | GAS TURBINE ENGINE TURBINE VANE AIRFOIL PROFILE | 01-02-2014 |
20140060084 | GAS TURBINE ENGINE AIRFOIL COOLING CIRCUIT ARRANGEMENT - A component for a gas turbine engine according to an exemplary aspect of the present disclosure including, among other things, an airfoil that extends between a leading edge and a trailing edge and a cooling circuit disposed inside of the airfoil. The cooling circuit includes at least one core cavity that extends inside of the airfoil, a baffle received within the at least one core cavity, a plurality of pedestals positioned adjacent to the at least one core cavity and a first plurality of axial ribs positioned between the plurality of pedestals and the trailing edge of the airfoil. | 03-06-2014 |
20140083114 | TURBINE BLADE ROOT PROFILE - A turbine blade for a gas turbine engine includes an airfoil that extends in a first radial direction from a platform. A root extends from the platform in a second radial direction and has opposing lateral sides that provide a firtree-shaped contour. The contour includes first, second and third lobes on each of the lateral sides and that tapers relative to the radial direction away from the platform. The first, second and third lobes each provide contact surfaces arranged at about 45° relative to the radial direction. A contact plane on each lateral side at an angle of about 11° relative to the radial direction defining a contact point on each of the contact surfaces. The first, second and third lobes each include first, second and third grooves that are substantially aligned with one another along an offset plane spaced a uniform offset distance from the contact plane. | 03-27-2014 |
20140083115 | ARTICLE WITH DIELECTRIC MIRROR COATING SYSTEM - An article includes a substrate having a surface arranged to receive radiation. A thermal barrier coating system is disposed on the surface and includes a ceramic-based coating. A dielectric mirror coating system is disposed on the ceramic-based coating such that the ceramic-based coating is located between the dielectric mirror coating system and the substrate. The dielectric mirror coating system includes a plurality of higher refractive index layers interleaved in an alternating stacked arrangement with a plurality of lower refractive index layers. Each of the plurality of higher refractive index layers and the plurality of lower refractive index layers has a thickness that is about ¼ of a preselected wavelength of the radiation. | 03-27-2014 |
20140083116 | GAS TURBINE ENGINE COMPONENTS WITH BLADE TIP COOLING - A turbine rotor blade for a turbine section of an engine is provided. The rotor blade includes a platform and an airfoil extending from the platform into a mainstream gas path of the turbine section. The airfoil includes a pressure side wall, a suction side wall joined to the pressure side wall at a leading edge and a trailing edge, and a tip cap extending between the suction side wall and the pressure side wall. The rotor blade further includes an internal cooling circuit having a tip cap passage configured to deliver cooling air to the tip cap and a flow accelerator positioned within the tip cap passage of the internal cooling circuit. | 03-27-2014 |
20140090402 | Combustor Bulkhead Assembly - A heat shield is disclosed. The heat shield may comprise a body having a back surface and an opposite front surface, wherein an opening in the body communicates through the front and back surfaces. The heat shield may further comprise at least one radial rail disposed on the back surface and extending radially outward from the opening for directing cooling air flow. | 04-03-2014 |
20140096538 | PLATFORM COOLING OF A TURBINE BLADE ASSEMBLY - A turbine blade generally includes a platform having a pressure side, a suction side, a leading edge, a trailing edge, a pressure side slash face and a suction side slash face. A platform cooling circuit extends within the platform. The platform cooling circuit may extend from the suction side of the platform to the pressure side of the platform. The platform cooling circuit generally defines a fluid flow path that directs a cooling medium from the platform suction side to the platform pressure side. | 04-10-2014 |
20140116066 | COMBUSTOR CAP ASSEMBLY - A combustor generally includes a shroud that that defines at least one inlet passage extends circumferentially inside the combustor. A first plate extends radially inside the shroud downstream from the inlet passage. The first plate defines at least one inlet port, at least one outlet port and at least partially defines at least one fuel nozzle passage. The shroud at least partially surrounds a sleeve that extends around the fuel nozzle passage. A tube at least partially surrounded by the sleeve may extend through the fuel nozzle passage. The tube, the sleeve, and the first plate may at least partially define an outlet passage. A first fluid flow path generally extends from the at inlet passage to the inlet port, and a second fluid flow path extends generally from the outlet port to the outlet passage. | 05-01-2014 |
20140130514 | TURBINE NOZZLE HAVING NON-LINEAR COOLING CONDUIT - A turbine nozzle having a non-linear cooling conduit is disclosed. In one embodiment, a turbine nozzle includes: an airfoil, at least one endwall adjacent the airfoil, and a fillet region connecting the airfoil and the at least one endwall, the fillet region including an outer surface. The turbine nozzle also includes a non-linear cooling conduit located within the fillet region and adjacent the outer surface of the fillet region, the non-linear cooling conduit allows fluid flow through the fillet region. The non-linear cooling conduit spans substantially along an axial length of the airfoil between a leading edge of the airfoil and a trailing edge of the airfoil. Additionally, the non-linear cooling conduit includes an arc profile substantially similar to an arc profile of the airfoil. | 05-15-2014 |
20140150455 | COATED ARTICLE - The present invention is a coated article containing at least two adjacent cooling holes that are substantially uncoated. In an embodiment, the coated article includes a base material, an outer surface and a coating on a portion of the outer surface. Here, the adjacent cooling holes are arranged along an axis on the outer surface and extend within the base material at an angle to the outer surface. The coating may include at least one unfinished edge oriented substantially parallel to the axis and set off at an oblique angle and a distance from the adjacent cooling holes. | 06-05-2014 |
20140157792 | SYSTEM AND METHOD FOR REMOVING HEAT FROM A TURBINE - A system for removing heat from a turbine includes a component in the turbine having a supply plenum and a return plenum therein. A substrate that defines a shape of the component has an inner surface and an outer surface. A coating applied to the outer surface of the substrate has an interior surface facing the outer surface of the substrate and an exterior surface opposed to the interior surface. A first fluid channel is between the outer surface of the substrate and the exterior surface of the coating. A first fluid path is from the supply plenum, through the substrate, and into the first fluid channel, and a second fluid path is from the first fluid channel, through the substrate, and into the return plenum. | 06-12-2014 |
20140165593 | GAS TURBINE ENGINE TURBINE BLADE LEADING EDGE TIP TRENCH COOLING - An airfoil for a gas turbine engine includes pressure and suction walls spaced apart from one another and joined at leading and trailing edges to provide an airfoil having an exterior surface that extends in a radial direction to a tip. A tip trench is provided in the tip and wrapping at least a portion of the airfoil from the pressure side wall around the leading edge to the suction side wall. The tip trench is provided by a recess. | 06-19-2014 |
20140208771 | GAS TURBINE ENGINE COMPONENT COOLING ARRANGEMENT - A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a body portion having an exterior surface and an internal surface. A cavity is disposed inside of the body portion. A cooling hole extends between the exterior surface and the internal surface and includes a metering section having an outlet and an inlet. The inlet is shaped dissimilar to the outlet. | 07-31-2014 |
20140216056 | HEAT EXCHANGE MODULE FOR A TURBINE ENGINE - A heat exchange module is provided for a turbine engine. The heat exchange module includes a duct and a plurality of heat exchangers. The duct includes a flowpath defined radially between a plurality of concentric duct walls. The flowpath extends along an axial centerline through the duct between a first duct end and a second duct end. The heat exchangers are located within the flowpath, and arranged circumferentially around the centerline. | 08-07-2014 |
20140250917 | COMBUSTION CHAMBER HEAT SHIELD AND SEAL ASSEMBLY AND A METHOD OF MANUFACTURING A COMBUSTION CHAMBER HEAT SHIELD AND SEAL ASSEMBLY - A gas turbine engine combustion chamber heat shield and seal assembly comprises a heat shield and a seal. The heat shield has an aperture and the seal is located in the aperture in the heat shield. The seal comprises an annular member having an upstream end, a middle and a downstream end. The upstream end of the seal has a diameter greater than the diameter of the aperture in the heat shield, the middle has a diameter less than the diameter of the aperture in the heat shield and the downstream end of the seal has a diameter greater than the diameter of the aperture in the heat shield. | 09-11-2014 |
20140260327 | COOLED ARTICLE - The present invention is an article containing internal cooling channels located near at least one surface. In an embodiment, the cooled article includes a base material, a first layer, and a second layer. Here, the first layer is bonded to the base material and the second layer is bonded to the first layer, wherein at least one closed cooling channel is disposed within a portion of the first layer and a portion of the second layer. | 09-18-2014 |
20140290272 | GAS TURBINE ENGINE COOLING ARRANGEMENT - A gas turbine engine comprises a compressor, a combustion chamber, an outer casing, an inner casing and a cooling arrangement. The outer casing surrounds the compressor and the combustion chamber and the combustion chamber has turbine nozzle guide vanes. The compressor has load carrying outlet guide vanes connected to the outer casing and the inner casing. The turbine nozzle guide vanes connect the outer casing and the inner casing. The cooling arrangement comprises a cooling air duct located between the compressor and the combustion chamber. The compressor outlet guide vanes carry at least one aerodynamic fairing. A support structure supports the cooling air duct from the inner casing at two spaced positions and the support structure forms a chamber with the inner casing. The support structure comprises at least one hollow duct and each hollow duct locates behind a respective one of the aerodynamic fairings. | 10-02-2014 |
20140311163 | METHOD OF MANUFACTURING A TURBOMACHINE COMPONENT, AN AIRFOIL AND A GAS TURBINE ENGINE - One embodiment of the present invention is a unique method of manufacturing a component for a turbomachine, such as an airfoil. Another embodiment is a unique airfoil. Yet another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for cooled gas turbine engine components. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. | 10-23-2014 |
20140311164 | GAS TURBINE ENGINE AND TURBINE BLADE - One embodiment of the present invention is a unique turbine blade for a gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and turbine blades 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. | 10-23-2014 |
20140331687 | High Efficiency Power Production Methods, Assemblies, and Systems - The present disclosure provides methods, assemblies, and systems for power production that can allow for increased efficiency and lower cost components arising from the control, reduction, or elimination of turbine blade mechanical erosion by particulates or chemical erosion by gases in a combustion product flow. The methods, assemblies, and systems can include the use of turbine blades that operate with a blade velocity that is significantly reduced in relation to conventional turbines used in typical power production systems. The methods and systems also can make use of a recycled circulating fluid for transpiration protection of the turbine and/or other components. Further, recycled circulating fluid may be employed to provide cleaning materials to the turbine. | 11-13-2014 |
20140338364 | TURBINE ROTOR BLADE FOR A TURBINE SECTION OF A GAS TURBINE - A turbine rotor blade includes a mounting portion that partially defines a cooling circuit within the turbine rotor blade and an airfoil portion that extends radially outward from the mounting portion. The airfoil portion further defines the cooling circuit. The turbine rotor blade further includes a platform portion that is disposed radially between the mounting portion and the airfoil. The platform portion includes a bottom wall, a top wall, a forward wall, an aft wall and a pair of opposing side walls. A cooling plenum that at least partially defines the cooling circuit is defined within the platform portion. The cooling plenum is at least partially defined between the forward wall, the aft wall and between the pair of opposing side walls. | 11-20-2014 |
20140352324 | DUAL PRESSURE REGULATOR SHUT OFF VALVE APPARATUS - A pre-cooler system is provided and includes first and second pre-coolers, each of which is sized to handle demands of one downstream flow system, a piping system by which the first and second pre-coolers are receptive of compressed air from first and second turbine engines, respectively, and by which the first and second pre-coolers are both coupled to first and second downstream flow systems that are each configured to apply the demands of one downstream flow system to the first and second pre-coolers, a first pair of dual pressure regulator shut off valves (PRSOVs) disposed in parallel with each other and between the first turbine engine and the first downstream flow system, the first pair of dual PRSOVs being arranged in series with the first pre-cooler and a second pair of dual PRSOVs disposed in parallel with each other and between the second turbine engine and the second downstream flow system, the second pair of dual PRSOVs being arranged in series with the second pre-cooler. | 12-04-2014 |
20140366556 | GAS TURBINE ENGINE VANE-TO-TRANSITION DUCT SEAL - A vane seal assembly for a gas turbine engine comprises of a case including a first connector. A notch in the case adjoins the groove. A vane having a second connector mates with the first connector. A seal assembly is provided between the vane and the case to provide a sealed cavity adjoining the notch. | 12-18-2014 |
20150007581 | SHROUD BLOCK SEGMENT FOR A GAS TURBINE - A shroud block segment for a gas turbine includes a main body having a leading portion, a trailing portion, a first side portion and an opposing second side portion that extend axially between the leading portion and the trailing portion. The main body further includes an arcuate combustion gas side, an opposing back side and a cooling chamber defined in the back side. A cooling plenum and an exhaust passage are defined within the main body where the exhaust passage provides for fluid communication out of the cooling plenum. An insert opening extends within the main body through the back side towards the cooling plenum. A cooling flow insert is disposed within the insert opening. The cooling flow insert comprises a plurality of cooling flow passages that provide for fluid communication between the cooling chamber and the cooling plenum. | 01-08-2015 |
20150013345 | GAS TURBINE SHROUD COOLING - A shroud segment for a casing of gas turbine includes a body configured for attachment to the casing proximate a localized critical process location within the casing. The body has a leading edge, a trailing edge, and two side edges. The critical process location is located between the leading edge and the trailing edge when the body is attached to the casing. A cooling passage is defined in the body along one of the side edges with one of an inlet or an outlet proximate the critical process location. The cooling passage is configured large enough to cool the one side edge adjacent the cooling passage to a desired level during operation of the gas turbine. The critical process locations may be related to temperatures, pressures or other measurable features of the gas turbine environment when in use. | 01-15-2015 |
20150033761 | HEAT TRANSFER ASSEMBLY AND METHODS OF ASSEMBLING THE SAME - A heat transfer assembly for controlling heat transfer of a turbine engine is provided. The turbine engine includes a housing and includes a compressor, a combustor and a turbine located within the housing. The heat transfer assembly includes a flow control device having a sidewall coupled to the turbine, the sidewall is in flow communication with a compressor vane. The sidewall is configured to define a first flow path from the compressor vane to a turbine vane and a second flow path from the compressor vane to a turbine blade. A heat exchanger is coupled to the housing and located between the compressor and the turbine, wherein the heat exchanger is in flow communication with at least one of the first flow path and the second flow path. A fluid supply device is coupled to the housing and in flow communication with the heat exchanger. | 02-05-2015 |
20150040582 | CROSSOVER COOLED AIRFOIL TRAILING EDGE - A cooling circuit for a turbine bucket having an airfoil portion includes a trailing edge cooling circuit portion provided with a first radially outwardly directed inlet passage intermediate leading and trailing edges of the airfoil portion of the bucket, extending from a platform portion of the bucket to a location adjacent a radially outer tip of the bucket, and connecting to a second radially inwardly directed passage extending from a location adjacent the radially outer tip to a location adjacent the platform portion. The second radially inwardly directed passage connects to a third trailing edge region passage, and a plurality of crossover passages connect a radially outer half of the second radially inwardly directed passage to a radially outer half of the third trailing edge region passage. | 02-12-2015 |
20150059357 | METHOD AND SYSTEM FOR PROVIDING COOLING FOR TURBINE COMPONENTS - A system for providing cooling for a turbine component that includes an outer surface exposed to combustion gases is provided. A component base includes at least one fluid supply passage coupleable to a source of cooling fluid. At least one feed passage communicates with the at least one fluid supply passage. At least one delivery channel communicates with the at least one feed passage. At least one cover layer covers the at least one feed passage and the at least one delivery channel, defining at least in part the component outer surface. At least one discharge passage extends to the outer surface. A diffuser section is defined in at least one of the at least one delivery channel and the at least one discharge passage, such that a fluid channeled through the system is diffused prior to discharge adjacent the outer surface. | 03-05-2015 |
20150075180 | SYSTEMS AND METHODS FOR PROVIDING ONE OR MORE COOLING HOLES IN A SLASH FACE OF A TURBINE BUCKET - A turbine bucket is disclosed herein. The turbine bucket may include a platform and a shank portion extending radially inward from the platform. The shank portion may include a slash face, a radial seal pin groove formed in the slash face, and at least one cooling hole disposed in the slash face about the radial seal pin groove. | 03-19-2015 |
20150082808 | GAS TURBINE ENGINES WITH TURBINE AIRFOIL COOLING - An airfoil for a gas turbine engine is provided. The airfoil includes a body with a leading edge, a trailing edge, a first side wall extending between the leading edge and the trailing edge, and a second side wall extending between the leading edge and the trailing edge. The body defines an interior cavity. The airfoil includes an interior wall disposed within the interior cavity of the body and extending between the first wall and the second wall to define a supply chamber and a leading edge chamber. The interior wall defines a cooling hole with a base portion and a locally extended portion to direct cooling air from the supply chamber to the leading edge chamber such that the cooling air impinges upon the leading edge. | 03-26-2015 |
20150096306 | GAS TURBINE AIRFOIL WITH COOLING ENHANCEMENT - A turbine nozzle vane segment includes one or more nozzle vanes extending between radially inner and outer side walls, each nozzle vane having a peripheral edge wall extending between a leading edge and a trailing edge of the vane. In one exemplary embodiment, at least one substantially radially-oriented cooling channel is formed in the peripheral edge wall at the leading edge, with openings at opposite ends of the cooling channel. The location and length of the cooling channels may vary about the peripheral edge wall, and the inner cavity of the vane may be provided with ribs extending along and adjacent the one or more cooling channels to reinforce the wall and to also provide additional cooling surface areas in the inner cavity. | 04-09-2015 |
20150107267 | REVERSE BULK FLOW EFFUSION COOLING - Effusion cooling holes formed through a transition component provided in a combustion section of a gas turbine engine. The transition component directs a hot working gas from a combustion basket to a first row of vanes in a turbine section of the engine. The effusion cooling holes are formed through an outer wall of the transition component in a direction so that the flow of air through the effusion holes is in a direction substantially opposite to the bulk flow direction of the working gas through the transition component. | 04-23-2015 |
20150121898 | TURBINE AND POWER GENERATION SYSTEM - A turbine and so on capable of enabling high reliability are provided. In the turbine of an embodiment, a turbine rotor is accommodated in a turbine casing, and is rotated by a working medium which is introduced after flowing in an inlet pipe of a combustor. A sleeve is provided at the turbine casing, and accommodates the inlet pipe therein. Here, the sleeve is thicker than the inlet pipe, and a cooling fluid whose temperature is lower than the working fluid flows between the inlet pipe and the sleeve. | 05-07-2015 |
20150292743 | FUEL NOZZLE HEAT SHIELD - An aft heat shield for a fuel nozzle tip includes: an annular shield wall; an annular shield flange extending radially outward from an aft end of the shield wall; an annular baffle flange surrounding the conical section, and disposed such that an axial gap is defined between the shield flange and the baffle flange, the baffle flange including a radially outer rim extending axially forward therefrom; and a plurality of impingement cooling holes passing through the baffle flange and oriented to as to direct air flow towards the shield wall. | 10-15-2015 |
20150308341 | COMPRESSOR INJECTOR APPARATUS AND SYSTEM - Cooling systems for high pressure compressor systems are provided. The cooling systems may comprise tangential on board injectors (“TOBIs”). The TOBIs may comprise one or more fluid channels configured to conduct cooling fluid flow to components of the compressor, including, for example, disk-hub portions of the compressor. In this regard, the TOBI may be configured to exhaust cooling air in a manner such that the exhausted air has a similar linear velocity of the disk-hub portion. The cooling air may also be exhausted in a manner that is substantially parallel to the disk-hub portion. | 10-29-2015 |
20150308342 | GAS TURBINE ENGINE VAPOR COOLED CENTRIFUGAL IMPELLER - A gas turbine engine radial impeller includes first and second impeller portions that are secured to one another along a neutral bending plane of the radial impeller. A vapor cooling cavity is provided between the first and second impeller portions. The neutral bending plane is arranged in the vapor cooling cavity. | 10-29-2015 |
20150322860 | VARIABLE VANE SEGMENT - A variable vane pack includes an inner platform, an outer platform, radially outward of the inner platform, a plurality of vanes connecting the inner platform to the outer platform, wherein the outer platform comprises a platform body and an impingement plate, the impingement plate having a radially inward impingement plate, a radially outward pressure distribution plate, and an impingement plenum defined between the radially inward impingement plate and the radially outward pressure distribution plate. | 11-12-2015 |
20150337733 | COOLING SUPPLY CIRCUIT FOR TURBOMACHINERY - Embodiments of the present disclosure include a cooling supply circuit within a turbine wheel, which may include: a substantially axial passage configured to communicate air along an axial length of the turbine wheel; a substantially radial inlet positioned within the turbine wheel between a hollow interior of the turbine wheel and the substantially axial passage, the inlet being configured to direct a rotor purge air into the substantially axial passage; and a substantially radial outlet positioned within the turbine wheel between the substantially axial passage and a cooled component coupled to a radial exterior of the turbine wheel, the outlet being configured to direct the rotor purge air towards the cooled component, wherein the outlet is axially displaced from the inlet. | 11-26-2015 |
20150345301 | ROTOR BLADE COOLING FLOW - A rotor blade includes and airfoil. The airfoil includes pressure and suction side walls which extend radially outwardly from a platform in span from a root to a tip and between a leading edge and a trialing edge. The tip includes a tip floor, a plurality of coolant outlets and a tip rail having a pressure side and suction side portions that extend radially outwardly from the tip floor. Cooling passages are circumscribed within the airfoil and are in fluid communication with one or more of the coolant outlets. A baffle extends radially outwardly from and transversely across the tip floor from the pressure side portion to the suction side portion to define first and second tip pockets. A slot is disposed along the suction side portion of the tip rail to provide for fluid communication out of one of the first or second tip pockets, thereby reducing pressure within the corresponding tip pocket. | 12-03-2015 |
20150345303 | ROTOR BLADE COOLING - A rotor blade includes a mounting portion comprising a mounting body that is formed to receive a coolant therein. An airfoil portion extends substantially radially outwardly from the mounting body and includes an airfoil body. The mounting body and the airfoil body define a plurality of primary cooling passages that extend substantially radially therein for routing the coolant through the airfoil body. Each of the primary cooling passages includes a cooling flow outlet that is formed along a tip portion of the airfoil body. The airfoil body further defines a plurality of trailing edge cooling passages, each having a coolant outlet that is formed along a trailing edge portion of the airfoil body. At least a portion of the trailing edge cooling passages are formed along a radially outer portion of the trailing edge proximate to the tip portion of the airfoil body. | 12-03-2015 |
20150345395 | TURBOFAN ENGINE WITH VARIABLE EXHAUST COOLING - Disclosed aircraft and turbofan engines have an active configuration (corresponding to flight, etc.) and an idle configuration (corresponding to ground idle). Turbofan engines comprise a core engine, a nacelle, a bypass duct therebetween, and a bypass splitter shell that extends at least partially between the nacelle and the core engine to define peripheral and interstitial bypass ducts. Bypass flow in the bypass duct splits into peripheral bypass flow and interstitial bypass flow. The relatively cool, slow interstitial bypass flow is directed into relatively hot, fast core exhaust flow from the core engine and into a mixed exhaust duct at least partially defined by the bypass splitter shell. The bypass splitter shell may be selectively positioned to increase (in the idle configuration) or to decrease (in the active configuration) the relative flow of the interstitial bypass flow, thereby cooling and/or slowing the mixed exhaust flow in the idle configuration. | 12-03-2015 |
20150345396 | GAS TURBINE ENGINE COMPONENT HAVING VASCULAR ENGINEERED LATTICE STRUCTURE - A component according to an exemplary aspect of the present disclosure includes, among other things, a wall and a vascular engineered lattice structure formed inside of the wall. The vascular engineered lattice structure includes at least one of a hollow vascular structure and a solid vascular structure configured to communicate fluid through the vascular engineered lattice structure. | 12-03-2015 |
20150354382 | EXHAUST FRAME COOLING VIA STRUT COOLING PASSAGES - A system is provided including a turbine exhaust section. The turbine exhaust section includes an exhaust flow path. The turbine exhaust section also includes an outer structure having an outer casing, an outer exhaust wall disposed along the exhaust flow path, and an outer cavity disposed between the outer exhaust wall and the outer casing. The turbine exhaust section further includes an inner structure having an inner exhaust wall disposed along the exhaust flow path, a bearing cavity disposed between the inner casing and a bearing housing. In addition, the turbine exhaust section includes a strut extending between the outer structure and the inner structure. The strut includes a first flow passage configured to flow a fluid from the bearing cavity to the outer cavity. The flow of fluid is thermally insulated from the strut. | 12-10-2015 |
20150361809 | COOLING PASSAGES FOR INNER CASING OF A TURBINE EXHAUST - An inner casing assembly for a turbine including: an annular inner casing including cooling passages, wherein each passage extends through a wall of the inner casing from a source of cooling fluid to an outer surface of the wall of the inner casing, and struts extending outward from the outer surface of the inner casing wherein the cooling passages are arranged on the inner casing such that a pair of the cooling passages is on opposite sides of each of the struts, and the cooling passages in each pair are equidistant to the corresponding strut. | 12-17-2015 |
20150361827 | Acoustic Treatment to Mitigate Fan Noise - A cooling manifold has a plurality of pieces. The pieces extend in a circumferential direction to abutting flanges. The flanges are secured together at circumferential ends of each piece. Cooling channels are formed in between inner and outer walls. Air inlets are formed in the pieces with the air inlets delivering air in the interior. There are fingers on an outer periphery. The fingers are aligned within an air outlet. The air can be delivered into the inlet, cool the interior, and leave through the outlet extending to a main conduit. The main conduit is secured directly to an outer periphery of the cooling manifold. | 12-17-2015 |
20150369055 | BELL MOUTH INLET FOR TURBINE BLADE - An airfoil assembly for a gas turbine engine is disclosed and includes a platform portion defining a portion of a gas flow path and a root portion for attachment of the turbine airfoil, the root portion including a bottom surface including a bottom area and a plurality of inlets that define a total inlet area as a ratio of the inlet area to the bottom area. An airfoil extends from the platform and including a plurality of cooling air passages in communication with the plurality of inlets. | 12-24-2015 |
20150377135 | METHOD AND SYSTEM FOR RADIAL TUBULAR DUCT HEAT EXCHANGERS - A method and system for a heat exchanger are provided. The heat exchanger includes a plurality of arcuate heat exchanger segments, each including a first header configured to extend circumferentially about at least a portion of a circumference of an internal surface of a fluid duct. The heat exchanger also includes a second header configured to extend circumferentially about the portion spaced axially apart from the first header in a direction opposite of fluid flow through the fluid duct and a first plurality of heat exchanger tubes extending generally axially between the first header and the second header, the first plurality of heat exchanger tubes each including a first flow path separate from a second flow path of any other of the first plurality of heat exchanger tubes, the first flow path changing direction along the flow path from the first header to the second header. | 12-31-2015 |
20160003077 | GAS TURBINE ENGINE TURBINE VANE RAIL SEAL - A stator vane seal assembly for a gas turbine engine that includes a vane that has a rail with a scallop that extends axially through the rail. A scallop seal obstructs the scallop. | 01-07-2016 |
20160003088 | COOLING DEVICE FOR THE CASING OF AN AIRCRAFT JET ENGINE COMPRISING A SUPPORTING DEVICE - A cooling device for the casing of a jet engine, includes a cooling tube and a supporting device, the supporting device including an attachment plate and an attachment clamp, the attachment clamp including a clamp body surrounding the cooling tube and an attachment element attached to the attachment plate, wherein the attachment plate includes: an opening; and an attachment tab secured to the main wall of the attachment plate and arranged on at one portion of the periphery of the opening, the attachment clamp passing through the opening, and the attachment element of the attachment clamp being attached to the attachment tab. | 01-07-2016 |
20160003152 | GAS TURBINE ENGINE MULTI-VANED STATOR COOLING CONFIGURATION - A stator for a gas turbine engine has a platform supporting multiple vanes that includes first and second vanes respectively. First and second regions are arranged at the same location on the first and second vanes. The first and second regions respectively include first and second cooling hole configurations that are different than one another. | 01-07-2016 |
20160010463 | GAS TURBINE ENGINE HIGH LIFT AIRFOIL COOLING IN STAGNATION ZONE | 01-14-2016 |
20160010563 | OIL LOSS PROTECTION FOR A FAN DRIVE GEAR SYSTEM | 01-14-2016 |
20160025010 | TURBINE ENGINE AND TURBINE ENGINE COMPONENT WITH COOLING PEDESTALS - A turbine engine component includes a surface to be cooled by a flow of cooling air and a plurality of pedestals projecting from the surface to be cooled. At least one of the pedestals includes a pedestal surface oriented such that a ray normal to the pedestal surface and directed away from the pedestal surface intersects a line defined by an intersection between the surface to be cooled and the pedestal. | 01-28-2016 |
20160032767 | GAS TURBINE ENGINE WITH AXIAL COMPRESSOR WITH INTERNAL COOLING PATHWAYS - A gas turbine engine may include an axial high pressure compressor having an air flow pathway positioned between the inner and outer rim of the rotor section. The air flow pathway includes an inlet port, a transition segment, an axial segment, and an outlet port. The pathway may be a tube having an ovoid cross sectional shape and is substantially co-planar to the outer surface of the outer rim. The pathway may traverse the rotor section from the first rotor segment to the rear hub. | 02-04-2016 |
20160047312 | GAS TURBINE SYSTEM - A gas turbine system ( | 02-18-2016 |
20160054003 | COMBUSTOR CAP ASSEMBLY - A combustor cap assembly includes an impingement plate coupled to an annular shroud and a cap plate which is coupled to the impingement plate and forms an impingement air plenum therebetween. The cap assembly further includes a flow conditioning plate which is coupled to a forward end portion of the shroud. The flow conditioning plate includes an inner band portion, an outer band portion and an annular portion which extends radially therebetween. The annular portion includes upstream side, a downstream side and a plurality of flow conditioning passages which provide for fluid communication through the upstream and downstream sides. The inner band portion of the flow conditioning plate at least partially defines an exhaust channel. The exhaust channel is in fluid communication with the impingement air plenum and an exhaust outlet. The exhaust outlet is positioned to route cooling air from the impingement air plenum into an annular flow passage defined within a combustor. | 02-25-2016 |
20160054004 | COMBUSTOR CAP ASSEMBLY - A combustor cap assembly includes an annular shroud and an impingement plate coupled to the shroud. The impingement plate at least partially defines a plurality of impingement cooling holes and a cooling flow return passage. A cap plate is coupled to the impingement plate. The cap plate includes an impingement side which faces a second side portion of the impingement plate where the impingement side is axially spaced from the second side portion to define an impingement air plenum therebetween. The cooling flow return passage is in fluid communication with the impingement air plenum. A fluid conduit extends from a first side portion of the impingement plate towards a first end portion of the shroud. The fluid conduit is in fluid communication with the cooling flow return passage and provides for fluid communication out of the impingement air plenum. | 02-25-2016 |
20160061043 | TURBINE BUCKET - A turbine bucket includes a leading edge, a trailing edge, a root portion, and a tip portion. The turbine bucket also includes one or more cooling passages extending through a body of the turbine bucket from an inlet to an outlet. The cooling passages are configured to route a cooling flow of fluid through the turbine bucket. The turbine bucket further includes a plenum defined within the tip portion to receive the fluid from the outlet of the cooling passages for expulsion of the cooling flow of fluid into a main flow path via at least one outlet hole proximate the trailing edge of the turbine bucket. | 03-03-2016 |
20160061113 | ACTIVELY COOLED BLADE OUTER AIR SEAL - A turbine component includes a main body with an upstream end and a downstream end. A cooling passage network is interior to the main body and has multiple cooling passages. A shared passage wall in the cooling passage network includes a cross passage opening connecting each passage partially defined by the shared passage wall. The cross passage further including a pressure balancing feature operable to reduce a local pressure differential in a fluid pressure in each adjacent passage at the cross passage. | 03-03-2016 |
20160061451 | GAS TURBINE ENGINES WITH PLUG RESISTANT EFFUSION COOLING HOLES - A combustor for a turbine engine is provided. A first liner has a first surface and a second surface. A second liner forms a combustion chamber with the second side of the first liner, and the combustion chamber configured to receive an air-fuel mixture for combustion therein. The first liner defines a plurality of effusion cooling holes configured to form a film of cooling air on the second surface of the first liner. The plurality of effusion cooling holes includes a first effusion cooling hole extending from the first surface to the second surface and including an inlet portion extending from the first surface, a metering portion fluidly coupled to the inlet portion, and an outlet portion fluidly coupled to the metering portion and extending to the second surface. The inlet portion is larger than the metering portion. | 03-03-2016 |
20160069190 | BEVELED COVERPLATE - A gas turbine engine includes a platform that has a gas path side, a non-gas path side, a first mate face, and a second mate face. The second mate face has a beveled edge sloping towards the first mate face. The gas turbine engine also includes a coverplate that includes a first bend, a flat portion substantially parallel to the first mate face and a first wing substantially parallel to the second mate face. | 03-10-2016 |
20160069193 | COOLANT FLOW REDIRECTION COMPONENT - A gas turbine engine includes a compressor section, a combustor fluidly connected to the compressor section, and a turbine section fluidly connected to the combustor and mechanically connected to the compressor section via a shaft. Multiple rotors are disposed in one of the compressor section and the turbine section. Each of the rotors includes a rotor disk portion having a radially inward bore, and is static relative to the shaft. Each rotor is axially adjacent at least one other rotor and a gap is defined between each rotor and an adjacent rotor. A cooling passage for a cooling flow is defined between the shaft and the rotors, and a cooling flow redirection component is disposed at the gap and is operable to redirect the cooling flow in the cooling passage into the gap. | 03-10-2016 |
20160069569 | FILM COOLING CIRCUIT FOR A COMBUSTOR LINER - A liner of a combustor wall assembly generally defines a combustion chamber and includes a film cooling circuit having a channel communicating through a hot face of a substrate of the liner. An aperture of the circuit is defined by the substrate, extends through the cold face and is in fluid communication with the channel. The hot face of the substrate is covered with a coating that extends over and thus defines in-part the channel. A film cooling hole extends through the coating and is in fluid communication with the channel. A method of manufacturing the circuit includes casting the substrate with the aperture and hole; then placing an insert into the channel prior to application of the coating over the substrate and insert. The insert is then removed and the film cooling hole is formed through the coating. | 03-10-2016 |
20160076451 | FILM HOLE WITH IN-WALL ACCUMULATOR - A gas turbine engine component is described. The component includes a component wall having an internal surface that is adjacent a flow of coolant and an external surface that is adjacent a flow of gas. The component wall includes a cooling hole that has an inlet defined by the internal surface and an outlet defined by the external surface. The cooling holes also has a metering location having the smallest cross-section area of the cooling hole, an internal diffuser positioned between the inlet and the metering location, an accumulation diverter portion of the internal diffuser and an accumulator portion of the internal diffuser. | 03-17-2016 |
20160084164 | PLATE FOR METERING FLOW - A cooling device for a gas turbine engine component comprises a gas turbine engine component having an upstream channel and a downstream channel that define a cooling flow path. A meter feature includes at least one hole to meter flow from the upstream channel to the downstream channel, and has an upstream side and a downstream side. An exit diffuser extends outwardly from the downstream side of the meter feature to control flow in a desired direction into the downstream channel. A gas turbine engine is also disclosed. | 03-24-2016 |
20160108740 | GAS TURBINE ENGINES WITH IMPROVED LEADING EDGE AIRFOIL COOLING - An airfoil for a gas turbine engine includes a body with a first side wall and a second side wall joined at a leading edge and a trailing edge, the first side wall having a first interior surface and the second side wall having a second interior surface. The airfoil further includes an internal wall disposed within of the body and extending between the first interior surface and the second interior surface to define a supply passage and a leading edge passage. The internal wall defines a plurality of cooling holes to direct cooling air from the supply passage to the leading edge passage such that the cooling air impinges upon the leading edge. The airfoil further includes a first plurality of grooves formed in the first interior surface, each the first plurality of grooves extending in a chordwise direction within the leading edge passage. | 04-21-2016 |
20160108755 | GAS TURBINE ENGINE COMPONENT - A gas turbine engine component includes an exterior pressure side with a plurality of cooling holes located in the exterior pressure side. A relief cut surrounds at least one of the plurality of cooling holes. | 04-21-2016 |
20160115871 | COOLING CONFIGURATION FOR A COMPONENT - A component includes at least one thermal riser that extends from an exterior surface of the component. At least one cooling passage extends through a wall and adjoins an interior cooling passage and provides an exterior surface. At least one cooling passage is configured to direct cooling fluid through the wall adjacent to at least one thermocouple. | 04-28-2016 |
20160123156 | COOLED COMPONENT - A cooled gas turbine engine component comprises a wall having first and second surfaces. The second surface has a plurality of recesses and each recess has a planar upstream end surface arranged so that it hangs over the upstream end of the recess. Each recess has a depth equal to the required depth plus the thickness of the thermal barrier coating to be deposited. The wall has a plurality of angled effusion cooling apertures extending from the first surface towards the second surface. Each effusion cooling aperture has an inlet in the first surface and an outlet in the end surface of a corresponding one of the recesses in the second surface. Each recess has smoothly curved transitions from the end surface and side surfaces to the second surface. Blocking of the effusion cooling apertures by thermal barrier coating is reduced. | 05-05-2016 |
20160130950 | GAS TURBINE ENGINE COMPONENT WITH RIB SUPPORT - A component according to an exemplary aspect of the present disclosure includes, among other things, an airfoil that includes a first sidewall and a second sidewall joined together at a leading edge and a trailing edge and extending from a base to a tip. A plenum is defined inside the airfoil. A first cooling cavity merges into the plenum and a second cooling cavity merges into the plenum. A rib extends from at least one of the first sidewall and the second sidewall at least partially into the plenum to separate the first cooling cavity from the second cooling cavity. | 05-12-2016 |
20160131037 | SUPPLY DUCT FOR COOLING AIR - In a featured embodiment, a gas turbine engine has a compressor section having a downstream rotor and a diffuser downstream of the compressor section. A combustor receives air downstream of the diffuser. A turbine section has at least one component to be cooled. A conduit is spaced from the diffuser and defines a cooling airflow path. The cooling airflow path is separate from an airflow downstream the diffuser, and passing to the combustor. The conduit passes cooling air to the component to be cooled. | 05-12-2016 |
20160146044 | INTERNALLY COOLED TURBINE PLATFORM - A turbine flowpath component for a gas turbine engine includes a platform defining at least one internal cooling cavity and a turbine flowpath component cover disposed radially outward of the platform. The turbine flowpath component cover defines a radially outward edge of the at least one internal cooling cavity. The turbine flowpath component cover further defines a transition region in which a radial depth of the at least one internal cooling cavity varies from a first depth to a second depth. | 05-26-2016 |
20160160648 | ROTOR DISK ASSEMBLY FOR A GAS TURBINE ENGINE - A rotor assembly for a gas turbine engine includes a rotor disk. The rotor disk includes a heat barrier feature radially inward of a rotor blade and radially outward of a rotor disk. | 06-09-2016 |
20160160652 | COOLED POCKET IN A TURBINE VANE PLATFORM - An airfoil includes a platform including a pocket, a continuous solid cover plate bonded over the pocket, and a plurality of apertures in the platform in fluid communication with the pocket and adjacent the continuous solid cover plate. A gas turbine engine including the airfoil and a method of cooling an article of a gas turbine engine are also disclosed. | 06-09-2016 |
20160160654 | TURBINE AIRFOIL SEGMENT HAVING FILM COOLING HOLE ARRANGEMENT - A turbine airfoil segment includes inner and outer platforms that are joined by at least one airfoil. The airfoil includes leading and trailing edges that are joined by spaced apart first and second sides to provide an exterior airfoil surface. At least one of the inner and outer platforms includes film cooling holes that have external breakout points that are located in substantial conformance with the Cartesian coordinates set forth in Table 1 for the inner platform or Table 2 for the outer platform. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate. The film cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.20 inches (5.0 mm). | 06-09-2016 |
20160160656 | TURBINE AIRFOIL SEGMENT HAVING FILM COOLING HOLE ARRANGEMENT - A turbine airfoil segment includes inner and outer platforms that are joined by at least one airfoil. The airfoil includes leading and trailing edges that are joined by spaced apart first and second sides to provide an exterior airfoil surface. At least one of the inner and outer platforms includes film cooling holes that have external breakout points that are located in substantial conformance with the Cartesian coordinates set forth in Table 1 for the inner platform or Table 2 for the outer platform. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate. The film cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.20 inches (5.0 mm). | 06-09-2016 |
20160169515 | EDGE COOLING FOR COMBUSTOR PANELS | 06-16-2016 |
20160177735 | TURBINE BLADE HAVING FILM COOLING HOLE ARRANGEMENT | 06-23-2016 |
20160177738 | GAS TURBINE ENGINE COMPONENT WITH FILM COOLING HOLE | 06-23-2016 |
20160177831 | COOLING-AIR SUPPLY DEVICE FOR A GAS TURBINE | 06-23-2016 |
20160194966 | TURBINE AIRFOIL SEGMENT HAVING FILM COOLING HOLE ARRANGEMENT | 07-07-2016 |
20160201473 | TURBINE BLADE HAVING FILM COOLING HOLE ARRANGEMENT | 07-14-2016 |
20160201474 | GAS TURBINE ENGINE COMPONENT WITH FILM COOLING HOLE FEATURE | 07-14-2016 |
20160201560 | COOLING AIR LINE FOR REMOVING COOLING AIR FROM A MANHOLE OF A GAS TURBINE | 07-14-2016 |
20160201989 | METHOD AND SYSTEM FOR RADIAL TUBULAR HEAT EXCHANGERS | 07-14-2016 |
20160251962 | GAS TURBINE | 09-01-2016 |
20160251974 | INCIDENT TOLERANT TURBINE VANE COOLING | 09-01-2016 |
20160251980 | INCIDENT TOLERANT TURBINE VANE GAP FLOW DISCOURAGEMENT | 09-01-2016 |
20160251981 | GAS TURBINE | 09-01-2016 |
20160376989 | CORE ASSEMBLY FOR GAS TURBINE ENGINE - A core assembly includes a core that includes an exterior surface that has a recessed area that extends along the exterior surface. An insert includes a contact surface that corresponds to the recessed area. | 12-29-2016 |
20170234142 | Rotor Blade Trailing Edge Cooling | 08-17-2017 |
20180023475 | GAS TURBINE ENGINE WITH HEAT PIPE FOR THERMAL ENERGY DISSIPATION | 01-25-2018 |
20180023906 | PASSIVE HEAT EXCHANGER VALVE | 01-25-2018 |
20190145272 | BLADE OUTER AIR SEAL FOR A GAS TURBINE ENGINE | 05-16-2019 |