Entries |
Document | Title | Date |
20090000305 | Dual Flow Turbine Engine Equipped with a Precooler - The invention concerns a precooler ( | 01-01-2009 |
20090193811 | BLEED AIRFLOW BALANCING CONTROL USING SIMPLIFIED SENSING - A simplified air bleed balancing control system for a pair of aircraft gas turbine engines reduces the number of pressure transducers and differential pressure transducers. Advantages include lower weight, less expensive system, better total system MTBF (mean time before failure), acceptable differential pressure transducer drift identification and compensation by the digital controller, and fewer maintenance tasks. | 08-06-2009 |
20090235668 | Insulator bushing for combustion liner - A combustor for a turbine including: a combustor liner, a first flow sleeve surrounding the combustor liner to define a first flow annulus, the first flow sleeve having cooling holes for directing compressor discharge air as cooling air into the first flow annulus, a transition piece body connected to the combustor liner to carry hot combustion gases to the turbine; a second flow sleeve surrounding the transition piece body, the second flow sleeve having cooling holes for directing compressor discharge air as cooling air into a second flow annulus between the second flow sleeve and the transition piece body; at least one dilution hole in the combustor liner for flowing compressor air into a combustion chamber defined by the combustor liner; and a bushing seated in at least one of the cooling or dilution holes and secured with respect thereto for defining a flow passage for compressor discharge air through the hole. | 09-24-2009 |
20090272120 | Diffuser for a gas turbine, and gas turbine for power generation - The invention relates to a gas turbine, for energy generation, with a compressor, arranged coaxially to a rotor, mounted such as to rotate, for the compression of an inlet gaseous fluid, at least partly serving for combustion of a fuel in a subsequent annular combustion chamber, with generation of a hot working medium, with an annular diffuser arranged coaxially to the rotor, between the compressor and the annular combustion chamber, for distribution and deflection of the fluid, whereby a part of the fluid is diverted as cooling fluid for the turbine stages after the combustion chamber, by means of a dividing element, arranged in the fluid flow. According to the invention, a compact diffuser and an economical gas turbine with an improved flow for the diversion of cooling air may be achieved, whereby the annular dividing element, arranged coaxially to the rotor, comprises at least one opening, facing the fluid flow and the dividing element is supported on the diffuser, by means of several hollow rib-like support elements, by means of which the cooling fluid, diverted through the opening, is first directed towards the rotor. | 11-05-2009 |
20090293495 | TURBINE AIRFOIL WITH METERED COOLING CAVITY - A turbine airfoil for a gas turbine engine includes: (a) spaced-apart pressure and suction sidewalls extending between a leading edge and a trailing edge; (b) a first cavity disposed between the pressure and suction sidewalls, the first cavity being adapted to be fed cooling air from a source within the engine, and connected to at least one film cooling hole which communicates with an exterior surface of the airfoil; (c) a second cavity disposed between the pressure and suction sidewalls, the second cavity being adapted to be fed cooling air from a source within the engine, and connected to at least one film cooling hole which communicates solely with the suction sidewall of the airfoil; and (d) a metering structure adapted to substantially restrict air flow into the second cavity. | 12-03-2009 |
20090293496 | GAS TURBINE ENGINES GENERATING ELECTRICITY BY COOLING COOLING AIR - A portion of cooling air for cooling the turbine section of a gas turbine engine is tapped and passed through a heat exchanger. The portion of the cooling air is cooled in the heat exchanger, and the heat taken out of the portion of the cooling air is utilized to generate electricity. | 12-03-2009 |
20100107649 | Gas Turbine Engine With Fuel Booster - A gas turbine engine including a compressor section, a combustor section including a combustor, a turbine section, a gaseous fuel supply conduit with an upstream section and a downstream section, a fuel booster, and a heat exchanger is provided. The fuel booster is located in the gaseous fuel supply conduit. The fuel booster has a driving expander with a driving fluid inlet and a driving fluid outlet for discharging expanded air and a fuel compressor with a low pressure fuel inlet connected to the upstream section of the gaseous fuel supply conduit and a high pressure fuel outlet connected to the downstream section of the gaseous fuel supply conduit. The heat exchanger is located between the driving fluid on the one side and the gaseous fuel on the other side so that a heat transfer between the air and the gaseous fuel is possible. | 05-06-2010 |
20100126181 | METHOD FOR CONTROLLING AN EXHAUST GAS RECIRCULATION SYSTEM - An embodiment of the present invention takes the form of a method and system that may reduce the level of SOx emissions by recirculating a portion of the exhaust of at least one turbomachine; the portion of exhaust may be mixed with the inlet air prior to re-entering the turbomachine. An embodiment of the present invention may incorporate an inlet bleed heat system to reduce the likelihood of the liquid products forming from SOx emissions. Here, a method may maintain a temperature of the inlet fluid above a condensation temperature. | 05-27-2010 |
20100162722 | TIP CLEARANCE CONTROL - Aspects of the invention relate to a system and method for actively managing blade tip clearances in a turbine engine, particularly under start up and steady state operating conditions. During start up, the rotor can be cooled while the blade ring can be heated. During steady state operation, the rotor may be heated while the blade ring may be cooled. The sources of the heating fluid and cooling fluid can be compressor delivery air, boiler steam and ambient air. The fluids can be compressed to maintain efficiency. | 07-01-2010 |
20100170264 | TURBINE SHROUD THERMAL DISTORTION CONTROL - A shroud suitable for use in a gas turbine engine exhibits substantially uniform thermal growth. | 07-08-2010 |
20100175387 | Cooling of Turbine Components Using Combustor Shell Air - A turbine engine assembly for a generator including a turbine engine having a compressor section, a combustor section and a turbine section. An air bleed line is in communication with the combustor section for receiving combustor shell air from the combustor section and conveying the combustor shell air as bleed air to a plurality of stages of the turbine section. Bleed air is controlled to flow through the air bleed line when an operating load of the turbine engine assembly is less than a base load of the engine to bypass air exiting the compressor section around a combustor in the combustor section and effect a flow of high pressure combustor shell air to the stages of the turbine section. | 07-15-2010 |
20100192593 | LINEAR QUADRATIC REGULATOR CONTROL FOR BLEED AIR SYSTEM FAN AIR VALVE - A system and method are provided for controlling the temperature of engine bleed air from a turbofan gas turbine engine. The system includes a fan air valve and a fan air valve controller. The fan air valve is adapted to receive a flow of fan air from a turbofan gas turbine engine intake fan. The fan air valve is coupled to receive valve position commands and is configured, in response to the valve position commands, to move to a valve position to thereby control the engine bleed air temperature. The fan air valve controller is configured to implement a linear quadratic regulator (LQR) control. The fan air valve controller is adapted to receive a plurality of sensor signals, each sensor signal representative of one or more system parameters, and is configured, in response to the sensor signals, to supply the valve position commands to the fan air valve. | 08-05-2010 |
20100229567 | TURBINE ENGINE ARRANGEMENT - A turbine engine arrangement incorporates an electrical machine combination which can act as a generator or motor. Thus, the combination during initial turbine engine arrangement start-up acts through this motor to drive shaft rotation. This shaft rotation creates core compressed air which is bled through a valve to an air turbine such that the electrical machine when acting as a motor or generator at low engine loads is appropriately cooled. The arrangement also includes a compressor fan which generates an airflow utilised by the air turbine in order to drive the electrical machine as a generator of electrical power when engine conditions dictate. A duct is provided to direct airflow exhausted from the air turbine to a heat exchanger arranged to exchange heat with a bypass flow from the compressor fan. In such circumstances, the generators can operate substantially independently of rotational speed of the engine shafts within the core whilst bled air through the valve acts to cool the electric machine during initial start-up. Generally, a gearbox is provided whereby matching of air turbine rotation with generator requirements is achieved whilst also through a coupling allowing connection to the primary drive shaft of the core to either allow engine start-up or driving of the generators when engine speed is appropriate. | 09-16-2010 |
20100242491 | Systems and Methods for Controlling Compressor Extraction Cooling - Embodiments of methods and apparatus for providing compressor extraction cooling are provided. According to one example embodiment, a method is disclosed for controlling compressor extraction cooling. The method can include providing a cooling medium. The method can include extracting air from a compressor associated with a gas turbine. The method can also include introducing the cooling medium to the compressor extraction air, wherein the compressor extraction air is cooled by the cooling medium prior to or during introduction to the turbine section. Furthermore, method can include selectively controlling at least one of the compressor extraction air or the cooling medium based at least in part on a characteristic associated with the gas turbine. | 09-30-2010 |
20100257869 | Diffuser arranged between the compressor and the combustion chamber of a gas turbine - The invention relates to a gas turbine, comprising an annular combustion chamber and an upstream diffuser, with a throughflow essentially parallel to a turbine longitudinal axis, at a distance from said axis at least partly less than the annular combustion chamber, in which a compressed gas may be divided into several partial flows at a branching point, whereby at least one of the partial flows is a cooling gas flow. A main deflection region is provided in said diffuser, directed at an angle to the turbine longitudinal axis towards the annular combustion chamber. | 10-14-2010 |
20100281879 | MULTI-SOURCE GAS TURBINE COOLING - A cooling arrangement for a gas turbine engine includes: (a) a turbine nozzle having: (i) spaced-apart arcuate inner and outer bands; and (ii) a hollow, airfoil-shaped turbine vane extending between the inner and outer bands, the vane disposed in a primary flowpath of the engine; (b) a supporting structure coupled to the outer band such that an outer band cavity is defined between the outer band and the stationary structure; (c) a first conduit passing through the outer band cavity and communicating with the interior of the vane, the first conduit coupled to a first source of cooling air within the engine; and (d) a second conduit communicating with the outer band cavity, the second conduit coupled to a second source of cooling air within the engine. | 11-11-2010 |
20100300113 | ANTI-VORTEX DEVICE FOR A GAS TURBINE ENGINE COMPRESSOR - An anti-vortex device for use in a compressor rotor assembly of a gas turbine engine is described. Spaced-apart radial passageways extend from an axially extending passage provided in a central area of the device to an outer peripheral rim surface thereof. The radial passageways channel air from the primary gaspath about the rotor assembly to the axially extending passage where the air is directed into a central axial passage of the rotor assembly. | 12-02-2010 |
20100319359 | SYSTEM AND METHOD FOR HEATING TURBINE FUEL IN A SIMPLE CYCLE PLANT - In certain embodiments, a system includes a fuel heater. The fuel heater includes a first heat exchanger configured to receive compressed air from a compressor and to transfer heat from the compressed air to a cooled intermediate heat transfer media to generate a heated intermediate heat transfer media. The fuel heater also includes a second heat exchanger configured to receive the heated intermediate heat transfer media from the first heat exchanger and to transfer heat from the heated intermediate heat transfer media to a fuel. The first heat exchanger is configured to receive the cooled intermediate heat transfer media from the second heat exchanger. | 12-23-2010 |
20100326088 | GAS TURBINE, LOAD COUPLING OF GAS TURBINE, AND COOLING METHOD OF GAS TURBINE COMPRESSOR - To include a cooling passage leading from a latter stage of a compressor, via an external cooler, to a hollow part provided in a load coupling that couples a rotor of the compressor and a rotor of a turbine to each other, and also leading from the hollow part to the latter stage of the compressor, and a centrifugal compressor that raises air pressure in the hollow part with rotation of the load coupling, in the hollow part. Therefore, pressure of the air in the hollow part is raised by the centrifugal compressor by using a centrifugal force resulting from the rotation of the load coupling at the time of operating the gas turbine. Cooled air flows from the hollow part to between compressor rotor blades and compressor vanes at the latter stage, and then flows into the hollow part, thereby enabling to efficiently reduce temperature in the hollow part. | 12-30-2010 |
20110067412 | GAS TURBINE AIRCRAFT ENGINES AND OPERATION THEREOF - There is disclosed an aircraft propulsion arrangement including a gas turbine aircraft engine having a compressor, an oil system configured to route engine oil through a heat exchanger mounted so as to define part of an aerodynamic surface to the flow of ambient air, and a duct arrangement fluidly connecting the compressor to the heat exchanger. There is also proposed a method of operating the engine, the method involving the steps of: (a) flowing engine oil through the heat exchanger and thus into heat-exchange relationship with said ambient air; and (b) directing a bleed flow of compressor gas drawn from the compressor along said duct arrangement and into heat-exchange relationship with said oil in said heat exchanger, wherein said directing step (h) is performed selectively. | 03-24-2011 |
20110067413 | EJECTOR CONTROLLED TWIN AIR SOURCE GAS TURBINE PRESSURIZING AIR SYSTEM - A passive pressurizing air system for a gas turbine engine includes a flow path for directing an air flow having a low temperature and low pressure, extending through a cavity to a pressurized area of the engine. The cavity contains pressurized air having a high temperature and high pressure. An air flow mixing apparatus is provided for adding the pressurized air from the cavity into the flow path to provide a mixed air flow having an intermediate temperature and intermediate pressure. | 03-24-2011 |
20110088405 | GAS TURBINE ENGINE TEMPERATURE MODULATED COOLING FLOW - A gas turbine engine cooling system includes a heat exchanger in fluid communication with a source of cooling air, a first cooling circuit including a first heat exchanger circuit in the heat exchanger and a first bypass circuit with a first bypass valve for selectively bypassing at least a portion of first airflow around the first heat exchanger circuit. A second cooling circuit may be used having a second heat exchanger circuit in the heat exchanger and a shutoff control valve operably disposed in the second cooling circuit upstream of the second heat exchanger circuit and the heat exchanger. A circuit inlet of the first cooling circuit may be used to bleed a portion of compressor discharge bleed air for the first airflow to cool turbine blades mounted on a rotor disk using an annular flow inducer downstream of the first bypass valve and the heat exchanger. | 04-21-2011 |
20110131999 | THERMOELECTRIC GENERATOR ON AN AIRCRAFT BLEED SYSTEM - A device for producing electrical power. A thermoelectric device is coupled to an aircraft bleed system for generating electrical power using temperature differentials between ram air and bleed air. | 06-09-2011 |
20110162383 | Ejector/Mixer Nozzle for Noise Reduction - An inlet bleed heat system in a gas turbine includes a compressor discharge extraction manifold that extracts compressor discharge air, an inlet bleed heat manifold receiving the compressor discharge air, and a plurality of acoustic dispersion nozzles disposed at an output end of the inlet bleed heat manifold that reduce a velocity of the compressor discharge air in the inlet bleed heat manifold. Noise is generated from the shearing action between the surrounding atmosphere and air jets from orifices. When the air jet velocity is slowed using, for example, a multi-stage ejector/mixer, noise can be abated. | 07-07-2011 |
20110162384 | TEMPERATURE ACTIVATED VALVES FOR GAS TURBINES - A system for modulating the amount of air supplied through a pressure boundary in a gas turbine is disclosed that includes a passageway located on the pressure boundary. Additionally, a temperature activated valve is mounted within the passageway and is configured to activate at a predetermined temperature threshold. Specifically, the temperature activated valve activates from a closed position to an open position when the local temperature at the temperature activated valve reaches or exceeds the predetermined temperature threshold to allow air to flow through the passageway. | 07-07-2011 |
20110247344 | REAR HUB COOLING FOR HIGH PRESSURE COMPRESSOR - In one exemplary embodiment, a gas turbine engine includes a turbine and a high pressure compressor. The high pressure compressor includes a last stage having a last stage compressor blade and a last stage vane. The gas turbine engine includes a first flow path through which bleed air flows to the turbine and a second flow path through which air from the last stage of the high pressure compressor flows. The bleed air in the first flow path exchanges heat with a portion of the air in the second flow path in a heat exchanger to cool the air in the second flow path. The cooled air in the second flow path is returned to the high pressure compressor to cool the high pressure compressor. | 10-13-2011 |
20110302928 | LIQUID-GAS HEAT EXCHANGER - The heat exchanger includes a large number of small, closely-spaced modules. Within each module of one embodiment, the fuel flows through a series of parallel micro-channels, while the air flows externally over rows of short, straight fins perpendicular to the direction of fuel flow. A theoretical model was developed to predict the thermal performance of the module for various operating conditions. To confirm the accuracy of the model, a module was constructed and tested using water to simulate the aircraft fuel. | 12-15-2011 |
20110314835 | COOLER IN NACELLE WITH RADIAL COOLANT - A cooling system in an aircraft gas turbine engine includes a heat exchanger positioned within an annular nacelle space surrounding a bypass duct of the engine. The heat exchanger has a radial coolant passage extending between the bypass duct and ambient air surrounding the nacelle, and a flow passage extending substantially normal to the radial passage for direction of a fluid to be cooled therethrough. A configuration of this nature may assist in defining a no-flow length of the heat exchanger in a third direction normal to the other two mentioned directions, which may allow for improved performance within a given radial envelope. | 12-29-2011 |
20120031105 | PRESSURE-ACTUATED PLUG - A plug for regulating a flow of gas in a system is disclosed. The plug includes a housing disposed on a temperature boundary in a system. The housing defines a passage for flowing gas therethrough. The plug further includes at least one pressure-actuated valve disposed in the passage and movable between an open position and a closed position. The at least one pressure-actuated valve moves from the open position to the closed position as the pressure of the gas increases and moves from the closed position to the open position as the pressure of the gas decreases. | 02-09-2012 |
20120036864 | GAS TURBINE ENGINE AND METHOD FOR COOLING THE COMPRESSOR OF A GAS TURBINE ENGINE - A gas turbine engine includes a compressor with rotor blades having roots connected into seats of a compressor drum. The rotor blade roots and/or the compressor drum have longitudinal passages for a cooling fluid, connecting higher pressure areas to lower pressure areas of the gas turbine engine. | 02-16-2012 |
20120060506 | GAS TURBINE ENGINE - A gas turbine engine has a compressor section with rotational compressor components rotatable with respect to static compressor components. A compressed air bleed arrangement is provided to cool one or more other rotational components of the gas turbine engine. The compressed air bleed arrangement takes a flow of compressed air from the compressor section along an off-take passage. The off-take passage opens in the compressor section at an off-take port. The off-take passage is rotatable, in use, with the rotational compressor components. The compressed air bleed arrangement is operable to provide the air in the off-take passage with higher static pressure than the air in the compressor section at the off-take port, by diffusing the air in the off-take passage. The off-take passage further includes off-take vanes, operable to increase the tangential velocity of the air in the off-take passage compared with the air at the off-take port. | 03-15-2012 |
20120060507 | GAS TURBINE ENGINE - A gas turbine engine comprising a turbine section cooling system and a method of cooling a turbine section of a gas turbine engine is provided. The gas turbine engine comprises in flow series a compressor section, a combustor, and a turbine section, the engine further comprising a turbine section cooling system. The turbine section cooling system including a first compressed air bleed arrangement and a second compressed air bleed arrangement. The first compressed air bleed arrangement bleeds a first flow of compressed air from a high pressure stage of the compressor section. The first flow of compressed air bypasses the combustor and arrives at the turbine section to form a sealing and/or cooling flow at a row of stator vanes upstream of an adjacent rotor disc. The second compressed air bleed arrangement bleeds a second flow of compressed air from one or more lower pressure stages of the compressor section. The second flow of compressed air bypasses the combustor and arrives at the turbine section to fowl a cooling flow. A first portion of the cooling flow is routed to a front face of the rotor disc and a second portion of the cooling flow is routed to a rear face of the rotor disc. | 03-15-2012 |
20120111020 | CONTROL METHOD FOR COOLING A TURBINE STAGE IN A GAS TURBINE - A control method for cooling a turbine stage of a gas turbine, whereby cooling air is bled from combustion air flowing in a compressor of the gas turbine, and is fed to a cooling circuit staring from a stator of the turbine stage; and cooling airflow is adjusted as a function of the pressure at the inlet of the cooling circuit, and as a function of the combustion air pressure at the exhaust of the compressor; more specifically, there is a feedback control setting a setpoint, which is predetermined as a function of the power output of the turbine to reduce contaminating emissions. | 05-10-2012 |
20120117977 | METHOD AND SYSTEM FOR CONTROLLING A SET POINT FOR EXTRACTING AIR FROM A COMPRESSOR TO PROVIDE TURBINE COOLING AIR IN A GAS TURBINE - A method for controlling the generation of turbine cooling air from air extracted from a compressor of a gas turbine including: extracting compressed air from a low pressure and a high pressure stage of the compressor; adding in an ejector the compressed air from the low pressure stage to the air from the high pressure stage and discharging the combined air as turbine cooling air; bypassing the ejector with a bypass portion of the extracted compressed air from the high pressure stage; in response to turning on the flow of extracted compressed air from the low pressure stage, changing a set point for an actual pressure ratio that includes a pressure of the turbine cooling air, and adjusting the bypass flow in response to the changed set point to cause the actual pressure ratio to approach the changed set point. | 05-17-2012 |
20120125010 | Spring loaded pressure relief door - A pressure relief apparatus for an engine compartment having a wall and an opening in the wall is provided. The pressure relief apparatus includes a door panel arranged in blocking relationship to the opening in the wall of the engine compartment, a plurality of hinges coupling the door panel to an inner surface of the wall of the engine compartment, a spring assembly coupled between each hinge and the inner surface of the wall, wherein the spring assembly includes a canister having an open end and a spring element housed within the canister, and a support fitting mounted to the engine compartment wall and having a portion extending toward the open end of the canister for applying a compressive force on the spring element. | 05-24-2012 |
20120137703 | METHOD FOR OPERATING AN AIR-STAGED DIFFUSION NOZZLE - A method is provided for operating an air-staged diffusion nozzle for a gas turbine combustor to cool the nozzle tip and improve mixing of gas fuel and air within a downstream burner space. Air is mixed with the gas-fuel in an outer swirler and expanded in a downstream burner tube space. Compressed air from a cooling air cavity in the nozzle flows through an inner swirler, passing downstream from the tip of the nozzle to the burner tube space, cooling the nozzle tip and improving the mixing of the gas-fuel with air, thereby reducing emissions from the gas turbine and reducing soot formation in startup. Direction and rotation of the discharged air from the nozzle tip into the burner space may be arranged to promote nozzle tip cooling and gas-fuel mixing with air. | 06-07-2012 |
20120159961 | GAS TURBINE ENGINE HEAT EXCHANGER - A gas turbine engine having a heat exchanger is disclosed. In one form the gas turbine engine includes a particle separator that can be used to separate particles or foreign objects and create a dirty flow and a clean flow. A blower can be used to discharge the particles or foreign objects from the separator. The heat exchanger includes a relatively warm flow path from a downstream region of a compressor and a relatively cool flow path from an upstream region of the compressor. The relatively cool flow path is merged with the dirty flow. In another embodiment, the gas turbine engine is a turbofan and the relatively cool flow path is merged with a bypass flow. In one embodiment of the engine the relatively warm flow path, after having exchanged heat with the relatively cool flow path is delivered to a working component without passing through a turbomachinery component. | 06-28-2012 |
20120167586 | Fuel Nozzle Passive Purge Cap Flow - A cooling circuit for a fuel nozzle in a gas turbine includes an end cap cavity receiving passive purge flow from a compressor of the turbine, and fuel nozzle swozzles disposed in a swozzle shroud that impart swirl to incoming fuel and air. Purge slots are formed in the swozzle shroud and through the fuel nozzle swozzles in fluid communication with the end cap cavity. The purge slots are positioned upstream of a quat fuel injection passage, and the passive purge flow enters fuel nozzle tip cavities of the fuel nozzle to provide tip cooling and tip purge volume without mixing the passive purge flow with quat fuel. | 07-05-2012 |
20120167587 | GAS TURBINE ENGINE WITH BLEED AIR SYSTEM - One embodiment of the present invention is a unique gas turbine engine. Another embodiment of the present invention is also a unique gas turbine engine. A further embodiment is a unique method for operating a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and bleed air systems therefor. 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 |
20120186267 | TURBINE INTEGRATED BLEED SYSTEM AND METHOD FOR A GAS TURBINE ENGINE - A bleed system for a gas turbine engine includes: (a) a bleed air turbine having a turbine inlet adapted to be coupled to a source of compressor bleed air at a first pressure; (b) a bleed air compressor mechanically coupled to the bleed air turbine, and having a compressor inlet adapted to be coupled to a source of fan discharge air at a second pressure substantially lower than the first pressure; and (c) a mixing duct coupled to a turbine exit of the bleed air turbine and to a compressor exit of the bleed air compressor. | 07-26-2012 |
20120227414 | GAS TURBINE ENGINE SWIRLED COOLING AIR - A gas turbine engine has in flow series a compressor section, a combustor, and a turbine section. The engine includes a turbine section rotor disc, and a stationary wall forward of a front face or rearward of a rear face of the rotor disc. The wall defines a cavity between the stationary wall and the rotor disc, and has a plurality of air entry nozzles through which cooling air can be delivered into the cavity at an inlet swirl angle. The engine further includes a cooling air supply arrangement which accepts a flow of compressed air and supplies the compressed air to the nozzles for delivery into the cavity. The cooling air supply arrangement and the nozzles are configured such that the inlet swirl angle of the air delivered into the cavity can be varied between a first inlet swirl angle and a second inlet swirl angle. | 09-13-2012 |
20120240594 | METHOD AND APPARATUS FOR PROTECTING AIRCRAFT ENGINES AGAINST ICING - An aircraft engine generates engine power by burning hydrocarbon fuel such as Jet-A. A minute quantity of the fuel is burned in such a manner as to generate no engine power, and the heat generated by the burning fuel is used to protect a region of a surface of a component of an aircraft. In one application, burner assemblies are located inside the splitter of a turbofan engine and the heat generated is used to deice or anti-ice the splitter and the inlet guide vanes of the engine. In another application, burner assemblies are located in an engine nacelle to deice or anti-ice the leading edge of the nacelle. | 09-27-2012 |
20120247120 | APPARATUS AND METHOD FOR COOLING A COMBUSTOR - A combustor includes a combustion chamber and an interior wall circumferentially surrounding at least a portion of the combustion chamber and defining an exterior surface. A plurality of turbulators are on the exterior surface. The combustor further includes means for preferentially directing fluid flow across a predetermined position of the turbulators. A method for cooling a combustion chamber includes locating a plurality of turbulators to an exterior surface of the combustion chamber and preferentially directing fluid flow across a predetermined position of the plurality of turbulators. | 10-04-2012 |
20120304662 | FUEL AIR HEAT EXCHANGER - A gas turbine engine with a fuel air heat exchanger located in the high pressure plenum. The heat exchanger includes at least one air conduit and at least one fuel conduit in heat exchange relationship with one another, with a fuel flow communication between a fuel source and fuel distribution members of the combustor being provided at least partly through the at least one fuel conduit, and the at least one air conduit defining a fluid flow communication between the high pressure plenum and an engine component to be cooled by the compressed air. | 12-06-2012 |
20130036747 | METHOD FOR COOLING A GAS TURBINE PLANT AND GAS TURBINE PLANT FOR IMPLEMENTING THE METHOD - A method is provided for operating a gas turbine plant, in which compressed air is extracted from a compressor and for cooling is directed in an internal cooling passage through thermally loaded components to the combustion chamber and/or to the turbine, then re-cooled, and added to the compressor main flow in the compressor. At least a portion of the recirculated air is supersaturated, or partially saturated, with drops of water during or before recirculation into the compressor and a cooling mist is created. A gas turbine plant is provided with a closed cooling circuit which, for implementing the method, includes an injection arrangement for introducing water into the recirculated cooling air. | 02-14-2013 |
20130042627 | COMBUSTION CHAMBER HEAD OF A GAS TURBINE WITH COOLING AND DAMPING FUNCTIONS - A combustion chamber head of a gas turbine has a substantially annular combustion chamber outer wall | 02-21-2013 |
20130055724 | GAS TURBINE ENGINE AIR CYCLE SYSTEM - An air cycle system for a gas turbine engine includes a compressor, a turbine and a heat exchanger fluidly connected between the compressor and the turbine. A fluid source communicates a fluid through the heat exchanger. The heat exchanger exchanges heat between the fluid and an airflow communicated through the heat exchanger from the compressor to provide a conditioned airflow. | 03-07-2013 |
20130067928 | METHOD FOR OPERATING A GAS TURBINE PLANT AND GAS TURBINE PLANT FOR IMPLEMENTING THE METHOD - A method is provided for operating a gas turbine plant including a compressor, which on an inlet side inducts intake air and compresses it, providing compressor exit air on a discharge side. The plant also includes a combustion chamber where fuel is combusted, using compressor exit air, forming a hot gas; and a turbine, where the hot gas is expanded, performing work. The method includes extracting compressed air from the compressor, directing it as cooling air flow into the combustion chamber and/or into the turbine for cooling thermally loaded components. The method also includes controlling at least one cooling air flow, for achieving specific operating targets, using a control element depending on an operating target. A gas turbine plant is also provided having at least one control element for cooling air flow control, and a gas turbine controller which controls the gas turbine plant based on selectable control parameter sets. | 03-21-2013 |
20130086920 | COMBUSTOR AND METHOD FOR SUPPLYING FLOW TO A COMBUSTOR - A device for supplying flow to a combustor includes a flow sleeve configured to circumferentially surround the combustor, wherein the flow sleeve defines a first annular passage around the combustor. A first section of the first annular passage converges at a first convergence rate. A second section of the first annular passage downstream from the first section converges at a second convergence rate that is less than the first convergence rate. A method for supplying flow to a combustor includes flowing a first portion of a working fluid substantially axially through a first annular passage, converging the first annular passage at a first convergence rate, and converging the first annular passage at a second convergence rate downstream from the first convergence rate, wherein the second convergence rate is less than the first convergence rate. | 04-11-2013 |
20130086921 | COMBUSTOR AND METHOD FOR SUPPLYING FLOW TO A COMBUSTOR - A device for supplying flow across a combustor includes an axial fluid injector configured to circumferentially surround at least a portion of the combustor. An inner annular passage extends through the axial fluid injector and provides fluid communication through the axial fluid injector and into a first annular passage that surrounds the combustor. An outer annular passage extends through the axial fluid injector radially outward from the inner annular passage and provides axial flow into the first annular passage. A method for supplying flow to a combustor includes flowing a first portion of a working fluid through a first axial flow path and flowing a second portion of the working fluid through a second axial flow path. | 04-11-2013 |
20130091858 | EFFUSION COOLED NOZZLE AND RELATED METHOD - A fuel nozzle for a turbine combustor includes a nozzle head configured to supply a fuel/air mixture to a burner tube attached to said nozzle head and extending downstream of the nozzle head. The burner tube is provided with plural holes for introducing a fluid into the burner tube to thereby treat (e.g., cool) an interior wall of the burner tube by effusion. | 04-18-2013 |
20130104564 | ACTIVE CLEARANCE CONTROL SYSTEM AND METHOD FOR GAS TURBINE | 05-02-2013 |
20130111921 | METHOD FOR CONTROLLING GAS TURBINE ROTOR TEMPERATURE DURING PERIODS OF EXTENDED DOWNTIME - A method for warming the rotor of a gas turbine during extended periods of downtime comprising feeding ambient air to an air blower; extracting compressed air from the air blower; feeding a portion of the compressed air to one side of a heat exchanger and steam (typically saturated) from e.g. a gas turbine heat recovery steam generator; passing the resulting heated air stream from the exchanger into and through into defined flow channels formed within the rotor; continuously monitoring the air temperature inside the rotor; and controlling the amount of air and steam fed to the heat exchanger using a feedback control loop that controls the amount of air and steam feeds to the exchanger and/or adjusts the flow rate of heated air stream into the rotor. | 05-09-2013 |
20130133334 | COOLING SYSTEM FOR ENGINE AND AIRCRAFT AIR - The combined cooling system uses a single heat exchanger to cool both engine air for use in an engine system and aircraft air for use in an aircraft system. More particularly, a bleed air path leads from the compressor stage to the heat exchanger where it is placed in thermal exchange contact with a flow of cooling air coming from a cooling path. From an outlet end of the heat exchanger, the bleed air splits into two paths: an aircraft air path leading to at least one aircraft system such as an Environmental control system (ECS), a wind de-icing system or the like, and an engine air path leading to at least one engine system such as a buffer air system for pressurizing the bearing cavities. | 05-30-2013 |
20130139521 | ON BOARD INERT GAS GENERATION SYSTEM - An on board inert gas generation system for an aircraft receives air from a relatively low pressure source such as low pressure engine bleed air or ram air and passes it to a positive displacement rotary compressor to increase the pressure thereof to be suitable for supply to an air separation module. The speed of the positive displacement compressor may be adjusted across a wide range in order to provide efficient operation in cruise and descent phases of aircraft flight. | 06-06-2013 |
20130152601 | CONTROLLER - A gas turbine engine has, in flow series, a compressor section, a combustor, and a turbine section. The gas turbine engine further has a system for cooling the turbine section and providing tip clearance control between turbine blades and circumferentially distributed segments forming an annular shroud surrounding the blades outer tips. The turbine section cooling sub-system diverts a first cooling air flow, regulated by a first valve arrangement, from the compressor section to a heat exchanger and then to the turbine section to cool its components. The tip clearance control sub-system supplies a second cooling air flow, regulated by a second valve arrangement, to an engine case where the segments are mounted, which regulates thermal expansion of the case and controls the clearance between the segments and outer tips. The system further includes a closed-loop controller which issues demand signals to the first and second valve arrangements. | 06-20-2013 |
20130152602 | CONTROLLER - A gas turbine engine has, in flow series, a compressor section, a combustor, and a turbine section. The engine further has a cooling system which diverts a cooling air flow received from the compressor section to a heat exchanger and then to the turbine section to cool a component thereof. The cooling air flow by-passes the combustor and is cooled in the heat exchanger. The cooling system has a valve arrangement which regulates the cooling air flow. The engine further has a closed-loop controller which estimates and/or measures one or more temperatures of the cooled component, compares values derived from the estimated and/or measured temperatures with one or more corresponding targets, and issues a demand signal to the valve arrangement based on the comparison and a value of the demand signal at a previous time step. | 06-20-2013 |
20130174572 | METHOD FOR COOLING A THERMAL PROTECTION FLOOR OF AN AFT AERODYNAMIC FAIRING OF A STRUCTURE FOR MOUNTING AN AIRCRAFT PROPULSION SYSTEM - A propulsion system for an aircraft, including a dual-flow turbojet and a mounting structure for mounting this turbojet on the wing surface or on the fuselage of an aircraft. The mounting structure includes an aft aerodynamic fairing including a thermal protection floor to protect the mounting structure from the heat of a primary airstream channelled by an exhaust nozzle of the turbojet, as well as an air inlet provided in a longitudinal aerodynamic wall washed by a secondary airstream of the turbojet and delimiting together with an other similar wall a cavity isolated from the secondary airstream for extracting a cooling airstream from the secondary airstream, and air circulation means fed by the air inlet and having at least one outlet aperture emerging in a space between the thermal protection floor and the exhaust nozzle. | 07-11-2013 |
20130186100 | SMALL ENGINE COOLED COOLING AIR SYSTEM - A system for managing thermal transfer in an aircraft includes a fuel stabilization unit, a fuel-air heat exchanger, and a turbine. The fuel-air heat exchanger is located downstream from the fuel stabilization unit. The fuel-air heat exchanger places deoxygenized fuel in a heat exchange relationship with compressor bleed air to produce heated deoxygenized fuel and cooled bleed air. The turbine is operationally connected to the engine compressor and receives cooled bleed air from the fuel-air heat exchanger. | 07-25-2013 |
20130192249 | Gas Turbine Engine System and Method for Controlling a Temperature of a Conduit in a Gas Turbine Engine System - According to one aspect of the invention, an engine system includes one or more fuel circuits configured to provide gaseous fuel from a fuel source for combustion. The system further comprises a conduit in fluid communication with the one or more fuel circuits, wherein the conduit is configured to receive a first gaseous fuel flow and an oxidant flow for a reaction within the conduit, wherein a temperature of the conduit is controlled by a second gaseous fuel flow along an outer wall of the conduit. | 08-01-2013 |
20130192250 | GAS TURBINE ENGINE BUFFER SYSTEM - A gas turbine engine includes a buffer system that can communicate buffer supply air to a portion of the gas turbine engine. The buffer system can include a first circuit and a second circuit. The first circuit selects between a first bleed air supply having a first pressure and a second bleed air supply having a second pressure that is greater than the first pressure to render a first buffer supply air having an intermediate pressure. The second circuit selects between a third bleed air supply and a fourth bleed air supply to communicate a second buffer supply air. | 08-01-2013 |
20130192251 | BUFFER SYSTEM THAT COMMUNICATES BUFFER SUPPLY AIR TO ONE OR MORE PORTIONS OF A GAS TURBINE ENGINE - A gas turbine engine includes a buffer system that can communicate a buffer supply air to a portion of the gas turbine engine. The buffer system includes a first bleed air supply having a first pressure, a second bleed air supply having a second pressure that is greater than the first pressure, and an ejector that selectively augments the first bleed air supply to prepare the buffer supply air for communication to the portion of the gas turbine engine. | 08-01-2013 |
20130192252 | GAS TURBINE ENGINE BUFFER SYSTEM - A gas turbine engine includes a buffer system that communicates a buffer cooling air to at least one bearing structure and at least one shaft of the gas turbine engine. The buffer system includes a first bleed air supply and a conditioning device that conditions the first bleed air supply to render the first buffer supply air at an acceptable temperature to pressurize the at least one bearing structure and cool the at least one shaft. | 08-01-2013 |
20130192253 | GAS TURBINE ENGINE BUFFER SYSTEM PROVIDING ZONED VENTILATION - A gas turbine engine includes a first zone and a second zone downstream from the first zone. A buffer system can communicate a buffer cooling air to at least the first zone. A bleed source can communicate a bleed air to the second zone. | 08-01-2013 |
20130213054 | Gas Turbine Inlet System with Solid-State Heat Pump - A gas turbine inlet system includes a first passage that delivers inlet air to a compressor, a second passage that delivers fuel to a combustor, and a heat pump that transfers heat from the inlet air to the fuel by consuming electric power. The heat transfer causes the turbine inlet air temperature to drop and fuel temperature to increase with favorable effects on both turbine output and heat rate. | 08-22-2013 |
20130219915 | FUEL AIR HEAT EXCHANGER - A fuel air heat exchanger for a gas turbine engine having fuel and air conduits in heat exchange relationship with one another, and a distribution conduit in heat exchange relationship with a component to be cooled. The distribution conduit is in fluid communication with the outlet of each air conduit. The heat exchanger also includes a secondary air inlet in fluid communication with the distribution conduit and a flow selection member selectively movable between first and second configurations. In the first configuration, the flow selection member closes the fluid communication between the secondary inlet and the distribution conduit. In the second configuration, the flow selection member opens the fluid communication between the secondary air inlet and the distribution conduit. | 08-29-2013 |
20130219916 | APPARATUS AND METHOD FOR CONDITIONING AIR RECEIVED BY A POWER GENERATION SYSTEM - According to one aspect of the invention, a method for conditioning air received by a power generation system includes flowing ventilation air through a turbine system to control a temperature of the turbine system and receiving the ventilation air from the turbine system and mixing the ventilation with an ambient air to form an intake air to be directed to a compressor, wherein a temperature of the ventilation air is greater than the ambient air. | 08-29-2013 |
20130219917 | GAS TURBINE ENGINE BUFFER COOLING SYSTEM - A gas turbine engine includes a heat exchanger, a bearing compartment, and a nozzle assembly in fluid communication with the bearing compartment. The heat exchanger exchanges heat with a bleed airflow to provide a conditioned airflow. The bearing compartment is in fluid communication with the heat exchanger. A first passageway communicates the conditioned airflow from the heat exchanger to the bearing compartment. A second passageway communicates the conditioned airflow from the bearing compartment to the nozzle assembly. | 08-29-2013 |
20130219918 | BUFFER COOLING SYSTEM PROVIDING GAS TURBINE ENGINE ARCHITECTURE COOLING - A gas turbine engine includes a buffer cooling system having a first heat exchanger, a first passageway, a second passageway and a third passageway. The first heat exchanger exchanges heat with a bleed airflow to provide a conditioned airflow. The first passageway communicates a first portion of the conditioned airflow to a high pressure compressor of the gas turbine engine, the second passageway communicates a second portion of the conditioned airflow to a high pressure turbine of the gas turbine engine, and the third passageway communicates a third portion of the conditioned airflow to a low pressure turbine of the gas turbine engine. | 08-29-2013 |
20130219919 | GAS TURBINE ENGINE BUFFER COOLING SYSTEM - A gas turbine engine includes a heat exchanger, a mid-turbine frame, a passageway that extends through at least a portion of the mid-turbine frame and a first nozzle assembly. The heat exchanger exchanges heat with a bleed airflow to provide a conditioned airflow. The mid-turbine frame is in fluid communication with the heat exchanger. The conditioned airflow is communicated through the passageway and is received by the first nozzle assembly to condition gas turbine engine hardware. | 08-29-2013 |
20130219920 | GAS TURBINE ENGINE COOLING SYSTEM - A gas turbine engine includes a heat exchanger, a diffuser case, a passageway and a nozzle assembly. The heat exchanger exchanges heat with a bleed airflow to provide a conditioned airflow. The diffuser case includes a plenum that receives the conditioned airflow. The passageway is fluidly connected between the heat exchanger and the diffuser case, and the conditioned airflow is communicated through the passageway and into the plenum. The nozzle assembly is in fluid communication with the plenum of the diffuser case to receive the conditioned airflow from the plenum. | 08-29-2013 |
20130247583 | GAS TURBINE ENGINES AND SYSTEMS AND METHODS FOR REMOVING PARTICULATE MATTER THEREFROM DURING OPERATION - Systems and methods are provided for removing particulate matter from gas turbine engines during operation are provided. Gas turbine engines are also provided. The particulate matter is suspended in a primary gas flow stream passing through an engine flowpath. A flowpath surface of the engine flowpath is electrostatically charged to a first polarity to thereby impart an electrostatic charge of the first polarity to the particulate matter. A bleed discharge duct is electrostatically charged to a second polarity and intersects the engine flowpath to define a bleed air flowpath. The second polarity is opposite to the first polarity. A bleed port is in fluid communication with the bleed discharge duct and has an outlet exterior of the gas turbine engine. | 09-26-2013 |
20130247584 | ACTIVE CONTROL OF COMPRESSOR EXTRACTION FLOWS USED TO COOL A TURBINE EXHAUST FRAME - A gas turbine includes at least one combustor and an exhaust frame; a compressor adapted to supply air to the combustor and to supply bleed air to the exhaust frame. A cooling air supply duct is arranged to supply ambient air to the exhaust frame and at least one ejector is. arranged to supply the bleed air to the cooling air supply duct upstream of the exhaust frame. A control valve is configured to control the supply of compressor bleed air to the cooling air supply duct and to the exhaust frame as a function of turbine exhaust temperature and/or turbine load conditions and cooling requirements at the various turbine load conditions. | 09-26-2013 |
20130255273 | Systems and Methods for Gas Turbine Combustors - A gas turbine system includes a compressor operative to output an airstream and a diffuser having an inlet to receive the airstream and an outlet to output the airstream. The outlet has an area larger than the inlet to diffuse the airstream. The gas turbine system also includes a fuel nozzle operative to receive fuel and emit the fuel in a combustor and at least one bleed duct having an inlet between the compressor and the outlet of the diffuser. The at least one bleed duct is operative to direct bleed air from downstream of the compressor to the fuel nozzle. | 10-03-2013 |
20130283813 | GAS TURBINE COMPRESSOR WITH BLEED PATH - A gas turbine engine includes a compressor for generating compressed air. The compressor includes a rotor defined by a plurality of axial disks including a first disk and a second disk. A first row of blades extends radially outwardly from the first disk, and a second row of blades extends radially outwardly from the second disk. A row of cantilevered vanes is located at an axial location between the first row of blades and the second row of blades. A bleed path extends at least partially through the second disk and includes an entrance at an axial location between the first row of blades and at least a portion of the row of cantilevered vanes. The entrance communicates with a compressed air flowpath through the compressor. | 10-31-2013 |
20130283814 | TURBINE COOLING SYSTEM - A system includes a turbine. The turbine includes a rotor, a stator, and a cooling insert. The rotor includes multiple turbine blades. The stator surrounds the rotor and includes an inner wall surrounding the multiple turbine blades, an outer wall surrounding the inner wall, and a cooling chamber between the inner wall and the outer wall. The cooling insert extends through an opening in the outer wall into the cooling chamber. The insert includes a side wall extending around an axis, an end wall crosswise to the axis, a set of lateral ports extending through the side wall, and end ports extending through the end wall. The cooling insert is configured to direct a cooling fluid through the set of lateral ports and end ports into the cooling chamber. | 10-31-2013 |
20130283815 | INTEGRAL COOLING FOR SERVO VALVE - A cooling structure for a servo valve includes a shroud to enclose at least a portion of the servo valve; and a base connected to the shroud to define a cooling chamber surrounding the servo valve, the base including an inlet port to receive cooling air, a flow channel connecting to the inlet port and a plurality of flow passages connecting the flow channel to the cooling chamber to allow cooling air flow from the inlet port into the cooling chamber. | 10-31-2013 |
20130291553 | INTEGRATED THERMAL PROTECTION AND LEAKAGE REDUCTION IN A SUPERSONIC AIR INTAKE SYSTEM - An air intake system suitable for a supersonic vehicle is disclosed. The system includes a channel comprising an inlet and a side wall and a plenum coupled to the side wall. The plenum is configured to accept a flow of coolant. In certain embodiments, the coolant is the waste coolant from an on-board electronics cooling system. The system also includes a porous region in the side wall configured to allow a flow of bleed air from the channel through the porous region of the side wall into the plenum so as to aid the transition to supersonic flow. In certain embodiments, the flow of the bleed air is reduced at supersonic speeds by pressurization of the plenum with the coolant. | 11-07-2013 |
20130318995 | SYSTEMS AND METHODS FOR DIRECTING COOLING FLOW INTO THE SURGE PLENUM OF AN EXHAUST EDUCTOR COOLING SYSTEM - An eductor assembly comprises a primary nozzle configured to discharge turbine exhaust gas therefrom. The eductor assembly further comprises a cooler plenum having an inlet and an outlet and a surge plenum at least partially surrounding the cooler plenum and the nozzle, the surge plenum for conducting a surge flow. Cooling air flows through a vent between the cooler plenum and the surge plenum when there is no surge flow. | 12-05-2013 |
20130318996 | COOLING ASSEMBLY FOR A BUCKET OF A TURBINE SYSTEM AND METHOD OF COOLING - A cooling assembly for a bucket of a turbine system includes a shroud assembly operably coupled to an outer casing of a turbine section. Also included is an airfoil having at least one cavity, wherein the at least one cavity is configured to receive a cooling flow from a cooling source through at least one channel disposed within the shroud assembly. | 12-05-2013 |
20130327055 | TRANSITION PIECE FOR A GAS TURBOMACHINE - A transition piece for a gas turbomachine includes a body having an outer surface and an inner surface that defines a flow duct, a plurality of openings that extend through the body, and an active control element provided at one or more of the plurality of openings. The active control element is configured and disposed to selectively establish a dimension of the one or more of the plurality of openings. | 12-12-2013 |
20130327056 | COMBUSTOR LINER WITH DECREASED LINER COOLING - A shell for a combustor liner includes a cold side, a hot side, a row of cooling holes and a jet wall. The jet wall projects from the hot side for creating a wall shear jet of increased velocity cooling flow in a tangential direction away from the row of cooling holes and along an adjacent heat shield cold side wall. | 12-12-2013 |
20130327057 | COMBUSTOR LINER WITH IMPROVED FILM COOLING - A heat shield for a combustor liner includes first linear film cooling slots through the heat shield and second linear film cooling slots through the heat shield. The first linear film cooling slots are run in a row and each of the first linear film cooling slots is angled from the row in a first direction. The second linear film cooling slots also run in the row and each of the second linear film cooling slots is angled from the row in a second direction opposite the first direction. The second linear film cooling slots alternate with the first linear film cooling slots in the row. The first and second linear film cooling slots are connected to form a single, multi-cornered film cooling slot. | 12-12-2013 |
20130333390 | Cooling air inlet, engine bleed air system and method for operating a cooling air inlet - A cooling air inlet which is particularly suitable for use in an engine bleed air system of an aircraft air conditioning system. The cooling air inlet comprises a cooling air inlet opening which opens into an inlet portion of a cooling air line. A cooling air inlet valve which is positionable in a first open position or a second open position is arranged in the inlet portion of the cooling air line. The cooling air inlet valve is adapted to effect the build-up of a first cooling air admission pressure in the inlet portion of the cooling air line in its first open position and the build-up of a second cooling air admission pressure in the inlet portion of the cooling air line in its second open position. The second cooling air admission pressure is greater than the first cooling air inlet pressure. | 12-19-2013 |
20140000278 | TURBOMACHINE THERMAL ENERGY EXCHANGE | 01-02-2014 |
20140000279 | SYSTEM AND METHOD FOR OPERATING A PRECOOLER IN AN AIRCRAFT | 01-02-2014 |
20140026588 | SYSTEM AND METHOD FOR RECIRCULATING AND RECOVERING ENERGY FROM COMPRESSOR DISCHARGE BLEED AIR - A system includes a gas turbine, an inlet bleed circuit, and a controller. The gas turbine includes a compressor and a turbine. The compressor is configured to produce pressurized air and bleed air. The turbine is configured to produce a first output. The inlet bleed circuit includes a turbo-expander configured to produce a second output from a non-zero first portion of the bleed air. The inlet bleed circuit is also configured to direct a part of the bleed air to an inlet of the compressor. The controller is configured to adjust the gas turbine and the inlet bleed circuit to control the second output of the turbo-expander. | 01-30-2014 |
20140053571 | SEAL FOR A PERFORATED PLATE - A cooling circuit of a gas turbine passes an airflow through a combustor section that includes a plurality of mixing tubes for transporting a fuel/air mixture and a perforated plate including a plurality of impingement holes and a plurality of tube holes for accommodating the mixing tubes. The tube holes and the mixing tubes form a plurality of annulus areas between the perforated plate and the mixing tubes. The impingement holes and the annulus areas are configured to pass the airflow through the perforated plate. A flow management device modifies an effective size of the annulus areas to control a distribution of the airflow through the impingement holes and the annulus areas of the perforated plate to enhance cooling efficiency. | 02-27-2014 |
20140060073 | MULTIPLE POINT OVERBOARD EXTRACTOR FOR GAS TURBINE - The disclosed overboard extractor provides overboard extraction of compressor fluid at multiple extraction points. The overboard extractor includes multiple extraction manifolds and a delivery manifold for extracting the compressor fluid from the multiple extraction points. The overboard extractor also includes extraction valves and a delivery valve to control the extraction flow through each extraction manifold. The overboard extractor can operate in a delivery mode, a closed loop mode, or a mixed mode. | 03-06-2014 |
20140075955 | EASILY ADAPTABLE COMPRESSOR BLEED SYSTEM DOWNSTREAM OF A VANE PLATFORM - Discrete bleed behind stator vane platform is provided. The discrete bleed behind stator vane platform relates to a system for bleeding off a working fluid from an inner volume (Vi) of a turbo-machine. The system includes a vane carrier with an annular rail and a vane device comprising at least one vane element, a vane platform and a vane root. The vane element is mounted to the vane platform and the vane root is mounted to the annular rail. A first annular cavity is formed between the vane platform and the annular rail, wherein an annular gap is formed between an edge of the vane platform and the vane carrier such that a part of the working fluid of the turbo-machine is bleedable through the annular gap into the first annular cavity. A second annular cavity is formed between the vane root, the annular rail and the vane carrier, wherein at least one inlet hole is formed into the annular rail for coupling the first annular cavity and the second annular cavity. | 03-20-2014 |
20140083112 | Cooled Combustor Liner Grommet - A combustor liner grommet is disclosed. The grommet may include a peripheral wall defining a hole in a combustor liner and further including at least one cooling air flow channel. The cooling air flow channel in the grommet wall may be a slot or a hole. The channel may increase cooling flow to the grommet and the combustor liner around the grommet to prevent cracking from heat stress. | 03-27-2014 |
20140096534 | Low Profile Compressor Bleed Air-Oil Coolers - An air-oil cooler (AOC) for a gas turbine engine is disclosed. The AOC may comprise an oil inlet, an oil outlet, and heat exchange elements between the oil inlet and the oil outlet. The AOC may be longitudinally positioned between a fan and a V-groove of the engine and radially spaced between a low pressure compressor and a low pressure compressor panel. A gas turbine engine comprising an AOC is disclosed. The AOC of the engine may comprise an oil inlet, an oil outlet, and heat exchange elements between the oil inlet and the oil outlet. The AOC of the engine may be longitudinally positioned between a fan and a V-groove of the engine and radially spaced between a low pressure compressor and a low pressure compressor panel. A method of operating an AOC for use on a gas turbine engine is also disclosed. | 04-10-2014 |
20140109590 | GAS TURBINE THERMAL CONTROL AND RELATED METHOD - Thermal control is provided for a gas turbine casing by supplying thermal control gas from a compressor to a space between an outer casing and an inner casing, and transferring the thermal control gas from the space through the opening in the inner casing via a plurality of holes defined through a plate attached to an outer surface of the inner casing. The holes are arranged with a predetermined non-uniform distribution corresponding to a desired preferential impingement pattern for providing non-uniform heat transfer. A gas turbine thermal control assembly includes structure providing preferential heat transfer from the inner casing during operation of the gas turbine via a thermal control gas flow path from radially outside of the inner casing into the interior of the gas turbine. | 04-24-2014 |
20140123674 | Inlet Bleed Heat System with Integrated Air Knife/Silencer Panels - The present application provides an inlet bleed heat system for supplying a flow of bleed air to a flow of incoming air into a compressor of a gas turbine engine. The inlet bleed heat system may include an air knife and a silencer panel. The air knife may include a compressor bleed air port in communication with the flow of bleed air and a discharge gap to discharge the flow of bleed air into the flow of incoming air. The air knife may and the silencer panel may form an integrated air knife/silencer panel. | 05-08-2014 |
20140123675 | AIR INJECTION SYSTEM IN A GAS TURBINE ENGINE - An air injection system for use in a gas turbine engine includes at least one outlet port through which air is extracted from the engine only during less than full load operation, at least one rotor cooling pipe, which is used to inject the air extracted from the outlet port(s) into a rotor chamber, a piping system that provides fluid communication between the one outlet port(s) and the rotor cooling pipe(s), a blower system for extracting air from the engine through the outlet port(s) and for conveying the extracted air through the piping system and the rotor cooling pipe(s) into the rotor chamber, and a valve system. The valve system is closed during full load engine operation to prevent air from passing through the piping system, and open during less than full load engine operation to allow air to pass through the piping system. | 05-08-2014 |
20140123676 | GAS TURBINE ENGINE END-WALL COMPONENT - An end-wall component of the mainstream gas annulus of a gas turbine engine has a cooling arrangement including one or more circumferentially extending rows of ballistic cooling holes through which, in use, dilution cooling air is jetted into the mainstream gas to reduce the mainstream gas temperature adjacent the end-wall. A portion of the cooling holes are first cooling holes angled such that the direction of the dilution cooling air jetted therethrough has, on entry into the mainstream gas annulus, a component in one tangential direction. A portion of the cooling holes are second cooling holes angled such that the direction of the dilution cooling air jetted therethrough has, on entry into the mainstream gas annulus, a component in the opposite tangential direction. The first and second cooling holes are arranged such that the cooling air from jets having entry components in opposing tangential directions collide and coalesce. | 05-08-2014 |
20140130510 | SYSTEMS AND METHODS FOR DIRECTING COOLING FLOW INTO THE SURGE PLENUM OF AN EXHAUST EDUCTOR COOLING SYSTEM - A method and apparatus is provided for cooling the external surface of an aircraft APU eductor assembly. A processor is configured to open a surge valve by a predetermined amount when the surge valve is closed and the temperature of the exhaust gas exceeds a predetermined temperature in order to cool the surge plenum surfaces. | 05-15-2014 |
20140144154 | GAS TURBINE ENGINE WITH BEARING BUFFER AIR FLOW AND METHOD - The gas turbine engine has a bleed air aperture formed in the radially outer wall upstream from the combustor and a bearing cavity formed within the radially inner wall, at least two bearing seals enclosing at least one bearing in the bearing cavity and separating the bearing cavity from associated buffer air entry points, an oil supply system including oil paths leading to each of the bearings; a buffer air supply system including buffer air paths leading to each of the entry points. | 05-29-2014 |
20140144155 | GAS TURBINE COMPRISING A HEAT SHIELD AND METHOD OF OPERATION - A turbine has a first and second inner wall, an inner casing and a shield element. The first and second inner walls are mounted to the inner casing such that an inner volume for working fluid is separated from an outer volume for cooling fluid. The first and second inner walls and inner casing are arranged to form a cavity in outer volume. The shield element is arranged inside cavity to separate cavity in an inner and outer region in a radially outer position to the inner region. A gap is formed between the first and second inner walls for working fluid between the inner volume and inner region. The shield element is arranged in cavity to form a fluid inlet for injecting cooling fluid from outer region to inner region for generating a predefined recirculation of working fluid and cooling fluid inside inner region. | 05-29-2014 |
20140165588 | TURBO COMPRESSOR FOR BLEED AIR - A disclosed bleed air system utilizes high pressure air from a high pressure compressor to drive the turbo compressor to increase a pressure of bleed air drawn from the low pressure compressor. Air drawn from the low pressure compressor is at a lower temperature and pressure than that encountered from the high pressure compressor. The turbo compressor increases the pressure of airflow and provides that airflow into the main bleed air passage to be communicated to systems utilizing the bleed air. | 06-19-2014 |
20140208768 | COOLANT SUPPLY SYSTEM - A cooling system for a gas turbine engine ( | 07-31-2014 |
20140208769 | INTEGRATED INDUCER HEAT EXCHANGER FOR GAS TURBINES - An integrated inducer heat exchanger is provided. The integrated inducer heat exchanger includes multiple airfoil devices disposed in an annular array within an inner circular casing and an outer circular casing forming multiple passages for allowing a flow of fluid from a forward side to an aft side of the integrated inducer heat exchanger. The integrated inducer heat exchanger also includes multiple annular manifolds arranged about the outer circular casing configured for supplying a flow of coolant at low temperature from one or more coolant sources and returning the flow of coolant at high temperature to the one or more coolant sources via an external heat exchanger for dissipating heat and multiple transfer tubes connecting the multiple annular manifolds with the multiple airfoil devices for transferring the flow of coolant within the airfoil devices for exchanging heat between the coolant and the fluid passing through the multiple passages. | 07-31-2014 |
20140245747 | GAS TURBINE ENGINE TWO DEGREE OF FREEDOM VARIABLE BLEED VALVE FOR ICE EXTRACTION - A gas turbine engine variable bleed apparatus includes a variable bleed valve door disposed in a bleed inlet in a transition duct, rotatable about two or more separate pivot points, operable to open and close an aft bleed slot extending outwardly from transition duct, and operable to open and close a forward bleed slot extending inwardly into transition duct. Door is operable to transition between a first position with aft bleed slot open and forward bleed slot closed to a second position with aft bleed slot closed and forward bleed slot open without fully closing door. Door is rotatable about an axis translatable between the two or more separate pivot points. Transition duct having a transition duct conical angle at least about 10 degrees greater than a booster conical angle of a booster outer shroud upstream of transition duct. | 09-04-2014 |
20140311157 | VANE CARRIER TEMPERATURE CONTROL SYSTEM IN A GAS TURBINE ENGINE - A vane carrier temperature control system for use in a gas turbine engine includes a first cooling air source, a second cooling air source, and an air temperature control system. The first cooling air source supplies a first portion of vane carrier cooling air extracted from a compressor section of the engine to a first section of a vane carrier that supports a plurality of rows of vanes within a turbine section of the engine. The second cooling air source supplies a second portion of vane carrier cooling air extracted from the compressor section to a second section of the vane carrier spaced from the first section in an axial direction defined by a direction of hot working gas flow through the turbine section. The air temperature control system controls a temperature of at least one of the first and second portions of vane carrier cooling air. | 10-23-2014 |
20140338360 | BLEED PORT RIBS FOR TURBOMACHINE CASE - An example turbomachine structure includes a case having a radially extending port, and at least one rib configured to direct flow through and past the port. | 11-20-2014 |
20150027129 | GAS TURBINE WITH ADJUSTABLE COOLING AIR SYSTEM - In order to improve the cooling of an air-cooled gas turbine in the partial load operating mode it is proposed to provide a connecting line between two cooling air lines with different pressure levels, which connecting line leads from the second cooling air line at a relative high pressure level to the first cooling air line at a relative low pressure level. In this context, a cooling device for cooling an auxiliary cooling air stream, flowing from the second cooling air line into the first cooling air line, and an adjustment element are arranged in the connecting line. In addition to a gas turbine, a method for operating such a gas turbine is the subject matter of the disclosure. | 01-29-2015 |
20150033756 | Gas Turbine with Primary and Secondary Lubricating Oil Cooler - A gas turbine including: a turbine package, a gas turbine, a ventilation system for cooling the interior of the turbine package, and a lubricating oil circuit. The lubricating oil circuit comprises a lubricating oil pump, a lubricating oil tank, a primary lubricating oil cooler, and a secondary lubricating oil cooler. The secondary lubricating oil cooler is arranged in the turbine package, in a position lower than a rotary shaft of the gas turbine. The ventilation system is arranged and designed such that at least part of a package cooling airflow contacts the secondary lubricating oil cooler to remove heat from the lubricating oil circulating in the secondary lubricating oil cooler. | 02-05-2015 |
20150047367 | COMPOSITE HEAT EXCHANGER - A heat exchanger assembly in one embodiment includes an inlet, an outlet, and a core. The inlet is configured to accept a fluid to be cooled and the outlet is configured to provide an exit for a flow of the fluid that has been cooled. The core includes at least one passageway formed therein configured for passage of the fluid to be cooled from the inlet to the outlet. The core also includes a receiving surface configured to receive a cooling flow therethrough. The core is configured to direct the cooling flow to pass by at least a portion of the at least one passageway to cool the fluid. The core is comprised of a high temperature geopolymer composite material. The composite material may be configured for use at temperatures above about 300 degrees Celsius, and may have a thermal conductivity below about 20 Watts/(meter*degree Kelvin). | 02-19-2015 |
20150047368 | SYSTEMS AND METHODS FOR CONTROLLING GAS TURBINES - Embodiments of the disclosure relate to systems and methods for controlling a gas turbine. In one embodiment, a method can be provided. Measurement data associated with the operation of a gas turbine may be received to determine if the current operating state of the gas turbine is associated with a predefined risk. Thereafter, the measurement data and one or more model operating parameters for operation of the inlet bleed heat system that minimizes the risk may be identified in order to generate one or more control signals operable to adjust the operation of the inlet bleed heat system for the gas turbine to adjust the one or more model operating parameters. | 02-19-2015 |
20150059355 | Method and System for Controlling Gas Turbine Performance With a Variable Backflow Margin - A system and method for controlling the performance of a gas turbine system is provided. A backflow margin pressure ratio for a component is determined. A modified backflow margin pressure ratio for the component is calculated based on the number of fired hours and starts. Bleed air along a first flow path is controlled based on the modified backflow margin pressure ratio. | 03-05-2015 |
20150068217 | INLET BLEED HEAT SYSTEM AND RELATED METHOD FOR A COMPACT GAS TURBINE INLET - An inlet system for a gas turbine includes an inlet air duct; a silencer disposed in the inlet air duct, the silencer including a plurality of panels with spaces between the panels; and a conduit with orifices disposed to inject inlet bleed heat into each of the spaces. A method of conditioning inlet air for a gas turbine includes flowing air through spaces between panels of a silencer in an inlet air duct of the gas turbine, and injecting inlet bleed heat through orifices and into each of the spaces. | 03-12-2015 |
20150089955 | GAS TURBINE WITH COOLING AIR COOLING SYSTEM AND METHOD FOR OPERATION OF A GAS TURBINE AT LOW PART LOAD - A gas turbine and method for operating a gas turbine, includes a compressor, a combustor, a turbine and a cooling air cooling system having at least a first cooling air line going from a first bleed of the compressor to the turbine, and at least one second cooling air line at a downstream position of the compressor relative to the first cooling air line. A heat exchanger is arranged in the second cooling air line for cooling the extracted air of higher pressure. The heat exchanger is connected with an air inlet side of the compressor such that heat is transferred in order to heat up the inlet air of the compressor. | 04-02-2015 |
20150113999 | Inlet System for a Precooler - An apparatus comprises a leading edge of an inlet. The leading edge of the inlet is positioned relative to a direction of air flow such that a total pressure of air along the leading edge of the inlet is equalized within selected tolerances. | 04-30-2015 |
20150292438 | METHOD AND APPARATUS FOR COOLING COMBUSTOR LINER IN COMBUSTOR - A method and apparatus for cooling a combustor liner in a combustor are disclosed. In one embodiment, a combustor is disclosed. The combustor includes a transition piece, and an impingement sleeve at least partially surrounding the transition piece and at least partially defining a generally annular flow path therebetween. The combustor further includes an injection sleeve mounted to one of the transition piece or the impingement sleeve and positioned radially outward of the impingement sleeve, the injection sleeve at least partially defining a flow channel configured to flow working fluid to the flow path. In another embodiment, a method for cooling a combustor liner in a combustor is disclosed. The method includes flowing a working fluid through a flow channel at least partially defined by an injection sleeve, and exhausting the working fluid from the flow channel into a flow path adjacent the combustor liner. | 10-15-2015 |
20150292742 | GAS TURBINE ENGINE COMBUSTOR BASKET WITH INVERTED PLATEFINS - A gas turbine engine combustor basket has nested outer and inner liners that are separated by a gap at their respective distal downstream ends for passage of cooling air between the liners. Radially inwardly projecting platefins formed on an inner circumferential surface of the outer liner maintain the cooling air passage gap. In some embodiments effusion cooling through holes are formed in the inner liner outer circumference, oriented in the air passage gap between the fins, so that cooling air passes through the effusion holes into the cooling air passage gap. | 10-15-2015 |
20150323186 | COOLED FUEL INJECTOR SYSTEM FOR A GAS TURBINE ENGINE AND METHOD FOR OPERATING THE SAME - A cooled fuel injector system of a combustor section of a gas turbine engine is provided. At least a part of the fuel injector system is exposed to core gas flow traveling through the engine. The cooled fuel injector system includes a source of a first cooling fluid and a fuel injector system component. The first cooling fluid is at a temperature lower than a temperature of the core gas flow proximate the fuel injector system. The fuel injector system component includes a vascular engineered structure lattice (VESL) structure, which VESL structure is in fluid communication with the source of the cooling fluid. | 11-12-2015 |
20150345393 | SYSTEMS AND METHODS FOR UTILIZING GAS TURBINE COMPARTMENT VENTILATION DISCHARGE AIR - Systems and methods for utilizing gas turbine compartment ventilation discharge air. In one embodiment, a system may include a gas turbine engine having a compressor. The system also may include a gas turbine compartment disposed about the gas turbine engine. Moreover, the system may include an inlet bleed heat (IBH) manifold in fluid communication with the compressor. The gas turbine compartment may be in fluid communication with the IBH manifold for providing the IBH manifold with ventilation discharge air from the gas turbine compartment. | 12-03-2015 |
20150345405 | GAS TURBINE ENGINE WITH ROTOR BORE HEATING - A gas turbine engine has a compressor rotor with blades and a disk. A bore is defined radially inwardly of the disk. A combustor includes a burner nozzle. A tap taps air that has been combusted in the combustor section through a valve, and into the bore of the disk. A method is also disclosed. | 12-03-2015 |
20150354455 | THERMALLY ISOLATED TURBINE SECTION FOR A GAS TURBINE ENGINE - A gas turbine engine includes a turbine section fluidly connected to a combustor by a primary flow path. The turbine section includes a first portion at a high pressure relative to a second portion. A thermally isolated cooling plenum is positioned radially inward of the primary flow path. The cooling plenum is defined by a forward wall, a shaft structure, an aft wall, and an inner diameter wall of the primary flow path. Air in the thermally isolated cooling plenum is thermally isolated from air in the primary flow path. At least one cooling air system is operable to provide cooling air to the thermally isolated cooling plenum. | 12-10-2015 |
20150354465 | TURBINE STAGE COOLING - A turbine cooling air generation system for a gas turbine engine includes a first fluid pathway connecting a compressor bleed outlet and a mixing valve, a second fluid pathway connecting the compressor bleed outlet and an input of a heat exchanger, and a third fluid pathway connecting an output of the heat exchanger and the mixing valve. The mixing valve is further connected to an input of a turbine stage active cooling system. | 12-10-2015 |
20150354819 | Combustor Cooled Quench Zone Array - In accordance with one aspect of the disclosure, a combustor is disclosed. The combustor may include a shell and a liner disposed within the shell. The combustor may further include a grommet at least partially defining a hole communicating through the shell and liner and a cooling channel communicating through the grommet. | 12-10-2015 |
20150377055 | GAS TURBINE AND OPERATING METHOD THEREOF - Provided is a gas turbine capable of achieving high-speed startup of the gas turbine through quick operation control of an ACC system during startup of the gas turbine, improving the cooling efficiency of turbine stationary components, and quickly carrying out an operation required for cat back prevention during shutdown of the gas turbine. Included are a pressurizing device ( | 12-31-2015 |
20150377056 | GAS TURBINE AND OPERATING METHOD THEREOF - Provided is a gas turbine capable of achieving high-speed startup of the gas turbine through quick operation control of an ACC system during startup of the gas turbine, improving the cooling efficiency of turbine stationary components, and quickly carrying out an operation required for cat back prevention during shutdown of the gas turbine. Included are a pressurizing device ( | 12-31-2015 |
20150377126 | Combined Gas Turbine Auxiliary Systems - The present application provides a gas turbine engine. The gas turbine engine may include a compressor, a combustor, a number of auxiliary systems, and a common auxiliary system manifold. The common auxiliary system manifold is in communication with the compressor, the combustor, and the auxiliary systems. | 12-31-2015 |
20160003192 | EXHAUST SYSTEM HAVING A FLOW PATH LINER SUPPORTED BY STRUCTURAL DUCT SEGMENTS - The exhaust system ( | 01-07-2016 |
20160025009 | METHOD OF SUPPLYING FUEL TO AN INTERNAL FUEL MANIFOLD - The described method of supplying fuel to an internal fuel manifold of a bypass gas turbine engine includes directing a fuel flow through a fuel fairing having an outer surface exposed to a cool bypass airflow, directing the fuel flow in the fuel fairing through a heat exchanging structure on the outer surface of the fuel fairing to cool the fuel flow being below a coking temperature of the fuel, and then feeding the cooled fuel flow from the fuel fairing to the internal fuel manifold. | 01-28-2016 |
20160025339 | ENERGY-EFFICIENT AND CONTROLLED VAPORIZATION OF CRYOFUELS FOR AIRCRAFT ENGINES - A method and apparatus of using cryogenic fuel in an engine for an aircraft wherein the cryogenic fuel is supplied to the engine for combustion. | 01-28-2016 |
20160040605 | GAS TURBINE ENGINE - A gas turbine engine including a compressor, a turbine having one or more stages and a combustor, the combustor being located between the compressor and turbine. The gas turbine engine further includes a bleed from a core defined by a core duct, the core duct surrounding and extending between the turbine and combustor at least. The bleed includes at least one inlet located downstream of the combustor and upstream of at least one of the turbine stages. The turbine is arranged in use to drive the compressor. The bleed is arranged to be controllable in use to selectively bleed air from the core through the inlet and to thereby control the power delivered by the turbine to the compressor. | 02-11-2016 |
20160040879 | COMBUSTOR PANEL WITH INCREASED DURABILITY - A liner panel for a combustor of a gas turbine engine includes a nominal wall thickness and a thickened wall thickness in the region of a hot spot. | 02-11-2016 |
20160040880 | COMBUSTOR HEAT SHIELD - There is provided a combustor comprising a dome and a shell extending from the dome defining a combustion chamber. A dome heat shield is mounted to the dome inside the combustion chamber. A front heat shield is mounted to the shell and spaced therefrom. The dome heat shield has a lip extending generally away from the dome heat shield and generally parallel to the shell and spaced inwardly of the front heat shield to define a gap between the lip and the front heat shield. The front heat shield has a leading edge opposite the lip. The combustor has impingement holes extending through the shell and disposed to direct impingement cooling jets to the upstream portion of the front heat shield. The leading edge, of the front heat shield has at least one scallop defining an opening and disposed to allow the impingement cooling jets to impinge directly on a portion of the peripheral lip adjacent the scallop. | 02-11-2016 |
20160054000 | COMBUSTOR FOR GAS TURBINE ENGINE - A combustor with a liner where each of the walls has a respective circumferential row of dilution holes defined therethrough adjacent a junction between the primary zone and the dilution zone. In the primary zone, the inner surface of each of the walls is covered by at least one heat shield attached thereto and spaced apart therefrom to allow air circulation between the inner surface and the at least one heat shield, the walls each having a plurality of cooling holes defined therethrough having a smaller diameter than that of the dilution holes. In the dilution zone, the inner surface of each of the walls is free of heat shields, and the walls each have a plurality of effusion cooling holes defined therethrough and having a smaller diameter than that of the dilution holes. | 02-25-2016 |
20160069266 | HEAT EXCHANGER ARRANGEMENT FOR TURBINE ENGINE - A turbine engine cooling arrangement according to an example of the present disclosure includes, among other things, a core passage for receiving a core flow for combustion, a first airflow source including a first passage adjacent the core passage for conveying a first airflow, a second airflow source including a second passage adjacent the first passage for conveying a second airflow, and a heat exchanger that is thermally connected with the first passage and the second passage for transferring heat between the first airflow and the second airflow. The first airflow and the second airflow stream over the heat exchanger in a parallel radial direction and a parallel axial direction. An engine inlet divides inlet air into the core flow, the first airflow, and the second airflow. A method of providing cooling air is also disclosed. | 03-10-2016 |
20160084165 | PRE-COOLER INLET DUCTS THAT UTILIZE ACTIVE FLOW-CONTROL AND SYSTEMS AND METHODS INCLUDING THE SAME - Pre-cooler inlet ducts that utilize active flow-control and systems and methods including the same are disclosed herein. The systems include a pre-cooler inlet duct for a jet engine that is configured to receive a pre-cooler air stream and to direct the pre-cooler air stream into a heat exchanger. The pre-cooler inlet duct includes a flow-directing surface, which defines at least a portion of the pre-cooler inlet duct, and an active flow-control device. The active flow-control device is located to modify a boundary layer fluid flow within a boundary layer adjacent the flow-directing surface to resist separation of the boundary layer from the flow-directing surface when the pre-cooler air stream flows through the pre-cooler inlet duct. The methods include methods of resisting boundary layer separation in the pre-cooler inlet duct by flowing the pre-cooler air stream across the flow-directing surface and modifying the boundary layer with the active flow-control device. | 03-24-2016 |
20160109128 | GAS TURBINE ENGINE WAVE GEOMETRY COMBUSTOR LINER PANEL - A panel for a combustor of a gas turbine engine includes a cold side defining at least one convex portion and at least one concave portion. The concave portion is in communication with a passage. A method of operating a combustor section of a gas turbine engine includes: directing an impingement flow toward a multiple of peaks on a cold side of a panel; directing the impingement flow from the multiple of peaks toward a multiple of troughs with a multiple of entrances on the cold side of the panel; and directing the impingement flow through the multiple of entrances and a respective multiple of effusion passages through the panel. | 04-21-2016 |
20160153326 | DEVICE FOR COOLING OIL FOR A TURBINE ENGINE | 06-02-2016 |
20160153359 | JET ENGINE COLD AIR COOLING SYSTEM | 06-02-2016 |
20160169109 | MODULATED COOLED P3 AIR FOR IMPELLER | 06-16-2016 |
20160169119 | METHOD TO CONTROL THE OPERATING TEMPERATURE OF A GAS TURBINE HEATER | 06-16-2016 |
20160169516 | GAS TURBINE ENGINE COMBUSTOR BULKHEAD ASSEMBLY | 06-16-2016 |
20160177830 | DIFFUSER CASE MIXING CHAMBER FOR A TURBINE ENGINE | 06-23-2016 |
20160178016 | CLUTCH ASSEMBLY WITH INTEGRATED SURFACE COOLING | 06-23-2016 |
20160178199 | COMBUSTOR DILUTION HOLE ACTIVE HEAT TRANSFER CONTROL APPARATUS AND SYSTEM | 06-23-2016 |
20160178202 | SYSTEM AND METHOD FOR UTILIZING COOLING AIR WITHIN A COMBUSTOR | 06-23-2016 |
20160186661 | COOLED COOLING AIR TAKEN DIRECTLY FROM COMBUSTOR DOME - A gas turbine engine includes a compressor, a turbine, and a combustor. The combustor includes a fuel injector and a vaporizer within the combustor positioned to receive liquid fuel from the fuel injector to vaporize the liquid fuel therein. The gas turbine engine includes an enclosed passage external to the combustor having a wall, a diffuser positioned to direct the air into the passage, causing the air to cool by transferring heat through the wall from the air within the passage to the vaporized fuel within the vaporizer, and a cooled cooling air passageway positioned to receive the air from the passage and direct the air after being cooled to at least one of the turbine and the compressor. | 06-30-2016 |
20160186998 | CONTOURED DILUTION PASSAGES FOR GAS TURBINE ENGINE COMBUSTOR - A wall assembly for use in a combustor of a gas turbine engine includes a support shell with a first inner periphery along an axis and a liner panel with a second inner periphery along the axis. The second inner periphery is smaller than the first inner periphery. Another wall assembly for use in a combustor of a gas turbine engine includes an annular grommet mounted between the support shell and the liner panel. The annular grommet defines a contoured inner wall. | 06-30-2016 |
20160195023 | METHOD TO OPERATE A COMBUSTOR OF A GAS TURBINE | 07-07-2016 |
20160201908 | VENA CONTRACTA SWIRLING DILUTION PASSAGES FOR GAS TURBINE ENGINE COMBUSTOR | 07-14-2016 |
20160201914 | SEALED COMBUSTOR LINER PANEL FOR A GAS TURBINE ENGINE | 07-14-2016 |
20160201917 | COOLED FUEL INJECTOR SYSTEM FOR A GAS TURBINE ENGINE | 07-14-2016 |
20170234223 | GAS TURBINE ELECTRICALLY DRIVEN SUPPLEMENTARY AIR SYSTEM FOR POWER AUGMENTATION AND EFFICIENCY IMPROVEMENTS | 08-17-2017 |
20170234227 | BACKUP SYSTEM FOR SUPPLYING COMPRESSED AIR TO A GAS TURBINE COMPONENT | 08-17-2017 |
20170234240 | INLET TURBINE AND TRANSMISSION FOR HIGH-MACH ENGINES | 08-17-2017 |
20180023476 | SHAFT ASSEMBLY OF A GAS TURBINE ENGINE AND METHOD OF CONTROLLING FLOW THEREIN | 01-25-2018 |