Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: An actuating device for opening and closing at least one shutter in a gas turbine engine, such as a turbojet engine, includes at least one actuator made of a two-way shape memory alloy having a first stable state at a first temperature in which state it actuates either the opening or the closing of the shutter, and a second stable state at a second temperature, in which state it actuates either the closing or the opening of the shutter, respectively.

Abstract: 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.

Abstract: A method to provide clearance control for a gas turbine having a multi-stage compressor and a turbine having turbine buckets rotating within a turbine shell, the method includes: selecting a first compressor stage from which to extract compressed air; ducting the compressed air from the first compressor stage to the turbine shell; passing the compressed air from the first compressor stage to thermally contract the turbine shell; selecting a second compressor stage from which to extract compressed air and deselecting the first compressor stage; ducting the compressed air from the second compressor stage to the turbine shell, and passing the compressed air from the second compressor stage to thermal expand the turbine shell.

Abstract: 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.

Abstract: A method of operating an integrated gasification combined cycle power generation system is provided. The method includes compressing air in an adiabatic air compressor to produce a compressed heated air stream, heating a nitrogen stream using the compressed heated air stream to produce a heated nitrogen stream and a cooled compressed air stream, and channeling the cooled compressed air stream to an air separation unit.

Abstract: 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.

Abstract: 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.

Abstract: A scoop for a fairing 1 of a core engine 2 of an aircraft gas turbine allows air to be supplied from a bypass flow in a bypass duct 3 to several cooling-air distributors in a core-engine ventilation compartment 4. The scoop includes a first tubular flow duct 5, whose inlet opening 6 is arranged in the bypass duct 3 and which extends through the fairing 1, as well as a second tubular flow duct 7, which at least partly encompasses the first flow duct 5 and whose inlet opening 8 is rearwardly offset relative to the inlet opening 6 of the first flow duct 5 in the direction of flow.

Abstract: A combustor (14) is placed next to a turbine (16), on the side opposite a compressor (12). A heat insulation device (20) for reducing the transmission of heat from the high-temperature side to the low-temperature side is provided between the combustor/turbine and the compressor. A connection shaft (18) has an axial hole (18a) open on the inlet side of the compressor and axially extending to near a turbine impeller, and also has a radial hole (18b) open near the turbine impeller to the outside of the connection shaft and radially extending to be in communication with the axial hole.

Abstract: 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.

Abstract: A precooler for cooling compressor bleed air for an environmental control system includes a heat exchanger in fluid communication with a source of cooling air and operable for cooling the bleed air. A variable bypass valve between a bleed air source and environmental control system is operable for bypassing at least a portion of the compressor bleed air around the heat exchanger. The cooling air may be a portion of fan air modulated by a variable fan air valve. The bleed air source may be selectable between the low pressure bleed air source and a high pressure bleed air source. One method includes flowing the compressor bleed air from a single low pressure source only and increasing thrust sufficiently to meet a minimum level of pressure of the bleed air during one engine out aircraft operating condition during approach or loitering.

Abstract: A method and apparatus for controlling a temperature a component of a gas turbine is disclosed. A compressed gas for use as a coolant is provided. The coolant is moisturized at a moisturizeing unit. A circulating unit circulates the moisturized coolant to the component of the gas turbine to control the temperature of the component. The coolant can be air, nitrogen, and a mixture of air and nitrogen in various embodiments. The component of the turbine can be a blade of a turbine section of the gas turbine, a turbine nozzle and a combustor, for example. A combustor can combust a mixture of fuel and the moisturized compressed coolant gas to reduce a NOx emission of the gas turbine.

Abstract: 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.

Abstract: A flow device adapted to operate as a restrictor as well as provide a pressure relief capability in the event of an over-pressurization event within a fluid system containing the device. The flow device includes an expandable orifice that has an outer perimeter, a plurality of cantilevered tabs surrounded by the outer perimeter, and an opening surrounded and defined by the tabs. The tabs project from the outer perimeter toward the opening, which restricts flow of the bleed air through the expandable orifice at a pressure below a predetermined pressure level, but then expands to relieve an over-pressure condition of the bleed air at a pressure above the predetermined pressure level as a result of the cantilevered tabs being deflected by the over-pressure condition. The device is adaptable for use in aircraft applications, including the regulation of bleed air used in anti-icing/de-icing systems.

Abstract: A combustion chamber head of a gas turbine has a substantially annular combustion chamber outer wall 18 as well as a substantially annular combustion chamber inner wall 42 and several burners 6 distributed around the circumference. The combustion chamber head 5 has an inflow-side wall 13 which together with a wall 14 facing the combustion chamber 7 forms a combustion chamber head volume 15. The inflow-side wall 13 is provided with at least one inflow opening 32, the wall 14 facing the combustion chamber 7 is provided with at least one outflow opening 17 for connecting the combustion chamber head volume 15 to the combustion chamber 7, and at least one cooling air duct 29 is provided in the wall 14 facing the combustion chamber 7. A method for cooling and damping of the combustion chamber head is also disclosed.

Abstract: 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.

Abstract: A method and system for increasing the efficiency of a turbomachine having an extraction system is provided. The extraction system may have multiple extraction ports from which the working fluid of the turbomachine is extracted to meet extraction load requirements of an independent process. The method and system may enable a control system to optimally determine the which extraction ports to draw the working fluid from for meeting the extraction load requirement. The optimal location may be determined by utilizing the data on the load requirement, where the extraction load requirement may be operationally distinct from the turbomachine load.

Abstract: On a gas-turbine engine, in particular an aircraft engine, with a fan, a fan casing and a fan duct as well as with high-pressure and low-pressure turbines arranged behind each other in flow direction in the casing of the engine, the auxiliaries connected are to be operated with only a small increase in engine power. For this purpose in the flow direction of the airflow (20) exiting from the high-pressure turbine (15) a bleeding mechanism is provided for bleeding radially into the fan duct (4) at least part of the airflow (20) leaving the high-pressure turbine (15) on the upstream side of the low-pressure turbine (16).

Abstract: A gas turbine, comprising a turbine and a compressor provided with a compressor housing, and to a method for the operation thereof. The compressor is tapped in order to cool the turbine by means of at least one tap line for removing compressed or partially compressed air. The tap line comprises a locking device, particularly a valve, in order to regulate the outflow of tapped air and thus the cooling of the housing.

Abstract: 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.

Abstract: 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.

Abstract: Methods, systems, and apparatus for controlling a turbine clearance in an aircraft engine are provided. A method includes activating a turbine clearance control based on a flight phase of an aircraft using the aircraft engine, and adjusting the turbine clearance based on a preselected turbine clearance value.

Abstract: A flow device adapted to operate as a restrictor as well as provide a pressure relief capability in the event of an over-pressurization event within a fluid system containing the device. The flow device includes an expandable orifice that has an outer perimeter, a plurality of cantilevered tabs surrounded by the outer perimeter, and an opening surrounded and defined by the tabs. The tabs project from the outer perimeter toward the opening, which restricts flow of the bleed air through the expandable orifice at a pressure below a predetermined pressure level, but then expands to relieve an over-pressure condition of the bleed air at a pressure above the predetermined pressure level as a result of the cantilevered tabs being deflected by the over-pressure condition. The device is adaptable for use in aircraft applications, including the regulation of bleed air used in anti-icing/de-icing systems.

Abstract: A system and method for reintroducing gas turbine combustion bypass flow. The system may include a combustor body, wherein the combustor body includes a reaction zone for primary combustion of fuel and air, and a casing enclosing the combustor body and defining an annular passageway for carrying compressor discharge air into the combustor body at one end. The system further may include a reintroduction manifold for receiving combustor bypass air extracted from the compressor discharge air in the annular passageway, and one or more reintroduction slots in communication with the reintroduction manifold for injecting the combustor bypass air into the combustor body downstream of the reaction zone. The method may include extracting combustor bypass air from the annular passageway, transporting the combustor bypass air to a reintroduction manifold, and reintroducing the combustor bypass air into the combustor body through one or more reintroduction slots in communication with the reintroduction manifold.

Abstract: 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.

Abstract: A resilient annular seal structure is disposed radially between an aft end portion of a combustor liner and a forward end portion of a transition piece, the resilient annular seal structure configured to form a first annular cavity radially between the forward end portion of the transition piece and the aft end portion of said combustor. At least one transfer tube radially extends from the second flow sleeve through the second flow annulus to the transition piece, and is arranged to supply compressor discharge cooling air radially from an area outside the first and second substantially axially extending flow annuli directly to the resilient annular seal structure and to the aft end of the combustor liner.

Abstract: A turboprop engine (1) includes an engine nacelle (3) and at least one bleed air line (25) on the low-pressure compressor (4) and at least one ejector (21) formed by a cooling air duct (24) and a nozzle (22) to create a cooling air flow within the engine nacelle during critical ground idle operation (controlled or uncontrolled), and without undesirably increasing fuel consumption or disturbing the work cycle of the engine (1). The ejector (21) is arranged within the engine nacelle (3) in the forward part of the turboprop engine (1), with the cooling air duct (24) appertaining to the ejector (21) connecting at least one air intake (23) disposed on the periphery of the engine nacelle (3) with the interior of the engine nacelle (3), and with the at least one nozzle (22) being arranged in the cooling air duct (24).

Abstract: 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.

Abstract: Means for cooling a bearing assembly supporting a rotor stage of a gas turbine engine after engine shutdown. The engine comprises a combustion section, a compressor for the delivery of air to the combustion section, and a bearing assembly supporting a rotor stage. The means for cooling a bearing assembly comprises a means operable to generate a signal representative of turbine duct temperature immediately prior to engine shutdown, a means for determining the duration for cooling after engine shutdown as a function of the signal representative of turbine duct temperature, and the compressor operable as a cooling means to deliver cooling air to the rotor stage after engine shutdown. Thereby the amount of heat conducted to the bearing assembly is limited such that the temperature of the bearing assembly is limited to below a predetermined temperature.

Abstract: 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.

Abstract: A heat exchanger system for a gas turbine engine includes: (a) a fan having at least two stages of rotating fan blades surrounded by a fan casing, the fan operable to produce a flow of pressurized air at a fan exit; (b) at least one heat exchanger having a first flowpath in fluid communication with the fan at a location upstream of the fan exit; and (c) a fluid system coupled to a second flowpath of the at least one heat exchanger. The first and second flowpaths are thermally coupled to each other.

Abstract: Embodiments of methods and apparatus for providing compressor extraction cooling are provided. According to one example embodiment, a method for providing compressor extraction cooling in a gas turbine comprising a compressor and a turbine section can be provided. 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.

Abstract: 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.

Abstract: An embodiment of the present invention takes the form of an application and process that incorporates a waste heat source to increase the temperature of the airstream entering an inlet section of a combustion turbine. An embodiment of the present invention may perform an anti-icing operation that reduces the need to operate the IBH system.

Abstract: A steam turbine is provided. The steam turbine includes a rotor shaft including a plurality of buckets coupled thereto. The steam turbine further includes a stationary component coupled to a steam turbine casing, wherein the stationary component is coupled upstream from the buckets such that a wheelspace is defined between the buckets and the stationary component. The stationary component includes a first ring coupled to the steam turbine, a second ring coupled to the steam turbine radially inward from the first ring, and at least one airfoil extending between the first ring and the second ring. The steam turbine includes a cooling fluid flowpath defined through at least the first ring, the airfoil, and the second ring. The cooling fluid flowpath is configured to channel a cooling fluid to the wheelspace.

Abstract: A method for adjusting a clearance between a blade tip of a turbine engine and a blade track spaced radially outward of the blade tip is disclosed herein. The method includes the step of operably coupling an elongate member to a blade track in a turbine engine. The method also includes the step of directing a fluid stream having a temperature in proximity to the elongate member. The temperature of the fluid stream can change over time. The method also includes the step of transferring heat between the fluid stream and elongate member to a change a size of the elongate member and move the blade track radially relative to a centerline axis of the turbine engine. An exemplary apparatus for carrying out the method is also disclosed.

Abstract: 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.

Abstract: A vortex reducer for the guidance of bleed airflows (21), is arranged in an inter-disk chamber (3) between two rotor disks (1, 2) of the compressor of a gas turbine with at least one shaft, and includes at least one ring (11) with circumferentially disposed hole passages, in which bleed air tubes (15) are arranged. In order to provide a vortex reducer, which requires small material input and, therefore, has low weight, while producing directed bleed airflows with low pressure losses, the hole passages include first hole passages (13) and second hole passages (14), and bleed air tubes (15) are provided only in the first hole passages (13), with the bleed air tubes (15) being evenly distributed on the circumference of the ring (11), and the second hole passages (14) being devoid of bleed air tubes (15).

Abstract: 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.

Abstract: 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.

Abstract: A gas-turbine compressor has a casing (1) in which a rotor hub (2) is rotatably borne, and a compressor duct (9) being disposed between the casing (1) and the rotor hub (2), in which at least one rotor (4), which is rotatable about a machine axis (5), and one stator (5) are arranged. Recesses (12) of a bleed-air tapping device (6) are provided in the casing (1), which—in at least one circumferential area (11)—are arranged circumferentially to each other recess (12) and include a circumferential leading edge (16) in the circumferential direction and a circumferential trailing edge (17), each of which includes an identical angle ?E with the surface (18) of the casing (1).

Abstract: A suspension of a multi-flow turbojet engine provided with an intermediate casing and an exhaust casing from a pylon that can be attached to the structure of an aircraft is disclosed. The suspension includes a forward attachment device between the hub of the intermediate casing and the pylon, a rear attachment device between the exhaust casing and the pylon, and a connection device rigidly connecting the intermediate casing to the pylon. The rear attachment device includes an actuator for compensating for the variations in diameter of the exhaust casing so as to keep the axis of the exhaust casing coaxial with the axis of the intermediate casing through the various phases of flight of the aircraft.

Abstract: A turbine engine including a controller controlling clearance between tips of moving blades of a high-pressure turbine and an outer casing surrounding the blades, by cooling the outer casing by the impact of air taken from a high-pressure compressor stage of the engine, and by electric heating of top and bottom portions of the outer casing.

Abstract: 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.

Abstract: 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.