Abstract: A system and method for flame stabilization is provided that forestalls incipient lean blow out by improving flame stabilization. A combustor profile is selected that maintains desired levels of power output while minimizing or eliminating overboard air bleed and minimizing emissions. The selected combustor profile maintains average shaft power in a range of from approximately 50% up to full power while eliminating overboard air bleed in maintaining such power settings. Embodiments allow for a combustor to operate with acceptable emissions at lower flame temperature. Because the combustor can operate at lower bulk flame temperatures during part power operation, the usage of inefficient overboard bleed can be reduced or even eliminated.

Abstract: 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.

Abstract: A system for directing air flow to separate plena of a compartment that is defined at least by a compartment wall includes a NACA scoop, and a Pitot scoop. The NACA scoop is formed in the compartment wall, and includes two side walls, a bottom wall, and an entrance lip that is defined by the compartment wall and is spaced apart from the bottom wall to form a NACA scoop air inlet. The Pitot scoop is longitudinally aligned with the NACA scoop, includes a Pitot scoop air inlet, a Pitot scoop air outlet, and a Pitot scoop flow passage.

Abstract: An electronic controller for a sub-system of an air management system including an area conversion module and a sub-system area control module. The area conversion module includes inputs for a bleed output pressure set point, a bleed output pressure measurement, a sub-system parameter set point, a sub-system parameter measurement, and a sub-system output pressure measurement; and sub-modules for calculating an area set point and an area measurement. The sub-modules generate the area set point as a function of the engine bleed output pressure set point, the sub-system parameter set point, and the sub-system output pressure measurement; and the area measurement as a function of the bleed output pressure measurement, the sub-system parameter measurement, and the sub-system output pressure measurement. The sub-system area control module generates a control output for the sub-system as a function of the calculated area set point and the calculated area measurement.

Abstract: A system may include a compressor, a heat exchanger and an ITM. The compressor is configured to receive an air stream and compress the air stream to generate a pressurized stream. The heat exchanger is configured to receive the pressured stream and indirectly heat the pressurized stream using heat from an oxygen stream from an Ion Transport Membrane (ITM). The ITM is configured to receive the heated pressurized stream and generate an oxygen stream and the non-permeate stream, wherein the non-permeate stream is passed to a gas turbine burner and the oxygen stream is passed to the heat exchanger.

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: A method for optimizing operability of an aircraft propulsive unit, and a self-contained power unit implementing the method. The method removes mechanical bleed constraints in engines during transient flight phases of an aircraft to optimize operability of the engine assembly during the phases. To this end, a supply of power is provided, particularly during the phases, by an additional indirectly propulsive engine power source. The method for optimizing operability of the propulsive unit of an aircraft including main engines as main drive sources includes, using a main engine power unit GPP as a power source, providing all the non-propulsive power and, during the transient engine phases, at most partially providing additional power to the body of the main engines.

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: A device (30) for optimizing the cooling in a gas turbine of the type comprising at least one compressor, equipped with a combustion chamber (10) and an outer casing (16) and inner casing (32), at least one turbine wheel (14), equipped with a series of blades 12), and at least one high-pressure rotor (38), equipped with one or more supporting bearings (34), the compressor being capable of generating cooling air sent to the turbine wheel (14) through a suitable channel (20). On the outer surface of the device (30), there are one or more grooves (26) which allow the passage of additional flow-rates of air from the compressor towards the channel (20), in order to increase the overall cooling air flow towards said turbine wheel (14). The device (30) can be easily installed without the necessity of dismantling the outer casing (16), and is assembled on pre-existing machines in substitution of the vent tube (22) of the seals (36) of the supporting bearings (34).

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: 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: An air bleed device for cooling components in a turbine engine, including an annular conduit having a substantially rectangular cross-section formed in a housing and having a radially internal wall swept by an airflow is disclosed. The device includes an air inlet orifice, and a flap valve for controlling the airflow entering through the orifice, formed by a plate borne by a maneuvering member mobile in translation parallel to the axis of the orifice between a position in which the plate closes off the orifice and a position in which the plate opens the orifice.

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 system for operating a power plant is provided and includes a grid configured to generate a normal load and an abnormal load, a turbomachine configured to provide power to the grid in accordance with the normal load by firing at normal temperatures and in accordance with the abnormal load by firing at higher-than-normal temperatures, a cooling system disposed to cool components of the turbomachine with fluid supplied by an external reservoir and a controller configured to identify when the grid generates the abnormal load and to responsively operate the cooling system.

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: Compact, high-pressure exhaust muffling devices are disclosed. Some example muffling devices may include an inner flow conditioner including an inlet configured to convey a pressurized fluid flow into an inner flow conditioner interior. The inner flow conditioner may include a plurality of inner flow conditioner holes. The muffling device may include an exhaust can disposed substantially around the inner flow conditioner and arranged to receive the pressurized fluid flow via the inner flow conditioner holes into an exhaust can interior. The exhaust can may include a plurality of exhaust screen holes arranged to discharge the pressurized fluid flow from the exhaust can interior. The exhaust can interior may be substantially devoid of flow obstructions between the plurality of inner flow conditioner holes and the exhaust screen holes.

Abstract: Embodiments of a gas turbine engine cooling system for deployment within a gas turbine engine are provided, as are embodiment of a method for producing a gas turbine engine cooling system. In one embodiment, the gas turbine engine cooling system includes an impeller having a hub, a plurality of hub bleed air passages, and a central bleed air conduit. The plurality of hub bleed air passages each have an inlet formed in an outer circumferential surface of the hub and an outlet formed in an inner circumferential surface of the hub. The central bleed air conduit is fluidly coupled to the outlets of the plurality of hub bleed air passages and is configured to conduct bleed air discharged by the plurality of hub bleed air passages to a section of the gas turbine engine downstream of the impeller to provide cooling air thereto.

Abstract: An auxiliary power unit is operable to provide bleed air to a vehicle system. A method includes diverting substantially all of the bleed air to an exhaust for a selected time period commencing with startup of the auxiliary power unit. After the selected time period, at least a portion of the bleed air is diverted to the vehicle system.

Abstract: In a cooling structure for a recovery-type air-cooled gas turbine combustor having a recovery-type air-cooling structure that bleeds, upstream of the combustor, and pressurizes compressed air supplied from a compressor, that uses the bled and pressurized air to cool a wall, and that recovers and reuses the bled and pressurized air as combustion air for burning fuel in the combustor together with a main flow of the compressed air, wall cooling in which cooling air is supplied to cooling air passages formed in the wall of the combustor to perform cooling involves a downstream wall region, closer to a turbine, that is cooled using the bled and pressurized air as the cooling air and an upstream wall region, closer to a burner, that is cooled using, as the cooling air, bled compressed air bled from a main flow of the compressed air through a housing inner space.

Abstract: A compressor bleed cooling fluid feed system for a turbine engine for directing cooling fluids from a compressor to a turbine airfoil cooling system to supply cooling fluids to one or more airfoils of a rotor assembly is disclosed. The compressor bleed cooling fluid feed system may enable cooling fluids to be exhausted from a compressor exhaust plenum through a downstream compressor bleed collection chamber and into the turbine airfoil cooling system. As such, the suction created in the compressor exhaust plenum mitigates boundary layer growth along the inner surface while providing flow of cooling fluids to the turbine airfoils.

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: An exemplary lubrication system includes a flow path that carries fluid from a fluid supply through at least a portion of an auxiliary power unit to an outlet separate from the fluid supply. The fluid supply is pressurized to move the fluid along the flow path.

Abstract: A bypass turbojet engine including an outer fan duct and an inner fan duct (IFD), between which the secondary flow passes, and a high pressure compressor having an outer casing is disclosed. The IFD is fastened on the outer casing.

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 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 fluidically-controlled valve is provided. The fluidically-controlled valve includes a main flow channel with a main flow entrance, a flow exit and a constricted channel section located between the main flow entrance and the flow exit. The fluidically-controlled valve also includes a control flow channel including a jet forming control entrance, a first branch channel, a second branch channel, a common channel section, and a convex channel wall. The common channel section follows the control entrance, the first branch channel emerges from the common channel section and leads to the main flow entrance, the second branch channel emerges from the common channel section and leads to the constricted channel section, and the convex channel wall extends from the common channel section into the first branch channel. The fluidically-controlled valve can be used in bypasses present in turbines or in swirlers of gas turbine combustors.

Abstract: Air bleed device for cooling components in a turbine engine, including an annular conduit (18) having a radially internal portion swept by an airflow (22) and which comprises at least one air inlet orifice (20) formed in an upstream radial wall (38) of the conduit (18), a flap valve (40) for controlling the airflow entering through the orifice (20) and of which the flapper (41) includes a plate (42) applied against the wall (38) and capable of being moved by sliding on said wall (38) by a maneuvering member (50) mobile in translation parallel to the wall (38) between a position in which the plate (42) closes off the orifice (20) and a position in which the plate (42) opens said orifice.

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 bleed air duct joint insulation means is disclosed in which a leak vent comprises valve means.

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: Device (10) for the tapping of air for the cooling of flaps of a turbojet nozzle, comprising an annular duct (18) having a radially internal wall (22) swept by a stream of air (24) and which comprises at least one air inlet orifice (20), a flap valve (40) for controlling the flow rate of air entering through the orifice (20), formed of an elastically deformable metal plate (42) of which a downstream end is fixed on an edge of the orifice (20), and of which an upstream end can be displaced by a manoeuvring member (46) mobile in translation parallel to the axis (52) of the orifice between a position where the plate (42) seals this orifice and a position where the plate (42) opens this orifice.

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 mid-turbine frame buffer system for a gas turbine engine includes a mid-turbine frame that supports a shaft by a bearing. An air compartment and a bearing compartment are arranged radially inward of the mid-turbine frame. The bearing compartment is arranged within the air compartment and includes first and second contact seals arranged on either side of the bearing. The air compartment includes multiple air seals. A high pressure compressor is fluidly connected to the air compartment and is configured to provide high pressure air to the air compartment. A method of providing pressurized air to a buffer system includes sealing a bearing compartment with contact seals, surrounding the bearing compartment with an air compartment, and supplying high pressure air to the air compartment.

Abstract: A compressor bleed air valve in an auxiliary power unit (“APU”) modulates bleed air flow both for accessory pneumatic systems powered by the bleed air and for reducing or removing surge in the compressor.

Abstract: At least one main air compressor makes a compressed ambient gas flow. The compressed ambient gas flow is delivered to both master and slave turbine combustors at a pressure that is greater than or substantially equal to an output pressure delivered to each turbine combustor from each turbine compressor as at least a first portion of a recirculated gas flow. A fuel stream is delivered to each turbine combustor, and combustible mixtures are formed and burned, forming the recirculated gas flows. A master and slave turbine power are produced, and each is substantially equal to at least a power required to rotate each turbine compressor. At least a portion of the recirculated gas flow is recirculated through recirculation loops. At least a second portion of the recirculated gas flow bypasses the combustors or an excess portion of each recirculated gas flow is vented or both.

Abstract: A device is provided for pivoting at least one pivotable element in a gas turbine engine between a first and a second position in order to influence a gas flow in an annular gas duct in at least one of the positions. The device includes a moveable annular member which is arranged externally around the gas duct and is connected to the pivotable element in order to effect the pivoting of the pivotable element. The annular member is more specifically arranged to be displaced in a substantially axial direction and arranged to pivot the pivotable element when it is displaced axially.

Abstract: A gas turbine inlet heating system is disclosed. In one embodiment, the system includes: a compressor having: an inlet bellmouth adjacent to a set of inlet guide vanes (IGVs); and an outlet fluidly connected to the inlet bellmouth; and a conduit coupled to an outlet of the compressor, the conduit including a control valve, the conduit for diverting a first portion of compressed air from the outlet of the compressor to the inlet bellmouth.

Abstract: A wing-engine combination includes: a wing with a main wing and an engine with a premixing chamber, a combustion chamber and a hot-air space, which includes: an engine bleed-air duct, which extends along the wingspan direction and along the leading edge of the main wing, including an engine bleed-air inlet device that is coupled to an engine hot-air space, and including an engine bleed-air outlet device having a discharge orifice on the main wing or a connecting part for coupling the engine bleed-air duct to a consumer of the engine bleed-air, an ambient-air duct, which extends along the engine bleed-air duct, including an ambient-air inlet device which is arranged on an aircraft component of the aircraft, which aircraft component faces the intended flow-around direction of the aircraft and includes an aperture for letting ambient air into the ambient-air duct, and including an ambient-air outlet device with a passage between the ambient-air duct and a premixing chamber of the engine so that the arrangement i

Abstract: A noise reduction device, for example for use in a bleed assembly of a gas turbine engine, comprises partitions having apertures which provide contractions and sudden expansions of flow passing through the device, turbulators being provided between the partitions to break up jets issuing from the Apertures. Breaking up the jets before impact with the partitions enables the partitions to be disposed closer together, so enabling adequate noise reduction to be achieved in a device of relatively small thickness.

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 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: A bleed flow discharge device (136) adapted to discharge a bleed fluid flow into a main fluid flow, wherein the bleed flow discharge device comprises an outer wall (135) defining a passage (137) for the bleed fluid flow, the outer wall comprising a wave-shaped edge (139) where the bleed fluid flow meets the main fluid flow.

Abstract: A method and system for an integrated ejector valve assembly is provided. The integrated ejector valve assembly includes a first valve assembly configured to control a flow of relatively lower pressure fluid from a first inlet port, a second valve assembly configured to control a flow of relatively higher pressure fluid from a second inlet port, a first actuation chamber configured to close the first valve assembly, a second actuation chamber configured to close the second valve assembly, and a third actuation chamber configured to open the second valve assembly.

Abstract: A system for managing fuel temperature in an engine includes a source of hot pressurized air and a turbine for converting hot pressurized air into cool expanded air. The system further includes a fuel tank for storing fuel, a fuel conduit fluidly connected to the fuel tank, and a first heat exchanger located on the fuel conduit. The first heat exchanger places the cool expanded air from the turbine in a heat exchange relationship with the fuel, thereby cooling the fuel.

Abstract: An aircraft compound cooling system includes a power thermal management system for cooling one or more aircraft components, an air cycle system, a vapor cycle system, and a turbine cooling circuit for cooling bleed air and cooling turbine components in a high pressure turbine in the engine. An air to air FLADE duct heat exchanger is disposed in a FLADE duct of the engine and a valving apparatus is operable for selectively switching the FLADE duct heat exchanger between the turbine cooling circuit and the air cycle system. A vapor cycle system includes a vapor cycle system condenser that may be in heat transfer cooling relationship with the air cycle system. An air cycle system heat exchanger and an engine burn fuel to air heat exchanger in the vapor cycle system condenser may be used for cooling a working fluid in a refrigeration loop of the vapor cycle system.