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: An inlet bleed heat and power augmentation system utilizing a bi-directional and common piping arrangement is disclosed. The piping arrangement includes a plurality of feed tubes arranged to communicate either steam to a compressor discharge plenum or compressed air from the compressor discharge plenum. Various embodiments of the invention are discussed including operation methods.

Abstract: Disparities between optimum engine thrust for aircraft operation and power necessary to drive an ancillary machine such an electrical power generator can cause problems within a turbine engine. Clearly, it is necessary to continue developing sufficient electrical power for control and other systems within an aircraft despite aircraft thrust requirements. Previously, some compressed air flow has been bypassed in order to maintain compressor stability, but unfortunately such bypass systems within turbine engines can have a thrust recovery regime such that there is excess thrust in comparison with that actually necessary.

Abstract: 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.

Abstract: A method of controlling variable extraction flow in a combustion turbine engine, wherein extraction flow comprises a supply of compressed air extracted from the compressor and supplied to the turbine through extraction conduits, and wherein the extraction conduits includes a variable extraction orifice, the method comprising the steps of: measuring a plurality of turbine engine operating parameters; monitoring, by a control unit, the measured operating parameters of the combustion turbine engine; setting the variable extraction orifices to a setting that allows an approximate maximum level of extraction flow; calculating, by the control unit, at least one calculated operating parameter based upon model-based control and the measured operating parameters, including at least a current turbine inlet temperature and a maximum turbine inlet temperature; and manipulating the setting for the supply of fuel to the combustor such that an increased and/or maximum level of engine output is determined by comparing the va

Abstract: A gas turbine engine includes a compressor, combustor, and turbine having a row of blades mounted inside a surrounding turbine shroud. A heat exchanger is used for cooling pressurized air bled from the compressor. A distribution network joins the heat exchanger to the turbine for selectively channeling air from the heat exchanger below the blades and above the shroud for controlling blade tip clearance.

Abstract: A system is provided for directing air from plural compressor ports to provide cooling and/or sealing air to an associated turbine site. A first flow from a pressure stage of the compressor has a first pressure and temperature. A second flow from another pressure stage of the compressor has a second pressure and temperature. The first and second pressures/temperatures are different. An ejector has two inlets for receiving the first and second flows, and output for combining the first and second flows into a third flow. The pressure and temperature of the third flow are different from the first and second pressures and temperatures. A bypass line is connected between the first flow and the third flow, and provides a bypass flow. A mixer combines the bypass flow and the third flow into a fourth flow. The fourth flow has a pressure and temperature intermediate the pressure and temperature of the bypass flow and the third flow. The mixer comprises inner and outer sections.

Abstract: A technique is provided for operating a gas turbine engine that has a combustor with a primary stage and one or more other stages and a first compressor providing an air flow to the combustor. This technique includes driving a variable load device with the rotating shaft of the gas turbine engine and sensing pressure of the air flow and an engine speed. In response to a decrease in loading of the engine by the variable load device: selectively bleeding the air flow as a function of the engine speed and regulating temperature in the primary stage of the combustor as a function of a ratio between fuel flow provided to primary stage and the pressure to prevent engine flame out. In one form, the combustor is arranged as a dry load emissions type and the variable load device includes an electric power generator.

Abstract: 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.

Abstract: In certain circumstances recovery of a fluid flow presented through a heat exchanger into another flow can create problems with respect to drag and loss of thrust. In gas turbine engines heat exchangers are utilized for providing cooling of other flows such as in relation to compressor air taken from the core of the engine and utilized for cabin ventilation and de-icing functions. By providing an outlet valve through an outlet duct in a wall of a housing the exhaust fluid flow from the heat exchanger can be returned to the by-pass flow with reduced drag effects whilst recovering thrust. The valve may take the form of a flap displaceable into the by-pass flow before an exit plan to create a reduction in static pressure drawing fluid flow through the heat exchanger.

Abstract: 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.

Abstract: 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.

Abstract: An air-bleed gas turbine includes a compressor; a combustor; a turbine; and an air bleed portion by which compressed air, which is supplied from the compressor to a circular space that surrounds the combustor, is partially supplied to a device outside the air-bleed gas turbine. The air bleed portion includes an air bleed passage that is communicated with an air bleed port which is connected to the device outside the air-bleed gas turbine, and a communication passage that is communicated with the circular space. The air bleed passage is communicated with the circular space through the communication passage. The flow passage area of the communication passage is increased as the distance from the air bleed port increases in the circumferential direction of the circular space.

Abstract: The bypass turbomachine comprises, between a primary flowpath (2) and a secondary flowpath (3), a structural intermediate casing (14). A bleed passage (29) allows the diversion of a portion of the gas stream toward the secondary flowpath (3), downstream of support arms (18) situated at the periphery of the intermediate casing. A bleed opening (28) emerges upstream of the support arms (18). A slide valve (22) slides between three positions, namely a closed position in which the slide valve (22) completely blocks off the bleed passage and opening, an intermediate position in which the bleed opening (28) emerging upstream of the support arms (18) is blocked off, while the bleed passage (29) for diverting a portion of the gas stream downstream of the support arms (18) is open, and an open position in which the bleed opening (28) emerging upstream of the support arms (18) and the bleed passage (29) emerging downstream of the support arms are both open.

Abstract: A gas turbine engine control component includes at least one electronic device, electronics such as an integrated circuit associated with the device, and a thermoelectric cooler for cooling the electronics mounted in a compartment. The thermoelectric cooler may be disposed in or on a wall of the compartment with a heat sink connected to a hot side of the thermoelectric cooler and a cold side of the thermoelectric cooler exposed to an interior of the compartment. Data about and/or operating instructions for the device may be stored in memory on the integrated circuit. The data may be calibration information for the device. A bus connector is connected to the integrated circuit for transferring operating instructions and/or data from the integrated circuit out and/or out of the component. A controller or control system incorporating these devices and components have the devices electronically connected to the integrated circuit.

Abstract: 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.

Abstract: A gas turbine engine with a turbine with an exhaust manifold thereabout from which fluids can be transferred to the turbine and an air compressor having an air transfer duct extending therefrom so as to be capable to provide compressed air in that air transfer duct at one end thereof to the air intake of an internal combustion engine provided as an intermittent combustion engine. The air intake and an exhaust outlet are each coupled to combustion chambers therein and a rotatable output shaft is also coupled to those combustion chambers for generating force. The exhaust outlet has an exhaust transfer duct extending therefrom so as to have the intermittent combustion engine be capable to provide exhaust therefrom in that exhaust transfer duct at one end thereof, and the exhaust transfer duct being connected at an opposite end thereof to the exhaust manifold to be capable of transferring intermittent combustion engine exhaust thereto.

Abstract: A system to mitigate thermal lag of start-up operations of a gas turbine is provided and includes a housing which is disposed upstream from and which is fluidly coupled to a compressor of the gas turbine, a first system coupled to an aft portion of the compressor and to the housing which, when activated, receives a first air supply from the compressor and delivers the first air supply to the housing, and a second system, coupled to the first system, which removes a second air supply from an intermediate portion of the compressor and delivers the second air supply to the first system prior to the activation of the first system.

Abstract: 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.

Abstract: A shroud suitable for use in a gas turbine engine exhibits substantially uniform thermal growth.

Abstract: A method for recovering energy from bleed air from a compressor of a gas turbine is provided. The method includes directing bleed air from a compressor of a gas turbine into a cooling unit. The bleed air is compressed air from the compressor, and the bleed air is not directed into a combustion chamber of the gas turbine. The method also includes cooling the bleed air within the cooling unit. The method further includes expanding the cooled bleed air within a turbo-expander to generate power.

Abstract: 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.

Abstract: A gas turbine engine includes a compressor, combustor, and high pressure turbine operatively joined together. The turbine includes a row of nozzle vanes followed by a row of rotor blades. The vanes and blades have corresponding forward and aft internal cooling channels. First, second, third, and fourth bleed circuits are joined in flow communication with different stages of the compressor for bleeding pressurized air therefrom at different pressures to provide coolant to the forward and aft channels of the turbine vanes and blades.

Abstract: A cooling system of a gas turbine engine, includes a heat exchanger having a common wall shared by a first air passage for directing a portion of a compressor air flow to be used as cooling air, and a second air passage for directing a portion of a bypass air flow, the portion of the compressor air flow being thereby cooled by the portion of the bypass air flow through the common wall

Abstract: Acoustic noise from relatively high pressure gas flows as a result of opening for example a bypass valve in a gas turbine engine can cause problems. By provision of an acoustic apparatus incorporating an inner chamber where there is a pressure reduction and then a chamber divider surface incorporating small apertures with a wide spacing distribution an upward shift in acoustic frequency of the gas flow is achieved beyond normal human comprehension. An outer chamber is provided to receive the fine gas flow jets from the apertures in the chamber divider surface for further pressure reduction and low pressure release through wider apertures in an exit surface.

Abstract: 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.

Abstract: An electrical machine is embedded into the compressor assembly of a gas turbine engine. An electrical system interface module distributes electrical current to and from the embedded electrical machine for starting the gas turbine engine and for operating accessory components. Accordingly, the gas turbine engine and accessory components can be started and operated without a power-takeoff shaftline and without an external accessory gearbox.

Abstract: A flow control system comprises includes a fluidic control device having a main fluid path for a flow of fluid through the device, and a control fluid path for a control flow of fluid through the device. At least part of the control fluid path coincides with at least part of the main fluid path to control the flow of fluid out of the fluidic control device. A valve is associated with the control fluid path. The valve is movable between an open condition to allow fluid to flow along the control fluid path to effect the aforesaid control of the fluid flow out of the fluidic control device, and a closed condition to inhibit or prevent fluid flow along the control fluid path.

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: 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 system and method for actively managing blade tip clearances in a turbine engine, particularly under steady state operating conditions such as at base load, involves routing a portion of air from a rotor cooling air circuit to a vane carrier or other stationary support structure surrounding the turbine blades. Because the temperature of the air is less than the temperature of the stationary support structure, the stationary support structure will thermally contract when the air is passed in heat exchanging relation therewith. In one embodiment, the air can be passed through one or more passages extending through at least a portion of the stationary support structure. The contraction of the stationary support structure reduces the blade tip clearance because the blades do not contract. Thus, fluid leakage through the clearances is minimized, which in turn can increase engine performance.

Abstract: A system for mixing flows from a compressor is provided. The system comprises an ejector that combines a first flow and a second flow, from the compressor, into a third flow. A bypass flow is connected between the first flow and the third flow. A mixer combines the bypass flow and the third flow into a fourth flow. The fourth flow has a pressure and temperature intermediate the respective pressures and temperatures of the bypass flow and the third flow.

Abstract: An engine combination for generating forces with a gas turbine engine generating force, and an internal combustion engine provided in the combination as an intermittent combustion engine generating force having an air intake, there being an air transfer duct connected from a compressor in the gas turbine engine to the air intake to transfer compressed air thereto.

Abstract: A power generation system for propelling, and generating electrical power in, an aircraft, having a gas turbine engine in an engine compartment in the aircraft with an air inlet in the aircraft that is curved along its extent in leading to an air compressor in the gas turbine engine having a compressor air transfer duct extending therefrom to an internal combustion engine provided as an intermittent combustion engine at an air intake thereof. An inlet duct manifold is positioned against the duct wall of the inlet duct to cover a perforated portion thereof on the air compressor side of a curve therein with the inlet duct manifold having an inlet air transfer duct extending therefrom that is coupled to the intermittent combustion engine air intake.

Abstract: A system for providing air to a compressor of an auxiliary gas turbine engine of an aircraft. A system turbine has an inlet adapted to receive compressed air from the aircraft. A system compressor is mechanically coupled to the system turbine and has an inlet adapted to receive atmospheric air. The outlets of the system turbine and the system compressor are fluidly connectable to the inlet of the compressor of the auxiliary gas turbine engine. A method includes providing compressed air from an aircraft to an inlet of a system turbine mechanically coupled to a system compressor. The method includes providing atmospheric air to an inlet of the system compressor. The method includes providing the inlet of the compressor of an auxiliary gas turbine engine with air from an outlet of the system turbine and with air from an outlet of the system compressor.

Abstract: An air cooled core mounted ignition system for gas turbine engine applications is provided. The ignition system includes an ignition exciter component directly mechanically and electrically connected to an igniter component. The housing member of the exciter component includes an air plenum configured to receive bleed air from the engine fan or compressor sections of the turbine engine, or other source. The bleed air provides a relatively low temperature air source for the purpose of cooling of the exciter. As such, the exciter component can be directly secured to the igniter, thereby eliminating the need for an ignition lead.

Abstract: A gas turbine engine includes an engine core extending along a core axis, a cooling-air delivery duct on the engine core, and a removable nacelle cowl overlying the engine core. A cooling-air nacelle-cowl duct delivers cooling air to the cooling-air delivery duct. At least a portion of the length of the cooling-air nacelle-cowl duct is integral with the nacelle cowl and not directly supported on the engine core.

Abstract: A turbine nozzle (2) that, among components constructing a turbine, reaches particularly high temperature is efficiently cooled with a relatively simple structure. A double casing structure in which a turbine casing (7) is provided outside a turbine shell (5) is formed. The turbine casing (7) functions as a flow path (24) for compressed air (20?21) before combustion. The turbine shell (5) covers a turbine nozzle (2) and a radial turbine impeller (3) and forms flow paths (15, 16) for combustion gas (10?11?12?13). The compressed air (21) before combustion flowing in the flow path (24), for compressed air, having air-tightness between itself and the outside air is blown to the turbine nozzle (2) through a through-hole (51) penetrating both wall surfaces of the turbine shell (5) so that the turbine nozzle (2) is evenly cooled and the compressed air used to cool the turbine nozzle is made to flow toward the turbine impeller (3).

Abstract: 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.

Abstract: 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.

Abstract: A Turbofan engine includes a fan mounted to a fan frame inside a fan nacelle. A booster compressor is joined to the fan inboard a flow splitter. A booster bleed system is disposed inside the splitter, and includes an inlet at the compressor outlet, and an outlet joined to the bypass duct following the fan.

Abstract: A method of controlling a load of a gas turbine engine that is part of an integrated gasification combined-cycle power generation system, which includes an air separation unit, that includes the steps of: (1) extracting an amount of compressed air that is compressed by a turbine compressor; (2) supplying the extracted amount of compressed air to the air separation unit; and (3) varying the amount of compressed air extracted from the turbine compressor based upon a desired load for the gas turbine engine.

Abstract: 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.

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: A combustor for a gas turbine includes a first combustor component and a second combustor component. The second combustor component is at least partially insertable into the first combustor component, and the first combustor component and second combustor component define a combustion fluid pathway. A combustor seal is located between the first combustor component and the second combustor component. The combustor seal defines at least one inner cooling pathway between the combustor seal and the second combustor component and at least one outer cooling pathway between the combustor seal and the first combustor component for cooling the first combustor component and second combustor component. A method for cooling a first combustor component and a second combustor component is also disclosed.

Abstract: A turbine engine has a circumferentially segmented shroud within a case structure. Each shroud segment is mounted for movement between an inboard position and an outboard position. One or more springs bias the shroud segments toward their inboard positions. One or more valves are positioned to vent one or more volumes so as to counter the spring bias to shift the shroud segments to their outboard positions.

Abstract: An improved inlet bleed heat system for a gas turbine engine is disclosed. The inlet bleed heat system provides improved mixing in an inlet region while permitting the engine to be operated at lower power settings. The inlet bleed heat system comprises a supply conduit, a plurality of feed tubes extending from the supply conduit, and a guide tube for receiving opposing ends of the feed tubes. The plurality of feed tubes each have a plurality of injection orifices and the feed tubes are oriented such that the injection orifices generally face into a flow of oncoming air with the feed tubes being positioned forward of a plurality of sound attenuating baffles.

Abstract: A gas turbine engine system including a plurality of bleed off doors forming a portion of a conical surface in the flow path between the low pressure compressor and the intermediate pressure compressor. A plurality of actuators are coupled to a unison ring and move the unison ring in a substantially axial direction. The plurality of bleed off doors are coupled to the unison ring and move therewith.

Abstract: 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.

Abstract: A system for controlling a flow of air applied to a gas turbine shroud in an airplane engine, the air flow being adjusted by a regulator valve of position that is controlled by a first control signal calculated on the basis of a first setpoint value corresponding to predetermined clearance between the rotor and the turbine shroud. The system further comprises means for calculating at least one second control signal based on a second setpoint value different from the first value representative of clearance in the turbine, and selector means for-selecting the second control signal to control the regulator valve in response to one or more engine parameters. With such a system, it is possible to control the air flow applied to the shroud in application of a plurality of control logic systems.