Abstract: An adjustable bleed apparatus and method for bleeding air to or from the impeller inlet region of a centrifugal compressor through a segmented annular bleed slot. An annular support member, connected to the fixed intake casing, supports a downstream annular shroud segment and an upstream annular shroud segment such that the downstream end of the downstream shroud segment is unconstrained. The spaced-apart distance between the shroud segments defines an annular bleed slot, which is segmented by bridge members. The connections between the shroud segments and the annular support member are configured such that the spaced apart distance of the annular bleed slot is adjustable. The connections between the annular support member and the fixed intake casing are configured such that the running clearance between the downstream shroud segment and the impeller is adjustable independently from the width of bleed slot.

Abstract: A turbomachine includes a first turbine portion and a second turbine portion. A flow splitter assembly is coupled between the first and second turbine portions. The flow splitter assembly includes a first end portion including a first mounting member and a second end portion including a second mounting member. A flow diverting member is positioned between the first and second end portions. The flow diverting member includes a first end section having a first mounting element coupled to the first mounting member and a second end section having a second mounting element coupled to the second mounting member. The flow diverting member includes a flow diverting surface. A first locking member engages the first mounting element and the first mounting member, and a second locking member engages the second mounting element and the second mounting member.

Abstract: The invention is directed to a turbine (3) for generating power in a drill string (2), including a turbine rotor (10) drivable by a fluid and a bypass device having a first bypass channel bypassing the turbine rotor (10). For varying the volumetric flow through the bypass channel, provision is made for a valve (28) and a control device driven by the fluid and having a drive element (24) by means of which the valve (28) is movable anywhere between a first and a second position. The drive element (24) produces a flow resistance in the fluid flow path downstream from the turbine rotor (10) and downstream from the valve (28) and in the flow direction takes support upon a spring (26).

Abstract: An active tip clearance control (ATCC) apparatus of an aircraft gas turbine engine is situated within a pressurized core case of the engine. First and second portions of bypass air flow passing through the respective compressed core compartment and the ATCC apparatus, are isolated from one another in order to improve both the ATCC performance and bypass air duct performance.

Abstract: The relief device is designed to divert a portion of a main flow toward a by-pass flow in a turbojet that includes a main flow stream, a by-pass flow stream and an intermediate casing. The relief device includes inner openings of the inner shroud, outer openings of the outer shroud, a relief circuit connecting the inner openings to the outer openings, and a blanking device capable of opening and closing the relief circuit. The relief circuit includes upstream holes of the intermediate casing, downstream holes of the intermediate casing, and an annular channel defined between the intermediate casing, the outer shroud and a connecting partition. The connecting partition rests on the intermediate casing so as to encompass the downstream holes of the intermediate casing in the annular channel and resting on the outer shroud so as to encompass the outer openings of the outer shroud in the annular channel.

Abstract: The present invention relates to a rotorcraft blade (1) provided with a dynamic balancing system having a balancing device (10). The balancing device (10) is embedded in the inside (INT) of the body of the blade (1) in such a manner that it does not open to the outside (EXT) of said blade (1).

Abstract: Provision of flow control arrangements which can be switched from high flow restriction to low flow restriction or vice versa are desirable. It is known to use switched vortex valves comprising a vortex chamber having a diverter portion which jets a through flow into either radial presentation to an outlet or tangential presentation with differential flow restrictions. Typically deflection is through cross flows presented from control ports as determined by mechanically actuated and controlled valves. Sensors are required to determine pressure differentials in order to actuate these moving part valves. By providing association between the control ports and an outlet path along with appropriate configuration and sizing of the flow regulators switching states can be provided through the cross flows presented via the flow regulators between the high and low through flow restriction. The flow regulators generally comprise orifice restrictors, diffusers/venturi arrangements or vortex throttles.

Abstract: A turbomachine includes a compressor having a compressor intake, and an intake system mounted upstream of the compressor intake. The intake system includes a housing and a plurality of selectively positionable vanes arranged within the housing. The plurality of selectively positionable vanes are moveable between a first position to remove a first amount of foreign particles, and a second position, to remove a second amount of foreign particles.

Abstract: A compressor having a variable-geometry ported shroud includes a compressor housing defining an inlet duct, a shroud, and a bypass passage. A variable-geometry port extends through the shroud into the bypass passage, and includes an adjustable mechanism that is selectively configurable to adjust the meridional location of the port between at least first and second meridional locations. In one embodiment an opening extends through the shroud, and the adjustable mechanism includes a bypass control device that is disposed within the opening and is axially movable therein. The bypass control device has an axial length less than that of the opening such that there is always a portion of the opening that remains unblocked by the bypass control device and forms a port through the shroud. The bypass control device is axially movable between at least first and second positions to place the port at the first and second meridional locations.

Abstract: A system for providing air from a multi-stage to a turbine includes a turbine having a high pressure input port and a low pressure input port. The system also includes a compressor having at least one high pressure extraction air output and at least one low pressure extraction air output. A valve is fluidly connected to the at least one high pressure extraction air output, at least one low pressure extraction air output and low pressure input port of the turbine. The valve is selectively operated to fluidly connect the at least one low pressure extraction air output with the low pressure input port during normal operating conditions and fluidly connect the at least one high pressure extraction air output and the low pressure input port during a turn down condition or below design temperature operation to enhance turbine engine performance.

Abstract: An engine system includes an internal combustion engine and an air intake assembly configured to supply a flow of air to one or more cylinders of the internal combustion engine. The air intake assembly includes an air supply conduit with a turbine therein, a bypass conduit configured to bypass the turbine, and a fresh air valve configured to selectively direct the flow of air to one or both of the air supply conduit and the bypass conduit. An exhaust conduit is configured to receive a flow of exhaust gas from the internal combustion engine and direct the flow of exhaust gas away from the internal combustion engine. A compressor in the exhaust conduit, which may be coupled to the turbine by a shaft, is configured to rotate and thereby reduce backpressure in the exhaust conduit during idle and partial throttle conditions.

Abstract: A turbocharger includes an annular bypass passage to allow exhaust gas to bypass the turbine wheel. An annular bypass valve is disposed in the bypass passage. The bypass valve comprises a fixed annular valve seat and a rotary annular valve member arranged coaxially with the valve seat. The valve member is disposed against the valve seat and is rotatable about the axis for selectively varying a degree of alignment between respective orifices in the valve seat and valve member. The turbine includes a nozzle having a vane assembly. In one embodiment the vane assembly has only fixed vanes. In another embodiment the vane assembly includes both fixed vanes and variable vanes, and the variable vanes are part of a rotor that rotates together with the rotary valve member.

Abstract: A gas turbine engine has an annular bypass duct defined between a fan nacelle and a core fairing through which fan air is discharged. One or more bifurcator members are positioned inside the exit of the bypass duct to bifurcate the air flow in the bypass duct around the bifurcator members prior to exiting the bypass duct. The bifurcator members modify the flow field of the air in the bypass duct radially and/or tangentially with respect to the axis of the annular bypass duct.

Abstract: Providing a double casing type pump and a head adjusting method thereof, wherein, the performance of the pump, especially, the head of the pump can be adjusted without large-scale work as well as without component disassembly. A double casing type pump, comprising: a plurality of impellers fixed to a rotation shaft, an inner casing that shrouding the impellers; an outer casing shrouding the inner casing and having a suction opening and a discharge opening; wherein, the pump further comprises: a space formed between the inner casing and the outer casing, the space communicating with the discharge opening; a bypass hole connecting the space to the working fluid passage in the inner casing, the pressure in the working fluid passage being lower than the stagnation fluid in the space; and, the bypass hole comprises a throat diameter adjusting member having a through-hole therein for determining the throat area of the bypass hole.

Abstract: A convertible fan system is disclosed having a fan rotor with a plurality of fan blades having an arcuate splitter that forms a portion of an inner flow passage and an outer flow passage wherein an inner portion of the fan blade pressurizes an inner flow in the inner flow passage and an outer portion of the fan blade pressurizes an outer flow in the outer flow passage and a vane system having an outer vane that is adapted to be able to vary the flow in the outer flow passage such that pressure ratio of the fan system can be varied.

Abstract: A method of starting up a turbine with a compressor have an inlet guide vane, a number of bleed valves, and a number of blades so as to limit the duration in a rotating stall window. The method may include starting rotation of the blades, increasing the speed of the rotation of the blades, closing one or more bleed valves so as to shift the rotating stall window to a higher rotational speed, partially opening the inlet guide vane to a position outside the rotating stall window, and opening the bleed valves while partially closing the inlet guide vane such that the number of blades pass through the rotating stall window.

Abstract: A bypass axial turbomachine includes a fan, a low pressure stator stage, a high pressure stator stage, the compressors being traversed by a flow referred to as primary flow of the turbomachine, at least one passage of discharge flow rate controlled from the primary flow in one of the stator stages toward the secondary flow, a heat exchanger of the surface air-oil (ACOC) type arranged flat on or in the wall surrounding the stator stages and defining the internal surface of the secondary flow, directly downstream of the junction of the passage of the discharge flow rate with said wall, so as to be run through by the secondary flow enriched with the discharge flow rate.

Abstract: Variable geometry turbine, in particular a gas turbine, comprising a housing (2, 22), a turbine rotor (4), a fluid inlet scroll (1) that surrounds the turbine rotor, a vaned nozzle interposed between the inlet scroll and the turbine rotor conceived to accelerate the flow of fluid, the nozzle comprising an axially adjustable ring (5, 25) conceived to vary the nozzle gap (3, 23) and having a wall (19) delimiting the nozzle gap, the wall having balance holes (6, 26, 26?) connecting the nozzle gap with a chamber (12) delimited by the housing and the ring, characterised in that the edges (16), formed by the holes with the surface (18) of the wall facing the nozzle gap, are rounded in the portion (17, 17?) located downstream the holes with respect to the fluid flow. Turbocharger and supercharged engine comprising the engine.

Abstract: A nacelle assembly for a high-bypass gas turbine engine includes a variable area fan nozzle having a first fan nacelle section and a second fan nacelle section. The second fan nacelle section being axially movable relative the first fan nacelle section to define an auxiliary port to vary a fan nozzle exit area and adjust fan bypass airflow, the second fan nacelle section includes at least one cowl with an inner portion, an outer portion and a multiple of stiffeners therebetween to increase a flutter margin.

Abstract: A turbocharger system comprises a first relatively small high-pressure (HP) turbocharger and a second relatively large low pressure (LP) turbocharger. The turbine of the LP turbocharger is connected in series downstream of the turbine of the HP turbocharger. A first exhaust bypass flow passage provides a bypass flow path around the HP turbine. A second exhaust bypass flow passage provides a bypass flow path around the LP turbine. A rotary valve is located at a junction of the first and second bypass flow passages and a first exhaust gas flow passage. The rotary valve comprises a valve rotor which is rotatable selectively to permit or block flow to the LP turbine and to permit or block flow to the first and second bypass paths.

Abstract: Gas turbine engine systems and related methods involving heat exchange are provided. In this regard, a representative heat exchange system for a gas turbine engine includes: a heat exchanger; and a flow restrictor operative to selectively restrict a flow of gas flowing along an annular gas flow path; in an open position, the flow restrictor enabling gas to flow along the gas flow path and, in a closed position, the flow restrictor restricting the flow of gas such that at least a portion of the gas is provided to the heat exchanger, the heat exchanger being located radially outboard of the flow restrictor.

Abstract: A turbocharger includes a turbine wheel mounted within a turbine housing and connected to a compressor wheel by a shaft. The turbine housing defines an exhaust gas inlet connected to a volute that surrounds the turbine wheel, and an axial bore through which exhaust gas that has passed through the turbine wheel is discharged from the turbine housing. The turbine housing further defines an annular bypass passage surrounding the bore and arranged to allow exhaust gas to bypass the turbine wheel. An annular bypass valve is disposed in the bypass passage. The bypass valve comprises a fixed annular valve seat and a rotary annular valve member arranged coaxially with the valve seat. The valve member is disposed against the valve seat and is rotatable about the axis for selectively varying a degree of alignment between respective orifices in the valve seat and valve member.

Abstract: A turbine assembly for an exhaust gas turbocharger has separate first and second volutes that are sequentially activated via a valve that receives exhaust gases from an engine. In a first position of the valve, only the first volute receives exhaust gas; in a second position, both volutes receive exhaust gases. In a third position, a bypass passage is also opened so that some exhaust gas bypasses the turbine wheel. Unlike conventional twin-scroll turbines, each volute receives exhaust gases from all engine cylinders, and the first volute feeds gas into the B-width portion of the wheel, while the second volute feeds gas into the wheel after the contour portion.

Abstract: The invention relates to an air discharge system for an aviation turbomachine compressor, the discharge system including at least one discharge valve (10) each valve comprising a valve body suitable for moving between an open position and a closed position of the discharge valve, a camshaft (22) comprising a cylindrical rod carrying a cam mounted to bear against the valve body in such a manner that rotating the camshaft causes the valve body to move between its open and closed positions, and means for keeping the cam bearing permanently against the valve body. The system further includes at least one main electric motor (30) having its rotor suitable for driving the camshaft of each discharge valve in rotation, and an electronic control system (32) for controlling each main electric motor.

Abstract: A gas turbine engine includes a compressor, combustor, and high pressure turbine operatively joined together. A first interstage bleed circuit is joined in flow communication between a first preultimate stage of the compressor and hollow blades in the turbine to provide thereto pressurized primary air. A second interstage bleed circuit is joined in flow communication between a second preultimate stage of the compressor and the turbine blades to provide thereto pressurized secondary air at a lower pressure than the primary air.

Abstract: To provide a turbofan jet engine for a supersonic aircraft, which enables a supersonic cruise with small additional drag due to engine installation while minimizing the jet noise at takeoff. In the turbofan jet engine, two fans, i.e., a front fan and an aft fan, are disposed, and an air inlet and an air outlet are provided on each of the front and aft fans, whereby at takeoff, air is introduced to each of the air inlets of the front and aft fans to drive the front and aft fans in parallel, and the air is then ejected from each of the outlets. On the other hand, during a supersonic cruise, an inlet of an aft fan duct is coupled to an outlet of a front fan duct, where external air is introduced only to an air inlet of the front fan duct and the air compressed through the front fan is fed to the aft fan duct and further compressed through the aft fan, and ejected from a nozzle, so that the front and aft fans are driven in series.

Abstract: A rotational flow balance system includes an actuator system which controls operation of both a fan duct blocker ring and a modulated exhaust cooling ring through a kinematic system. The kinematic system is controlled by the single actuator system such that the modulated exhaust cooling ring will remain in a fixed position while the fan duct blocker ring can be moved to satisfy operational requirements.

Abstract: A gas turbine engine includes a main compressor section for compressing air, a main combustor section positioned downstream of the main compressor section, a main turbine section positioned downstream of the main combustor section, and a spool extending from the main compressor section to the main turbine section. A second turbine is fluidically connected to the main compressor section by a bleed passage. An electrical generator is mechanically driven by both the spool and the second turbine.

Abstract: The bleed valve actuating system is for use in a gas turbine engine. The system comprises an actuator mounted adjacent the engine case and having a main actuation axis extending generally parallel to a surface of the engine case. A drive rod extends from the bleed valve to the actuator generally perpendicularly with reference to the main actuation axis of the actuator. The drive rod connected to the actuator via a linkage mechanically connects the actuator to the drive rod. The linkage is configured to convert axial motion of the actuator along the main actuation axis into generally perpendicular motion of the drive rod.

Abstract: A method for regulating fluid flow through a gas turbine engine is provided. The method includes coupling an outer fairing to a radially outer duct wall and coupling an inner fairing to a radially inner duct wall. An annular valve is coupled between the radially outer and the radially inner duct walls such that an outer bypass flow area is at least partially defined between the annular valve and the outer fairing, and such that an inner bypass flow area is at least partially defined between the annular valve and the inner fairing. The annular valve is selectively positioned between a first operational position and a second operational position, such that at least one of the outer bypass and the inner bypass flow areas is varied during a transition from a first turbine operation to a second turbine operation.

Abstract: A fuel pump includes a bearing element disposed in proximity to an impeller. The bearing element is configured to form a fuel seal with the impeller and to drain fuel leaked from the pump chamber back into the pump chamber. The bearing element minimizes leakage of fuel from the pump chamber into an engine coupled to the impeller. The fuel pump also includes redundant lip seals configured to provide redundant sealing relative to a shaft of the impeller within the fuel pump in order to minimize engine oil from entering the pump chamber and to minimize fuel from entering the engine. Integration of both the bearing element and the redundant lip seals as part of the fuel pump results in the fuel pump having a relatively compact size.

Abstract: A turbine-intake tower for delivering wind to a turbine has a hollow support column, an intake nozzle assembly rotatably coupled to the support column, and a tower nozzle disposed within the support column. The intake nozzle assembly is configured to receive and to accelerate wind. The tower nozzle is configured to receive the wind from the intake nozzle assembly and to further accelerate the wind received from the intake nozzle assembly for delivery to the turbine.

Abstract: A flow control arrangement is provided for a gas turbine engine in which an inlet slot is positioned prior to a stationary structure such as a stator or pylon in order that air flow is bled or removed from a mainflow. The removed air passes through a passage duct and is re-injected through an outlet nozzle with an askew angle consistent with an angle of rotor blades of a turbine. In such circumstances, distortion in the flow due to the stationary structure is relieved such that there is less instability downstream from that structure.

Abstract: An exhaust turbo supercharger is provided that is able to effectively prevent exhaust gas that has flowed into between a turbine housing and a bearing housing from leaking to outside of the device, while allowing an excellent level of workability in assembly and disassembly. Included is a turbine housing, into which exhaust gas from an internal combustion engine is introduced; a turbine wheel which is provided within the turbine housing and which is rotation-driven by the exhaust gas; a turbine shaft, one end of which is inserted into the turbine housing, and to which the turbine wheel is attached; a bearing which supports the turbine shaft; and a bearing housing which is connected to the turbine housing and inside of which the bearing is housed.

Abstract: A gas turbine engine is disclosed having a compressor, a combustion chamber downstream of the compressor, and a turbine downstream of the combustion chamber. The compressor includes a first group of compressor blade wheels and a second group of compressor blade wheels downstream of the first group of compressor blade wheels and rotating in an opposite direction such that a deviation of a performance, defined as air mass flow per second between the first group of compressor blade wheels and the second group of compressor blade wheels, is minimized, and a detachment of the air mass flow from blades of a first blade wheel of the second group of compressor blade wheels is substantially eliminated. A bypass enables the air mass flow flowing from the first group of the compressor blade wheels to substantially serve all of the second group of compressor blade wheels substantially simultaneously.

Abstract: A turbine engine component of a turbofan engine fitted with a bypass air valve includes at least one turbine engine component having a surface with at least one aperture, said turbine engine component located from between a bypass fan duct and a turbine exhaust nozzle of the turbofan engine; a bypass air valve includes a liner concentrically disposed about the turbine engine component and parallel to a centerline of the turbofan engine, the liner has a surface including at least one aperture and at least one impermeable region, and means for actuating the liner about the turbine engine component; and a flow transfer location comprising an area proximate to a turbine exhaust stream flow.

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: Systems and methods for controlling compressor extraction air flows from a compressor of a turbine system during engine turn down are provided. In one embodiment, a method for controlling compressor extraction air flows from a compressor of a turbine system during turn down includes a control unit monitoring one or more operating parameters of a turbine system associated with an exit temperature of a combustor of the turbine system. The method further includes the control unit detecting one or more operating parameters meeting or exceeding a threshold associated with a decrease in the exit temperature of the combustor. In response to detecting one or more operating parameter meeting or exceeding the threshold, a control signal is transmitted to at least one variable orifice located in the turbine system causing at least one variable orifice to alter at least one extraction air flow from the compressor.

Abstract: A running-gap control system of an aircraft gas turbine with a core engine including a turbine whose blade rows have a running gap 13 to the turbine casing 12, with the turbine casing 12 being surrounded by a core-engine ventilation compartment 15 enclosed by a fairing 2 of the core engine, with the fairing 2 forming an inner wall of a bypass duct 1. Air from the bypass duct 1 can be introduced via at least one inlet nozzle 5 into a cooling-air distributor 7 by a control system 6, 10, 11 and the air is subsequently returned to the bypass flow 3.

Abstract: A gas turbine engine is disclosed which comprises a bypass duct, a core engine, and a fluid mixing arrangement. The fluid mixing arrangement is configured to mix a bypass flow of fluid and a secondary flow of fluid, the secondary flow of fluid being drawn from the core engine. The arrangement comprises a flow-duct terminating with an outlet and being arranged to direct said secondary flow through the outlet and into the bypass flow. The arrangement is characterised by the provision of a delta-shaped wing in the region of the outlet, said wing extending at least partially across the duct and being configured to generate lift from said secondary flow effective to produce at least one trailing vortex extending into said bypass flow. The fluid mixing arrangement can be used as a ventilation arrangement or as part of a bleed valve arrangement in the gas turbine engine.

Abstract: A turbo fan engine is equipped with a main body portion that has a high-pressure turbine and a low-pressure turbine, an output portion that is provided separately from the main body portion, a gas flow channel that communicates with a space between the high-pressure turbine and the low-pressure turbine and is connected to the output portion, and a flow rate adjustment portion that is provided in the gas flow channel to adjust a flow rate of a gas flowing through the gas flow channel.

Abstract: A gas turbine engine has a plurality of radial struts in a bypass duct. At least one strut has a scoop incorporated with the fairing of the strut and in communication with an air passage of an engine secondary air system. The scoop faces a bypass air flow to scoop a portion of the bypass air flow using available dynamic pressure in the bypass duct. Scooped air may be provided, for example, to an active tip clearance control apparatus in a long duct turbofan engine.

Abstract: (A1)) A turbofan engine includes core and fan nacelles that provide a bypass flow path having a nozzle exit area. The bypass flow path carries a bypass flow to be expelled from the nozzle exit area. A turbofan is arranged within the fan nacelle and upstream from the core nacelle for generating the bypass flow. A flow control device includes a surface in the bypass flow path including an aperture. The flow device is adapted to introduce a fluid into the bypass flow path for altering a boundary layer of the bypass flow that effectively changes the nozzle exit area. In one example, bleed air is introduced through the aperture. In another example, pulses of fluid from a Helmholz resonator flow through the aperture. By decreasing the boundary layer, the nozzle exit area is effectively increased. By increasing the boundary layer, the nozzle exit area is effectively decreased.

Abstract: A compressor having a variable-geometry ported shroud includes a compressor housing defining an inlet duct, a shroud, and a bypass passage. A variable-geometry port extends through the shroud into the bypass passage, and includes an adjustable mechanism that is selectively configurable to adjust the meridional location of the port between at least first and second meridional locations. In one embodiment an opening extends through the shroud, and the adjustable mechanism includes a bypass control device that is disposed within the opening and is axially movable therein. The bypass control device has an axial length less than that of the opening such that there is always a portion of the opening that remains unblocked by the bypass control device and forms a port through the shroud. The bypass control device is axially movable between at least first and second positions to place the port at the first and second meridional locations.

Abstract: The invention relates to a bypass turbojet engine comprising 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. In characteristic manner, the IFD is fastened on the outer casing. The invention is applicable to a turbojet engine having a high pressure compressor of the axial-centrifugal type.

Abstract: A gas turbine engine (10) includes a fan (14), a nacelle (28) arranged about the fan, and an engine core at least partially within the nacelle. A fan bypass passage (30) downstream of the fan between the nacelle and the gas turbine engine conveys a bypass airflow (1) from the fan. A nozzle (40) associated with the fan bypass passage is operative to control the bypass airflow. The nozzle includes a shape memory material having a first solid state phase that corresponds to a first nozzle position and a second solid state phase that corresponds to a second nozzle position.

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: An example turbofan engine sound control system includes a core nacelle (12) housing a compressor and a turbine. The core nacelle is disposed within a fan nacelle (34). The fan nacelle includes a turbofan. A bypass flow path downstream from the turbofan is arranged between the two nacelles. A controller (50) is programmed to manipulate the nozzle exit area to control sound propagating from the engine. In one example, the controller manipulates the nozzle exit area using hinged flaps (42) to control engine sound. The hinged flaps open and close to adjust the nozzle exit area and the associated bypass flow rate.

Abstract: The invention relates to a turbine engine that includes a spool supporting at least one of a compressor and a turbine. A turbofan is coupled to the spool The spool is arranged in a core nacelle, and the turbofan is arranged upstream from the core nacelle. A fan nacelle surrounds the turbofan and the core nacelle and provides a bypass flow path. A structure extends radially between the core and fan nacelles to support the core nacelle relative to the fan nacelle. Surfaces are supported relative to the fixed structure and are moveable between closed and open positions to selectively obstruct bypass flow through the bypass flow path, thereby changing the effective area of the exit nozzle. A change in the effective area of the nozzle exit can be used to improve the efficiency and operation of the turbine engine.

Abstract: A turbomachine system comprises a first turbocharger comprising an exhaust gas flow first turbine for location in an exhaust path and a first compressor driven by said first turbine; a turbomachine for location in the exhaust path upstream or downstream of said first turbocharger and comprising an exhaust gas flow second turbine and a second compressor driven by said second turbine. The first turbine has an outlet that is in fluid communication with an inlet of the second turbine. One of said first and second turbines is a radial outflow turbine. The arrangement provides for a relatively compact package. The radial outflow turbine may have a particular structure in which there is provided a deflector member at or near its inlet for directing the gas outwards, a stator for introducing swirl and a downstream turbine rotor. A shroud is fixed to blades of the turbine rotor to prevent leakage and to provide additional structural rigidity.