Abstract: The present invention provides a non-axisymmetric flow member disposed in a duct of a gas turbine engine. The flow member narrows the area aft of a vane to reduce the cross sectional area through which a wake from the vane traverses.

Abstract: A film cooling apparatus with a cooling hole (46) in a component wall (40). A first surface (42) of the wall is subject to a hot gas flow (48). A second surface (44) receives a coolant gas (50). The coolant flows through the hole, then downstream over the first surface (42). One or more pairs of cooperating electrodes (60-61, 62-63, 80-81) generates and accelerates a plasma (70) that creates a body force acceleration (71, 82) in the coolant flow that urges the coolant flow to turn around the entry edge (57) and/or the exit edge (58) of the cooling hole without separating from the adjacent surface (47, 42). The electrodes may have a geometry that spreads the coolant into a fan shape over the hot surface (42) of the component wall (40).

Abstract: A method of assembling an air distribution system for use in a rotor blade of a wind turbine. The rotor blade includes a sidewall at least partially defining a cavity extending from a blade root towards a blade tip. The method includes positioning at least a portion of a manifold within the cavity and coupling the manifold to the sidewall. The manifold extends from the blade root towards the blade tip and has a root end and an opposing tip end. A passage is defined from the root end to the tip end. A flow control device is coupled to the manifold root end and configured to channel air through the manifold. A bypass flow assembly is coupled to the manifold and configured to channel air through the air distribution system with the flow control device in a non-operating configuration.

Abstract: A radial compressor includes an impeller which is rotatively driven, axially introduces air taken in through an air inlet passage formed in a housing, pressurizes the introduced air, and discharges the pressurized air in a radial direction, wherein an annular concave groove is formed in a peripheral wall of the air inlet passage of the housing, a rear end portion of an opening of the annular concave groove, which rear end portion meets the housing peripheral wall, is provided in the vicinity of a blade front end surface of the impeller, and the rear end portion of the opening of the annular concave groove is formed such that an axial projecting amount X thereof relative to the blade front end surface of the impeller is set to ?1T?X?1.5T (where T denotes the thickness of the distal portion of a blade).

Abstract: A rotor blade assembly and a method for adjusting a loading capability of a rotor blade are disclosed. The rotor blade assembly includes a rotor blade having surfaces defining a pressure side, a suction side, a leading edge, and a trailing edge extending between a tip and a root. The rotor blade further defines a span and a chord. The rotor blade assembly further includes a spoiler assembly operable to alter a flow past a surface of the rotor blade. The spoiler assembly is incrementally deployable from the surface along one of a length or a width of the spoiler assembly.

Abstract: A free end of a blade of a fluid flow machine has a skeleton line camber distribution having an excessive value to a relative skeleton line camber of at least ?*=0.35 for a related running length of s*=0.1 in a blade profile flow line section between the free end and a blade section at 30% of a main flow path width from the free end. S* is a local running length relative to a total running length of the profile skeleton line and ?* is an angular change of the skeleton line relative to a total camber of the skeleton line from a leading edge to a related running length s*. The skeleton line camber distribution runs between leading edge point V (s*=0, ?*=0) and trailing edge point H (s*=1, ?*=1).

Abstract: Plasma generators (48, 49, 70, 71) in an endwall (25) of an airfoil (22) induce aerodynamic flows in directions (50) that modify streamlines (47) of the endwall boundary layer toward a streamline geometry (46) of a midspan region of the airfoil. This reduces vortices (42) generated by the momentum deficit of the boundary layer, increasing aerodynamic efficiency. The plasma generators may be arrayed around the leading edge as well as between two airfoils (22) in a gas turbine nozzle structure, and may be positioned at correction points (68) in streamlines caused by surface contouring (66) of the endwall. The plasma generators may be oriented to generate flow vectors (74) that combine with boundary layer flow vectors (72) to produce resultant flow vectors (76) in directions that reduce turbulence.

Abstract: A pressure relief device includes a main body and at least two air control units. The main body has at least one passage therein and two outlets formed at upper and lower ends of the passage. A lid is pivotally connected to each of the outlets. The two air control units are each disposed in the passage close to the lid. An air control device drives the air control units to open/close the lid which is adapted to open/close the passage.

Abstract: It is intended to provide a wind turbine blade with superior aerodynamic characteristics and rigidity, a wind turbine generator equipped with the wind turbine blade, and a method of designing the wind turbine blade. The wind turbine blade 1 is provided with a blade tip part 2, a blade root part 4 connected to a hub 112, and an airfoil part 6 disposed between the blade tip part 2 and the blade root part 4. The airfoil part 6 includes a flatback airfoil 10 with a blunt trailing edge 8 at least in an area in a longitudinal direction of the wind turbine blade 1 where a blade thickness ratio X is 40% to 50%. X indicates a ratio of a maximum thickness tMAX to a chord length LC. In the area where the blade thickness ratio X is 40% to 50%, the flatback airfoil has a trailing edge thickness ratio Y which is 5% to YH%, Y indicating a ratio of a trailing edge thickness tTE to the maximum thickness tMAX and YH has a relationship of YH=1.5X?30 with the blade thickness ratio X.

Abstract: A fluid flow machine has a flow path provided by a casing (1) and a rotating shaft (2), in which rows of rotor blades (3) and stator blades (4) are arranged, and includes at least one annular groove-type recess (5) being disposed in a blade (3, 4) tip area in an annulus duct wall of the casing (1) and/or the shaft (2). An upstream end point (E) of the recess (5) in a flow direction is set at a distance (e)>0 forward of a forward blade tip point (A), and a downstream end point (F) of the recess (5) is set at a distance (f) rearward of point (A), where: 0.5 L>(f)>0, and L is a distance between point (A) and a rearward blade tip point (B).

Abstract: A compressor typically for use in a turbocharger comprises a downstream radial compressor impeller wheel, an upstream axial compressor impeller wheel and an intermediate stator. The compressor housing has an inlet with inner and outer walls that define between them an MWE gas flow passage. An upstream opening defined by the flow passage provides communication between the passage and the intake and at least one first slot downstream of the upstream opening provides communication between the passage and the inner surface of the inner wall. The stator comprises a plurality of fixed vanes and is disposed in the inner wall of the inlet between the radial and axial impeller wheels. The position of the slot can be at one of several positions along the gas flow passage, in other embodiments there are second and third slots and the flow passage is divided into two parts. All the arrangements are designed to improve the compressor map width.

Abstract: A nacelle assembly includes an inlet lip section and an airfoil adjacent to the inlet lip section. The airfoil is selectively moveable between a first position and a second position to adjust the flow of oncoming airflow and to influence an effective boundary layer thickness of the nacelle assembly.

Abstract: A turbine stator vane with an airfoil extending from an endwall, a vortex flow retaining chamber formed within the endwall and wrapping around a leading edge region and opening onto a surface of the endwall in front of the leading edge region of the airfoil. A vortex flow tube is formed within the endwall and extends from one side to the opposite side of the endwall and passes through the vortex flow retaining chamber. Cooling air supply holes open into the vortex flow tube to produce a vortex flow in a direction opposite to a vortex flow of the hot secondary gas that flows into the vortex flow retaining chamber. A row of exit cooling holes are connected to the vortex flow tube and open onto the endwall surface to discharge the hot secondary flow that flows into the chamber and is mixed with the cooling air supplied through the supply holes.

Abstract: A fluid flow machine has a main flow path 2 which is confined by a hub 3 and a casing 1 and in which at least one row of blades 5 is arranged. A gap 11 is provided on at least one blade row 5 between a blade end and a main flow path confinement, with the blade end and the main flow path confinement performing a rotary movement relative to each other. At least one reversing duct 7 is provided in the area of the blade leading edge in the main flow path confinement at a discrete circumferential position. The reversing duct 7 connects two openings 12, 13 arranged on the main flow path confinement.

Abstract: An instability mitigation system is disclosed, comprising a rotor having a row of blades arranged around a centerline axis, and a mitigation system comprising at least one plasma actuator mounted on a blade that facilitates the improvement of the stability of the rotor, and a control system for controlling the operation of the mitigation system. An instability mitigation system further comprising a detection system for detecting an onset of an instability in a rotor and a control system for controlling the detection system and the mitigation system are disclosed.

Abstract: A system and method is provided for detecting and controlling rotating stall in the diffuser region of a compressor. A pressure transducer is placed in the gas flow path downstream of the impeller, preferably in the compressor discharge passage or the diffuser, to measure the sound or acoustic pressure phenomenon. Next, the signal from the pressure transducer is processed either using analog or digital techniques to determine the presence of rotating stall. Rotating stall is detected by comparing the detected energy amount, which detected energy amount is based on the measured acoustic pressure, with a predetermined threshold amount corresponding to the presence of rotating stall. Finally, an appropriate corrective action is taken to change the operation of the compressor in response to the detection of rotating stall.

Abstract: An exhaust diffuser system and method for a turbine engine includes an inner boundary and an outer boundary with a flow path defined therebetween. The inner boundary is defined at least in part by a hub that has an upstream end and a downstream end. The outer boundary has a region in which the outer boundary extends radially inward toward the hub. The region can begin at a point that is substantially aligned with the downstream end of the hub or, alternatively, at a point that is proximately upstream of the downstream end of the hub. The region directs at least a portion of an exhaust flow in the diffuser toward the hub. As a result, the exhaust diffuser system and method can achieve the performance of a long hub system while enjoying the costs of a short hub system.

Abstract: A row of stator blades in a compressor of a combustion turbine engine, the combustion turbine engine including a diffuser located downstream of the compressor, and the row of stator blades disposed in close proximity to the diffuser; the row of stator blades comprising: a plurality of stator blades that include at least one of an inboard forward notch and an outboard forward notch.

Abstract: A method of operating a compressor having a row of blades for preventing a compressor stall is disclosed, the method comprising the steps of mounting a plasma generator in a casing or a shroud radially outwardly and apart from the blade tips wherein the plasma generator comprises a radially inner electrode and a radially outer electrode separated by a dielectric material; and supplying an AC potential to the radially inner electrode and the radially outer electrode.

Abstract: A nozzle segment is disclosed that includes a first platform configured to be connected with a casing of a gas turbine engine, a second platform configured to be connected with the casing, the second platform being disposed apart from the first platform in a radial direction of the gas turbine engine, and a plurality of blades disposed between the first and second platforms and connected thereto, a portion of an inside surface of the first platform or the second platform being scalloped so as to increase a throat area of the nozzle segment in order to increase the nozzle segment's mass flow rate handling capability.

Abstract: In a compressor of a turbocharger of an internal combustion engine comprising a housing with an inflow channel; a compressor impeller arranged in the inflow channel; and a bypass channel which has a first flow opening upstream of the compressor impeller inlet, a second flow opening downstream of the compressor impeller inlet and an axial annular chamber which connects the first and the second flow openings. An axial vane structure is provided in the axial annular chamber of the bypass channel, which imparts a swirl direction corresponding to that of the air mass flow in the inflow channel to the air mass flowing through the bypass channel from the second to the first flow opening.

Abstract: An axial compressor comprises a stator component and a rotor component, which cooperate to perform work on a fluid flow in a primary flow-passage defined by the stator and rotor components, the stator component and the rotor component further defining a secondary flow-passage which interconnects a higher pressure region and a lower pressure region of the primary flow-passage, the rotor component being provided with at least one secondary rotor element which, in normal operation of the machine, pumps a bypass flow of fluid through the secondary flow passage from the lower pressure region to the higher pressure region.

Abstract: An instability mitigation system is disclosed, comprising a detection system for detecting an onset of an instability in a rotor during the operation of the rotor, a mitigation system that facilitates the improvement of the stability of the rotor when the onset of instability is detected by the detection system, a control system for controlling the detection system and the mitigation system.

Abstract: An instability mitigation system is disclosed, comprising a stator stage located axially proximate to a rotor, the stator stage having a row of a plurality of stator vanes arranged around a centerline axis, and a mitigation system comprising at least one plasma actuator mounted on the stator vane that facilitates the improvement of the stability of the rotor, and a control system for controlling the operation of the mitigation system. An instability mitigation system further comprising a detection system for detecting an onset of an instability in a rotor and a control system for controlling the detection system and the mitigation system are disclosed.

Abstract: With highly loaded rotors and stators problems can occur with secondary flows sweeping low momentum fluid across the blades reducing efficiency. By provision of collector slots to collect the secondary air and direct that air to a return slot in a rotor hub it is possible to provide impetus to the collected secondary flow to an outlet slot such that there is dispersal of the secondary flow and therefore reduce the effects upon the overall performance of a gas turbine engine incorporating the arrangement.

Abstract: An active flow control system for manipulating a boundary layer of air across a wind turbine rotor blade. The wind turbine rotor blade has at least one sidewall defining a cavity therein. The sidewall extends between a leading edge and an axially spaced trailing edge. The active flow control system includes an air discharge assembly that is coupled to the sidewall. The air discharge assembly is configured to selectively discharge air from within the wind turbine rotor blade into the boundary layer. An air suction assembly is coupled to the sidewall and to the air discharge assembly. The air suction assembly is configured to channel air from the boundary layer to the air discharge assembly.

Abstract: A turbomachine includes at least one rotor 1 featuring a hub and one stator 2, with a casing 5 confining the flow through the rotor 1 and the stator 2 to the outside. It also includes running-gap adjacent hub and casing surfaces LNGO, with openings being provided along the circumference on at least one running-gap adjacent hub and casing surface LNGO which form at least one dynamically operating supply point DAV which is connected via a least one line 6 to at least one opening on a blade-passage confining surface SKO.

Abstract: A high drag vortex generator for inhibiting the formation of boundary layer gas flow is mounted in a duct having a strong adverse pressure gradient with high speed gas flow over a primary aerodynamic. The vortex generator includes a mounting for supporting the high drag vortex generator at a spatial interval in the high speed gas flow above the primary aerodynamic surface. The high drag vortex generator is supported from the mounting and extends from the mounting in the high speed gas flow less than the full distance to the primary aerodynamic surface. Each high drag vortex generator includes first and second vanes. The first and second vanes have flow confining components at substantial right angles to one another. The vanes flow confining components deflect fluid flow towards and/or away from the primary aerodynamic surface inhibiting boundary layer formation downstream of the vortex generator.

Abstract: A fluid flow machine includes a flow path formed by a casing 1 and a rotating shaft 2, in which rows of blades 3 are provided, with at least one annular groove-type recess 4 being arranged in a blade 3 tip area in an annulus duct wall of the shaft 2. A cross-section and the position of the recess 4 are defined as follows: the axial arrangement is selected such that a partial length A of the recess 4 extends by max. 30 percent of the chord length of a blade 3 from the blade 3 leading edge against the direction of flow, and up to max. 50 percent of the chord length of the blade 3 from the blade 3 leading edge in the direction of flow.

Abstract: The present invention provides a turbine blade that is exposed to a high temperature atmosphere at relatively high temperatures and to a low temperature atmosphere at relatively low temperatures, including: a concave-curve-like pressure surface that is concaved along a flow direction of fluid; a convex-curve-like suction surface that is convexed along the flow direction of the fluid; and a concave and convex formation device which causes a concave and convex portion along the flow of the fluid to appear on the suction surface when the turbine blade is exposed to the low temperature atmosphere, and which turns the suction surface into a smooth surface along a convex curve surface of the suction surface when the turbine blade is exposed to the high temperature atmosphere.

Abstract: A fluid flow machine includes a main flow path in which at least one row of blades (3, 4) is arranged, and a blade shroud (2) which forms a confinement of the main flow path in the area of a blade row (3, 4) and is arranged in a recessed cavity (10) of a surrounding physical structure (11). The cavity (10) so formed between the shroud (2) and the surrounding physical structure (11) is provided with an annular connection (14) to the main flow path at least on the outflow side of the blade shroud (2), and with the blade shroud (2) being penetrated by at least one secondary flow path (16) which, commencing at a location of high pressure on the shroud (2), issues at a surface wetted by the main flow.

Abstract: Systems and methods involving variable throat area vanes are provided. In this regard, a representative gas turbine engine includes: a vane extending into a gas flow path and having: an interior operative to receive pressurized air; a pressure surface portion; and a first port communicating between the interior and pressure surface portion, the first port being operative to receive the pressurized air from the interior and emit the pressurized air, wherein the emitted pressurized air displaces the gas flow path such that a throat area defined, at least in part, by the vane is modified.

Abstract: Reinjecting air to the inlet of a compressor in order to improve its pumping margin. The inlet stator of the compressor includes injection holes passing through at least some of the vanes in the vicinity of their pivots.

Abstract: The invention relates to a wind turbine blade comprising one or more turbulence generating strips, where the strips are placed on a surface of the blade. The blade is characterized in that at least one joint area of the turbulence generating strips and the surface of the blade are completely or partially covered by sealing means. The invention further relates to a pitch controlled wind turbine comprising at least two pitch controlled wind turbine blades and pitch controlling means for pitching the blades. The pitch controlled wind turbine is characterized in that the blades comprises one or more turbulence generating strips, wherein at least one joint area of the turbulence generating strips and a surface of the blades are completely or partially covered by sealing means.

Abstract: A wind turbine comprising at least a blade (11) having an aerodynamic profile with a leading edge (13), a trailing edge (15) and suction and pressure sides between the leading edge (13) and the trailing edge (15) that includes an anti-noise device (31) placed on the suction side formed by elements (33) that modify the frequency spectra of the boundary layer noise. Preferably the anti-noise device (31) is placed on the suction side between two sections corresponding to chord positions in the range of 40%-95% of the chord length, measured from the leading edge (13).

Abstract: In a transonic region with a Reynolds number not more than a critical Reynolds number, a flow velocity distribution on an extrados of an airfoil has a single supersonic maximum value within a range of up to 6% from a leading edge on a chord, or a shape factor has a maximum value in a region of 6 to 15% from the leading edge on the chord, the value being nearly constant in a region of 30 to 60% and gradually can increase up to 2.5 in a region downstream of 60% of chord. A pressure loss in a low Reynolds number region can be drastically reduced, while conventionally keeping low the pressure loss in a high Reynolds number region. Moreover, this pressure-loss reduction effect in the low Reynolds number region is exerted even if an inflow angle is changed in a wide range.

Abstract: An axial flow turbine stage structure has: an annular diaphragm inner ring; an annular diaphragm outer ring arranged radially outside and coaxially with the diaphragm inner ring and separated from the diaphragm inner ring by an annular flow path interposed between them; stationary blades arranged peripherally at intervals in the annular flow path and rigidly secured to the diaphragm inner ring and the diaphragm outer ring; and moving blades rigidly secured to the outer periphery of a rotatable rotor and arranged peripherally at intervals respectively at axially downstream sides of the stationary blades. Through holes are formed in the diaphragm outer ring so as to allow axial upstream side and axially downstream side of the stationary blades to communicate with each other.

Abstract: A blade surrounded by: a leading edge portion, which is where fluid flows in; a pressure surface which is a concave curve shape concaved along a direction the fluid flows; a suction surface which is a convex curve shape convexed along a direction the fluid flows; and a trailing edge portion, which is where the fluid flows out, in which a disturbance addition portion is provided on the suction surface closer to the leading edge portion than a maximum velocity point, where the flow velocity of the fluid is maximized on the suction surface, for disturbing a laminar flow on the suction surface and for transitioning to a turbulent flow.

Abstract: A fan assembly for creating an air current is described. The fan assembly includes a nozzle mounted on a base housing a device for creating an air flow through the nozzle. The nozzle includes an interior passage for receiving the air flow from the base, a mouth through which the air flow is emitted, the mouth being defined by facing surfaces of the nozzle, and spacers for spacing apart the facing surfaces of the nozzle. The nozzle extends substantially orthogonally about an axis to define an opening through which air from outside the fan assembly is drawn by the air flow emitted from the mouth. The fan provides an arrangement producing an air current and a flow of cooling air created without requiring a bladed fan. The spacers can provide for a reliable, reproducible nozzle of the fan assembly and performance of the fan assembly.

Abstract: An airfoil disclosed herein comprises a pressure surface 42 exposed to a stream of fluid, a suction surface 40 exposed to the stream of fluid and a passage 56 extending from a passage intake end 60 to a passage discharge end 66. The intake end has an intake opening 62 penetrating the pressure surface for extracting fluid from the fluid stream. The discharge end has a discharge opening 68 penetrating the suction surface upstream of a natural separation point 52. The discharge end is configured to inject the extracted fluid into the fluid stream at a jet angle whose components include at least one of a nonzero streamwise angle ? in a prescribed angular range and a nonzero cross-stream angle ?.

Abstract: A turbomachine has at least one stator (4) and at least one downstream rotor (7), with the stator (4) being provided with stationary stator blades and the rotor (7) comprising several rotor blades attached to a rotating shaft. A casing (2) confines the passage of fluid through the rotor (7) and the stator (4) in the outward direction. A mechanism for peripheral jet creation is provided in the area of at least one vane of the stator (4), with at least one nozzle (21) issuing fluid at the radially outer boundary of a main flow path.

Abstract: A system and method is provided for detecting and controlling rotating stall in the diffuser region of a compressor. A pressure transducer is placed in the gas flow path downstream of the impeller, preferably in the compressor discharge passage or the diffuser, to measure the sound or acoustic pressure phenomenon. Next, the signal from the pressure transducer is processed either using analog or digital techniques to determine the presence of rotating stall. Rotating stall is detected by comparing the detected energy amount, which detected energy amount is based on the measured acoustic pressure, with a predetermined threshold amount corresponding to the presence of rotating stall. Finally, an appropriate corrective action is taken to change the operation of the compressor in response to the detection of rotating stall.

Abstract: A blade row for a rotary machine includes a radially inner endwall, a radially outer endwall having a radially inner surface, and an airfoil including a root section coupled to the inner endwall and a tip section coupled to the outer endwall such that the airfoil extends radially between the inner and outer endwalls. The radially outer endwall includes a projection extending radially inward from the outer endwall inner surface such that the outer endwall inner surface includes a non-axisymmetric shape.

Abstract: A supersonic aircraft engine axial fan that includes a rotating blade row having blades that receive a supersonic entry flow in the absolute frame and decelerate the flow to a lower supersonic or subsonic velocity exit flow while adding work to the flow to increase stagnation pressure. It is preferred that the lower velocity be subsonic. When the fan is combined with a suitably designed inlet, the propulsion system is compact and lighter in weight than conventional engines for supersonic aircraft.

Abstract: Devices for intaking gases from the boundary layer through ducts hollowed out in blades. These devices are suitable for variable setting blades, installed on the stator using pivots. The seal is maintained despite the variation in the relative position between the moving pivots through which ducts pass, and the fixed intake header pipe.

Abstract: A gas turbine engine plasma blade tip clearance control system includes an annular shroud surrounding rotatable blade tips and an annular plasma generator spaced radially outwardly and apart from the blade tips. An exemplary embodiment of the annular plasma generator is mounted to the annular shroud and includes radially inner and outer electrodes separated by a dielectric material disposed within an annular groove in a radially inwardly facing surface of the annular shroud. The plasma generator is operable for producing an annular plasma between the annular shroud and blade tips and an effective clearance produced by the annular plasma between the annular shroud and blade tips that is smaller than a clearance between the annular shroud and blade tips.

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: A centrifugal compressor for compressing a fluid comprises a compressor wheel having a plurality of circumferentially spaced blades, and a compressor housing in which the compressor wheel is mounted. The compressor housing includes an inlet duct through which the fluid enters in an axial direction and is led by the inlet duct into the compressor wheel, and a wheel shroud located radially adjacent the tips of the blades. A bleed port is defined in the wheel shroud at a location intermediate the leading and trailing edges of the blades for bleeding off a bleed portion of the fluid being compressed by the compressor wheel. A converging injection nozzle is defined in the duct wall upstream of the leading edges of the blades for injecting the bleed portion of the fluid back into the main fluid flow stream approaching the compressor wheel.

Abstract: A plasma boundary layer lifting system includes at least one gas turbine engine vane having a spanwise extending airfoil with an outer surface extending in a chordwise direction between opposite leading and trailing edges and chordwise spaced apart plasma generators for producing a plasma extending in the chordwise direction along the outer surface. Each plasma generator may include inner and outer electrodes separated by a dielectric material disposed within a spanwise extending groove in the outer surface. The airfoil may be hollow having an outer wall and the plasma generators being mounted on the outer wall. A method for operating the system includes forming a plasma extending in the chordwise direction along the outer surface of the airfoil. The method may further include operating the plasma generators in steady state or unsteady modes.

Abstract: A trailing edge vortex reducing system includes a gas turbine engine airfoil extending in a spanwise direction, one or more spanwise extending plasma generators in a trailing edge region around a trailing edge of the airfoil. The plasma generators may be mounted on an outer wall of the airfoil with first and second pluralities of the plasma generators on pressure and suction sides of the airfoil respectively. The plasma generators may include inner and outer electrodes separated by a dielectric material disposed within a grooves in an outer hot surface of the outer wall of the airfoil. The plasma generators may be located at an aft end of the airfoil and the inner electrodes flush with the trailing edge base. A method for operating the system includes energizing one or more of plasma generators in steady state or unsteady modes.