JP2012514716A - Plasma assisted compressor duct - Google Patents

Abstract

A first compressor having a first flow path; a second compressor having a second flow path located axially rearward from the first compressor; and an air flow from the first compressor to the second A compression system is disclosed comprising a transition duct capable of flowing to two compressors, wherein the transition duct has at least one plasma actuator mounted therein. At least a portion of the second flow path can be located radially inward from a portion of the first flow path. The first compressor can also be a booster having a row of booster blades arranged circumferentially around the longitudinal axis.
[Selection] Figure 2

Description

  This invention relates generally to compressors, and more particularly to a compression system having a transition duct having a plasma actuator.

  In a gas turbine engine, air is pressurized with a compression module during operation. Air carried through the compression module is mixed with fuel in the combustor and ignited to generate hot combustion gases that flow through the turbine stage, which powers the fan and compressor rotor Then, energy is extracted from the high-temperature combustion gas and engine thrust is generated to propel the aircraft in flight or to power a load such as a generator.

  The compressor includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of rotor blades that extend radially outward from the disk. More specifically, each rotor blade extends radially from the platform adjacent to the disk to the tip. The gas flow path through the rotor assembly extends radially inward by the rotor blade platform and radially outward by a plurality of shrouds.

  The stator assembly includes a plurality of circumferentially spaced stator vanes or airfoils that direct compressed gas entering the compressor toward the rotor blades. The stator vanes extend radially between the inner and outer bands. The gas flow path through the stator assembly extends radially inward by the inner band and radially outward by the outer band. The rotor stage comprises rotor blades arranged circumferentially around the rotor hub. Each compression stage includes a vane stage and a rotor stage.

  Modern high bypass ratio gas turbine engines have a booster (low pressure compressor) and a high pressure compressor between which a transition duct is located. The geometry of a conventional transition duct or gooseneck duct is governed by the level of end wall curvature. This is because excessive bending leads to boundary wall delamination of the end walls and thus high efficiency loss. In order to ensure a smooth aerodynamic transition without flow separation, conventional transition duct designs must have some minimum axial length for a given change in the annular flow radius. This is undesirable because an increase in transition duct length directly leads to an increase in engine length, which increases engine weight and reduces engine backbone stiffness. This reduction in stiffness makes it difficult to maintain the desired clearance above the rotor tip, reducing engine efficiency and operability range.

  As compressors and booster rotors approach the limit of their ability to apply work / pressure to the air, they tend to be less efficient, and if this limit is exceeded, they will stall (required Pressure increase cannot be generated, leading to backflow through that stage and loss of engine thrust). Booster rotors that are designed very close to that limit after the booster can experience significant operability problems. This is a concern in conventional booster system designs that are limited to lower radii at the rear rotor stage. These can be corrected by pushing the rear end of the booster outward, which is possible by using a plasma actuator in the transition duct.

US Patent No. 3300121

  Therefore, it would be desirable to have a shorter transition duct design with an enhanced pressure distribution without causing flow separation in the duct. It would be desirable to have a booster system with a higher radius for the rear rotor stage without causing flow separation in the transition duct.

  The above-described need includes a first compressor having a first flow path, a second compressor having a second flow path located axially rearward from the first compressor, and a first air flow. Filled by an exemplary embodiment that provides a compression system comprising a transition duct capable of flowing from one compressor to a second compressor, wherein at least one plasma actuator is mounted therein. May be.

  In another aspect of the invention, the duct comprises an inlet portion, an outlet portion located at a distance axially rearward from the inlet portion, an axially arcuate inner wall extending between the inlet portion and the outlet portion, and the inlet An axially arcuate outer wall extending between the portion and the outlet portion, an axially arcuate channel between the inner wall and the outer wall, and at least one plasma actuator mounted in the duct.

  In another aspect of the invention, a gas turbine engine comprises a duct located between a first compressor and a second compressor, in which at least one plasma actuator is mounted. .

  The subject matter which is regarded as the invention is pointed out with particularity and clearly claimed in the concluding portion of the specification. However, the invention can best be understood by reference to the following description taken in conjunction with the accompanying drawings.

1 is a cross-sectional view of an exemplary gas turbine engine assembly comprising a compression system according to an exemplary embodiment of the present invention. FIG. 2 is an enlarged axial cross-sectional view from FIG. 1 showing a portion of a booster system according to an exemplary embodiment of the present invention. 1 is a schematic view of a gooseneck duct having a plasma actuator according to an exemplary embodiment of the present invention. FIG. 2 is an enlarged axial cross-sectional view of a portion of an exemplary duct having a plasma actuator system in an excitation mode. FIG. 1 is a schematic view of a gooseneck duct of a booster system according to an exemplary embodiment of the present invention. 1 is a schematic diagram of a booster system having a plasma actuator according to an exemplary embodiment of the present invention superimposed with a conventional booster flow path for comparison. FIG. 2 is a plot of pressure distribution in a booster system according to an exemplary embodiment of the present invention when a plasma actuator is excited and not excited.

  Referring to the drawings in which like reference numbers indicate like elements throughout the various views, FIG. 1 illustrates a longitudinal axis 11 and a first compressor 21 and a second axially rearwardly located from the first compressor 21. FIG. 2 shows a cross-sectional view of an exemplary gas turbine engine assembly 10 having a compression system 20 with a compressor 22 of FIG. In the exemplary embodiment shown in FIG. 1, the first compressor 21 is a booster 40, which is alternatively referred to herein as a low pressure compressor. The exemplary booster 40 shown in FIGS. 1 and 2 has four rotor stages, each rotor stage having between 50 and 90 booster rotor blades. The exemplary booster system 50 has a row of stator vanes (alternatively referred to herein as booster inlet guide vanes “IGVs”) located axially forward from the first booster rotor stage. The exemplary booster system 50 has a row of stator vanes (alternatively referred to herein as booster outlet guide vanes 44 “OGV”) located axially rearward from the last booster rotor stage. The OGV 44 has 120 vanes circumferentially spaced around the longitudinal axis 11. Further, the second compressor 22 shown in FIG. 1 is an axial high pressure compressor 14 (“HPC”). The exemplary HPC 14 shown in FIGS. 1 and 2 has seven rotor stages, each rotor stage having between 24 and 96 HPC rotor blades. The exemplary HPC 14 has a circumferential row of 40 stator vanes (also referred to herein as HPC inlet guide vanes “IGV”) located axially forward from the first HPC rotor stage. The exemplary embodiment of the gas turbine engine assembly 10 shown in FIG. 1 further includes a combustor 16, a low pressure turbine 19 coupled axially downstream from the high pressure turbine 18 and the core gas turbine engine 12, and the core gas turbine engine. 12, and a fan assembly 13 coupled to the upstream in the axial direction. The fan assembly 13 includes an array of fan blades 17 extending radially outward from the rotor disk 29. In the exemplary embodiment shown in FIG. 1, the fan assembly 13, booster 40, and low pressure turbine 19 are coupled together by a first rotor shaft 28, and the compressor 14 and high pressure turbine 18 are coupled to a second rotor shaft. 27 together.

  In operation, air flows through the fan assembly blade 17, a portion of that air flows as a bypass air stream 15, and a portion of the air includes a first compressor 21 and a second compressor 22. 20 flows into the core air stream 25. In the exemplary embodiment shown in FIGS. 1 and 2, the first compressor 21 is a booster 40 (low pressure compressor) and the second compressor 22 is a high pressure compressor 14. The core air stream 25 entering the compression system 20 is first carried through the first flow path 23 and compressed by the first compressor 21 (shown as a booster 40 in the figures herein). The core air stream 25 then passes through the arcuate third flow path 33 of the duct 30 (also referred to herein as the transition duct 30 or gooseneck duct 38) to the second compressor 22 (herein). To the second flow path 24 (shown as high pressure compressor 14 in the figure), where the core air stream 25 is further compressed. The air stream leaving the compression system 20 is conveyed to the combustor 16. Air is mixed with fuel in a combustor and burned. Combustion products from the combustor 16 are utilized to drive a high pressure turbine (HPT) 18 and a low pressure turbine (LPT) 19. In the exemplary embodiment shown herein, LPT 19 drives booster 40 and fan assembly 13 via fan rotor shaft 28 and HPT drives high pressure compressor 14 via HP rotor shaft 27. . The engine 10 can be operated in a range of operating conditions between designed operating conditions and non-design operating conditions. In the exemplary embodiment shown in FIGS. 1 and 2, the booster 40 rotor may have an operating speed between 1500 rpm and 2700 rpm, and the high pressure compressor 14 rotor operates between 6000 rpm and 12000 rpm. You may have speed.

  In the exemplary embodiment shown in FIGS. 1 and 2, the booster rotor stage pitch line is located radially with a higher radius than the high pressure compressor rotor stage pitch line. This is especially true for modern high bypass ratio engines. As used herein, the “pitch line” of the rotor stage is defined as the axis passing through the radial midpoint between the root and tip of the leading edge of the airfoil of the rotor blade of the rotor stage. . The transition duct 30 flows the core air flow 25 from the first flow path 23 of the booster 40 to the second flow path 24 of the high pressure compressor. 3 and 5 schematically show an axial cross-sectional view of an exemplary embodiment of a transition duct 30 according to the present invention. The terms “duct”, “transition duct” and “gooseneck duct” have the same meaning and are used interchangeably herein. Duct 30 includes an inlet portion 34 and an outlet portion 35 located axially rearward from the inlet portion. The inlet portion 34 has an inlet end 47 having an inlet area 36 and the outlet portion 35 has an outlet end 48 having an outlet area 37. The inlet portion 34 is located axially near the booster 40 and the outlet portion is located axially near the high pressure compressor 14. The inlet portion 34 is located radially outward from the outlet portion 35 and the centerline axis 11. The duct 38 includes an inner wall 31 and an outer wall 32 that form a flow path 33 therebetween. The duct 38 may have an annular shape around the longitudinal axis 11. In the exemplary embodiment shown in FIGS. 1, 2, and 3, the support frame struts 46 extend radially through the third flow path 33 of the duct 38 at several circumferential positions. The third flow path has a generally annular shape in the axial direction with respect to the longitudinal axis 11 and the struts 46 extend through the third flow path at a circumferential location in some applications. Due to the generally annular configuration of the duct 38 with the inlet portion 34 located radially outward from the outlet portion 35, the third flow path 33 and the duct 38 are, for example, those shown in FIGS. It has a gooseneck shape. Inner wall 31 and outer wall 32 have an arcuate shape in the axial direction, such as the one shown in FIGS.

  Referring to FIG. 5, the inlet end 47 of the duct 30 is located at a higher radius than the outlet end 48 with respect to the longitudinal axis 11. The outlet end 48 is located an axial rear distance 76 (“D”) from the inlet end 47. In the exemplary embodiment of the invention shown herein, the ratio of the inlet outer radius 71 (“RI”) to the outlet outer radius 72 (“RO”) is about 1.8. For the same duct shaft length, this ratio is about 1.6 or less in conventional designs. In the exemplary embodiment shown herein, the axial distance D76 is between about 16 inches and 18 inches. In the exemplary embodiment shown herein, the inlet area 36 is about 598 square inches and the outlet area 37 is about 570 square inches. A slight decrease in the outlet area 37 from the inlet area 36 may help to further reduce flow separation in the duct 38. In other embodiments of the invention, the inlet area 36 and outlet area 37 may have other suitable values. In an alternative embodiment, it may be advantageous to have an outlet area 37 that is larger than the inlet area 36 to improve the pressure distribution in the duct 38 using known design methods. The present invention allows the design of a booster system having a short duct axial length (“D”) compared to the inlet and outlet radii (“RI” and “RO”). In an exemplary embodiment of the invention, the aspect ratio, defined as the ratio (RI-RO) / D, is between about 0.5 and 0.8. Due to the geometrical nature of the cross-sectional shape of the third flow path 33 such as that shown in FIGS. 3 and 5, the transition duct 30 is alternatively referred to herein as the gooseneck duct 38.

  As is apparent from the exemplary embodiments shown herein, the inner wall 31 and the outer wall 32 have a substantial curvature in the axial direction. In the exemplary embodiment of the invention shown in FIGS. 1, 2, 3, 5 and 6, flow separation in the duct 38 is reduced by using a plasma actuator 60. The terms “plasma actuator” and “plasma generator” as used herein have the same meaning and are used interchangeably. For example, plasma actuators such as those shown as items 60, 61, 62, and 63 in the figures herein enhance the local axial moment of air flow near the walls 31, 32, and increase the inner and outer walls 31, 32. Minimize flow separation in the duct 30 in areas with sharp radii of curvature. The plasma actuator used as shown in the exemplary embodiment of the present invention generates a flow of ions and body forces acting on the fluid near the walls 31, 32, and the fluid from the walls 31, 32 In a state where the flow separation is reduced, the flow approaches the walls 31 and 32 so as to flow in a desired fluid flow direction.

  FIG. 4 illustrates an exemplary embodiment of a plasma actuator 60 for reducing flow separation in a transition duct 38 located between two compressors such as the booster 40 and HPC 14 shown in FIGS. This is schematically illustrated in the direction cross-sectional view. The exemplary embodiment of the invention shown herein facilitates improving the pressure distribution in the duct 38 in a gas turbine engine 10, such as the aircraft gas turbine engine illustrated in cross-section in FIG. 7) and / or increase the efficiency of the compression system. The exemplary gas turbine engine plasma actuators shown in FIGS. 1-6 include plasma actuators such as those shown as items 60, 61, 62 or 63 located on the inner wall 31, outer wall 32 or hub portion 45 of the booster OGV 44. . A plasma actuator such as item 60 shown in FIG. 4 is located in a groove 68 in a wall such as the inner wall 31. The plasma actuator 60 may be continuous in the circumferential direction located in the annular groove. Alternatively, the plasma actuator 60 may be split, in which case the plurality of plasma actuators 60 are located in corresponding groove sections spaced circumferentially of the walls 31, 32. The exemplary embodiment shown in FIG. 4 includes a plasma actuator 60 located in the groove 68 of the inner wall 31 of the duct 38. Alternatively, the plasma actuator 60 may be located elsewhere in the duct 38 where flow separation is likely, such as where the wall of the duct 38 has a sharp radius of curvature in the direction of air flow.

  The exemplary embodiment shown in FIG. 4 shows an annular plasma generator 60 attached to the inner wall 31 and includes a first electrode 64 and a second electrode 66 separated by a dielectric material 65. The dielectric material 65 is disposed in the annular groove 68 of the duct 38. An AC (alternating current) power supply 70 is connected to the electrodes 64, 66 to supply a high voltage AC potential in the range of about 3-20 kV to the electrodes 64, 66. When the AC amplitude is sufficiently large, the air is ionized in the region of maximum potential to form plasma 80. The plasma 80 generally begins near the edge 67 of the first electrode 64 exposed to air and extends over the area 69 projected by the second electrode 66 covered by the dielectric material 65. Plasma 80 (ionized air) in the presence of an electric field gradient creates a force on the air flow 25 near the wall 31 and induces a virtual aerodynamic shape that causes a change in pressure distribution across the inner wall 31 of the annular duct 38. To do. The air near the electrodes is weakly ionized and usually there is little or no air heating. The air flow 25 near the wall 31 tends to remain stuck to the wall 31, reducing the flow separation in the duct 38 and improving the pressure distribution due to the reduced pressure loss in the duct 38. Bring.

  FIG. 6 illustrates a booster system 50 according to an exemplary embodiment of the present invention. The booster system 50 shown in FIG. 6 has a final rotor stage 57 having a pitch line radius 54 that is larger than a conventional booster system. This is possible with the present invention by using a plasma actuator, such as that shown in FIG. 6 as items 60, 61, 62, in the duct 38 that receives flow from the last stage of the booster. The conventional flow path 90 of the conventional booster system is shown by a dotted line in FIG. 6 for comparison with the booster system 50 of the exemplary embodiment of the present invention. There are several benefits associated with having a booster stage further radially outward, such as the last rotor stage 57. A rotor stage 57 having a larger pitch line radius 54 has a higher tip speed compared to conventional designs. Since the ability of the rotor to work on the fluid is directly related to the tangential speed of the rotor, the exemplary embodiment of the present invention shown in FIG. In some applications, it is possible to reduce the number of stages required in the booster system by using the present invention for the desired pressure ratio, resulting in a significant weight reduction for the engine 10.

  In the exemplary embodiment of the invention shown in FIG. 6, the booster system 50 includes a plurality of first rotor blades 56 circumferentially spaced around the rotor hub 41 and extends from the longitudinal axis 11. The first rotor stage 55 having the pitch line radius 53 and the last rotor stage 57 located axially rearward from the first rotor stage 55 are included. The last rotor stage 57 has a plurality of last rotor blades 58 spaced circumferentially around the rotor hub 41 and has a second pitch line radius 54 extending from the longitudinal axis 11. The booster system has a gooseneck duct 38 located axially rearward from the last rotor stage 57 and receives the air flow 25 exiting from the last rotor stage 57. The gooseneck duct 38 has an inlet end 47 and an outlet end 48 located at a distance axially rearward from the inlet end 47 with at least one plasma actuator mounted therein. The geometry of the gooseneck duct 38 and the arrangement of the plasma actuators, such as those shown as items 60, 61, 62 in FIG. 6, for example, have been previously described herein. Unlike conventional booster systems, the last rotor stage 57 has a higher pitch line radius 54 “B” compared to the first rotor stage 55 pitch line radius 53 “A”. In the exemplary embodiment of the invention shown herein, the ratio B / A is at least 0.9.

  The exemplary booster system 50 shown in FIG. 6 has a gooseneck duct 38 located at the rear end, which has an axially arcuate inner wall 31 and an axially arcuate outer wall 32. The outlet end 48 has an outlet area 37 and the inlet end 47 has an inlet area 36. The gooseneck duct geometry (see FIG. 5) is such that the ratio RI / RO of the inlet outer radius 71 to the outlet outer radius 72 is at least 1.6. Plasma actuators, such as those shown as items 60, 61, 62 in FIG. 6, for example, are located in the duct 38 as previously described herein.

  A gas turbine engine 10 having a booster system 50 with a gooseneck duct 38 having a plasma actuator as described herein uses an AC potential from an AC power supply 70 to produce a first electrode 64 and a second electrode. It can be operated by exciting the electrode 66. By exciting the electrodes 64, 66 and generating the plasma 80, flow separation in the duct 38 can be reduced, which provides an advantage and improvement in the pressure distribution in the booster system 50. In one method, a plasma actuator, such as item 60 in FIG. 6, can be energized continuously throughout the engine operation period. Alternatively, the plasma actuator can be excited only during selected portions of the engine operation plan. The timing and duration of plasma actuator excitation can be determined by well-known engine test methods for determining engine operability.

  FIG. 7 shows an exemplary pressure distribution in the duct 38 at the outlet end 48 as determined by known fluid flow analysis methods. The horizontal axis indicates the normalized pressure in the duct 38, and the vertical axis indicates the radial span position in the duct 38. The distribution identified by reference numeral 91 indicates the radial pressure distribution at the outlet end 48 of the duct 38 when the plasma actuator 60 is not excited by the AC power supply 70. The distribution identified by reference numeral 92 shows the radial pressure distribution at the same location (exit end 48 of duct 38) when plasma actuator 60 is excited by AC power supply 70. It is clear that near the wall 32 where the plasma actuator is located (near the 1.0 span position), the normalized pressure increases from about 0.79 to about 0.86.

  As used herein, an element or step listed in the singular and following the word “a” or “an” is intended to refer to the plural element or step, unless such exclusion is expressly recited. It should be understood that it is not excluded. When introducing elements / components / steps etc. that design and / or manufacture the components and systems described and / or illustrated herein, the articles “a”, “an”, “the” and “ “side” is intended to mean that there is one or more of element (s) / component (s) / other. The terms “comprising”, “including” and “having” are inclusive and include the element (s) / component (s) / additional element (s) / component (s) other than those described. Yes) / is intended to mean that there may be others. Moreover, references to “one embodiment” of the present invention are not intended to be interpreted as excluding the existence of additional embodiments that also incorporate the recited features.

  Although the methods described herein and items such as vanes, outer bands, inner bands and vane sections are described in the context of a compressor used in a turbine engine, the vanes and vane sections described herein and It will be appreciated that their manufacturing or repair methods are not limited to compressors or turbine engines. The vanes and vane segments illustrated in the figures included herein are not limited to the specific embodiments described herein, but rather are independent of the other components described herein. Available separately.

  This document uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to make and use the invention. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are those where they have structural elements that are not different from the literal words of the claims, or they have equivalent structural elements that are only slightly different from the literal words of the claims. If included, it is intended to be within the scope of the claims.

  Although what has been described herein is considered to be a preferred and exemplary embodiment of the present invention, other variations of the present invention should be apparent to those skilled in the art from the teachings herein. Accordingly, it is desired that all such modifications fall within the true spirit and scope of the invention be assured by the appended claims.

DESCRIPTION OF SYMBOLS 10 Gas turbine engine assembly 11 Longitudinal axis 12 Core gas turbine engine 13 Fan assembly 14 Axial flow high pressure compressor 15 Bypass air flow 16 Combustor 17 Fan blade 18 High pressure turbine 19 Low pressure turbine 20 Compression system 21 First compressor 22 First 2 compressor 23 first flow path 24 second flow path 25 core air flow 27 second rotor shaft 28 first rotor shaft 29 rotor disk 30 duct 31 inner wall 32 outer wall 33 third flow path 34 inlet portion 35 outlet portion 36 inlet area 37 outlet area 38 gooseneck duct 40 booster 41 rotor hub 44 booster outlet guide vane 45 hub portion 46 support column 47 inlet end 48 outlet end 50 booster system 53 first pitch line radius 54 second pitch line Half 55 First rotor stage 56 First rotor blade 57 Last rotor stage 58 Last rotor blade 60 Plasma actuator 61 Plasma actuator 62 Plasma actuator 63 Plasma actuator 64 First electrode 65 Dielectric material 66 Second electrode 67 First electrode Edge 68 Groove 69 Area projected by second electrode 70 AC power supply 71 Outer inlet radius 72 Outer outlet radius 76 Axial rear distance 80 Plasma 90 Conventional channel 91 Radial pressure when plasma actuator is not excited Distribution 92 Radial pressure distribution when the plasma actuator is excited

Claims (21)

A first compressor having a first flow path;
A second compressor located second axially from the first compressor and having a second flow path;
A transition duct located between the first compressor and the second compressor capable of flowing an air flow from the first compressor to the second compressor, wherein the at least one plasma A compression system comprising a transition duct in which an actuator is mounted. The compression system of claim 1, wherein at least a portion of the second flow path is located radially inward from a portion of the first flow path. The compression system of claim 1, wherein the first compressor is a booster having a row of booster blades disposed circumferentially about a longitudinal axis. The compression system of claim 3, wherein the second compressor is an axial compressor having a row of compressor blades circumferentially disposed about the longitudinal axis. The compression system of claim 1, wherein the transition duct comprises an axially arcuate inner wall and an axially arcuate outer wall. The compression system of claim 5, wherein the inner wall and the outer wall form a third flow path having an inlet portion and an outlet portion located at a distance axially rearward from the inlet portion. The compression system of claim 6, wherein the inlet portion has an inlet area and the outlet portion has an outlet area that is larger than the inlet area. The compression system of claim 5, wherein the at least one plasma actuator is located on the inner wall. The compression system of claim 5, wherein the at least one plasma actuator is located on the outer wall. The compression of claim 1, further comprising an outlet guide vane positioned between the first compressor and the transition duct, wherein the outlet guide vane includes the hub portion having a plasma actuator positioned at the hub portion. system. The compression system of claim 2, wherein the plasma actuator is circumferentially continuous about a longitudinal axis. The compression system of claim 1, further comprising a plurality of plasma actuators disposed circumferentially about the longitudinal axis. The compression system of claim 1, wherein the plasma actuator comprises a first electrode and a second electrode separated by a dielectric material. 14. The compression of claim 13, further comprising an AC power supply connected to the first electrode and the second electrode to supply a high voltage AC potential to the first electrode and the second electrode. system. An entrance portion;
An outlet portion located at a distance from the inlet portion at an axial rear;
An axially arcuate inner wall extending between the inlet portion and the outlet portion;
An axially arcuate outer wall extending between the inlet portion and the outlet portion;
An axially arcuate flow path between the inner wall and the outer wall;
Said duct comprising at least one plasma actuator mounted in the duct. The duct of claim 15, wherein the inlet portion has an inlet area and the outlet portion has an outlet area that is larger than the inlet area. The duct of claim 15, wherein the at least one plasma actuator is located on the inner wall. The duct of claim 15, wherein the at least one plasma actuator is located on the outer wall. The duct of claim 15, wherein the plasma actuator comprises a first electrode and a second electrode separated by a dielectric material. The duct of claim 13, further comprising an AC power supply connected to the first electrode and the second electrode to supply a high voltage AC potential to the first electrode and the second electrode. . A gas turbine engine comprising a duct positioned between a first compressor and a second compressor, the duct having at least one plasma actuator mounted therein.

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