Patent Application
20120195736
The present application relates generally to gas turbine engines and more particularly relates to plasma actuation systems that produce swirling flows at the end walls of turbo-machinery and the like so as to reduce to end wall blockages and losses therein.
Aerodynamic instabilities such as rotating stall and surge impose fundamental limits on the stability of compressors. Rotating stall may occur as the mass flow through the compressor is decreased at a certain speed. Stall cells may be created and may rotate around the circumference of the compressor as opposed to moving in the axial flow direction. Such stall cells may reduce substantially the efficiency of the compressor and also may increase the structural load on the airfoils in the localized region. Compressor surge may result in the reversal of the flow through the compressor and the expulsion of the previously compressed air. Compressor surge may result when the compressor does not have the capacity to absorb momentary disturbances. Recovery from compressor surge typically involves a complete restart of the engine.
Compressors thus are generally designed with a safety margin or a stall margin against rotating stall and the like. Current compressor designs, however, may increase the tip clearance to blade height ratio and thus may result in a significant decrease in the stall margin. Known approaches to stall margin improvement, however, such as casing treatments, oscillating inlet guide vanes, rotor tip injections, and the like, may have an impact on the efficiency of the compressor and may result in significant penalties in terms of weight or the use of “expensive” high pressure air from downstream stages.
For a turbine, the clearance gap between the end walls and the blades may be a significant source of typical aerodynamic losses. The clearance flows also interact strongly with other secondary flows present in the blade passage. As a result, losses due to clearance flows may account for nearly a third of the total losses of the turbine.
There is thus a desire for improved compressor designs and/or flow control systems so as to provide a robust stall margin even with the use of smaller blade heights. By avoiding known aerodynamic instabilities such as those described above, compressor designs may have increased safety throughout a mission, increased tolerance for stage mismatch during transient operations, and the opportunity to match stages at maximum efficiency so as to reduce the fuel burn therethrough while maintaining high efficiency. Likewise, there is strong need to develop flow control devices that can mitigate losses due to clearance flows in a turbine.
The present application provides a plasma actuation system for a turbo-machinery device. The plasma actuation system may include an end wall, a number of end wall actuators positioned about the end wall, and a blade positioned adjacent to the end wall. The end wall actuators are oriented to produce a swirling flow between the end wall and the blade.
The present application further provides a method of reducing a blockage and losses about an end wall and a blade tip of a turbo-machinery device. The method may include the steps of actuating a number of end wall actuators, generating circumferential and/or intermediate momentum in a flow therethrough, and creating a swirling flow near the end wall and the blade tip so as to reduce the blockage and losses thereabout.
The present application further provides a plasma actuation system for a turbo-machinery device. The plasma actuation system may include an end wall with a number of circumferential momentum end wall actuators and/or a number of intermediate momentum end wall actuators and a blade with a number of circumferential momentum blade actuators and/or a number of intermediate momentum blade actuators. The circumferential momentum end wall actuators, the intermediate momentum end wall actuators, the circumferential momentum blade actuators, and the intermediate momentum blade actuators are oriented to produce a swirling flow between the end wall and the blade.
These and other features and improvements of the present application will become apparent to one of ordinary skill in the art upon review of the following detailed description when taken in conjunction with the several drawings and the appended claims.
Referring now to the drawings, in which like numerals refer to like elements throughout the several views,
The gas turbine engine 10 may be one of any number of different gas turbine engines offered by General Electric Company of Schenectady, New York and the like. The gas turbine engine 10 may have other configurations and may use other types of components. Other types of gas turbine engines also may be used herein. Multiple gas turbine engines 10, other types of turbines, and other types of power generation and propulsion equipment also may be used herein together. Other types of rotary machines also may be used herein.
Generally described, the compressor 15 and the turbine 40 include a number of circumferentially spaced blades 50 positioned on a shaft 55 for rotation therewith. The blades 50 may be positioned within an end wall 60. The end wall 60 may be a casing or any type of other type of structure. A tip clearance space 65 may exist between the end wall 60 and a tip 70 of the blade 50. The blades 50 may rotate while the end wall 60 is stationary. Likewise, the blade 50 may be in the form of a stationary stator and the end wall 60 may be positioned on a rotating shaft thereabout.
As is shown in
In use, an air flow located above the dielectric layer 190 and between the conductive layers 200, 210 may be ionized in a desired fashion to create a region of a discharge plasma 240. The actuator 170 thus may be oriented to impart momentum to a flow therethrough via the discharge plasma 240. In this example, multiple actuators 170 in different orientations may be used to create a swirling flow 250 from the tip clearance flow 75 and the incoming flow 80 with momentum injection as will be described in more detail below.
Likewise, the intermediate momentum end wall actuators 130 may generate the plasma 240 with force extending in any desired direction between axial and circumferential. The intermediate momentum end wall actuators 135 may alter the intermediate momentum of the flow therethrough.
The combination of the different actuators 170 within the plasma actuation system 100 thus may be used to generate the swirling flows 250 about the tip 70 and the end wall 60 so as to reduce the blockage 90 and other losses near the tip 70. Specifically, the actuators 170 alter the axial, the circumferential momentum, and/or the intermediate momentum of the flows therethrough to create the swirling flow 250. Hence, the plasma actuation system 100 may inject an optimal combination of axial, circumferential, and/or intermediate momentum into the tip gap flows. Energizing the clearance flow by injection of momentum in optimal directions and locations thus reduces the losses and blockage introduced by the interaction of the clearance flow with the main flow.
The location of the actuators 170 may be chosen based on a specific turbo machinery design so as to reduce the blockage 90 and losses in and about the tip/end wall region. The actuators 170 also may be excited at different forcing frequencies so as to minimize the losses and blockages introduced in and about the tip/end wall region. For example, the blade passing frequencies and variations thereon may be used. The actuators 170 also have the relatively fast response time so as to enable active feedback control. Multiple actuators 170 may be used in series to augment the force imparted to the flow 250.
The appropriate injection of momentum by the actuators 170 may energize end wall boundary layers so as to minimize end wall boundary layer separation, reduce blade loading at the tip, and minimize blockage and losses. The swirling flows 250 produced by the actuators 170 thus may improve the aerodynamic performance stability characteristics of the overall turbo-machinery device 105. Such increased stability may lead to increased safety throughout the mission, increased tolerances for stage mismatch during part speed operation and transients, and an opportunity to match stages at the compressor maximum efficiency point so as to reduce fuel burn. Moreover, the actuators 170 do not use the “expensive” compressed air from upstream stages. Reduction in tip clearance flows also may lead to reduced fuel burn.
It should be apparent that the foregoing relates only to certain embodiments of the present application and that numerous changes and modifications may be made herein by one of ordinary skill in the art without departing from the general spirit and scope of the invention as defined by the following claims and the equivalents thereof.
