U.S. patent application number 13/015786 was filed with the patent office on 2012-08-02 for plasma actuation systems to produce swirling flows.
This patent application is currently assigned to GENERAL ELECTRIC COMPANY. Invention is credited to Grover Bennett, Katherine Essenhigh, Giridhar Jothiprasad, Robert Murray, Seyed Gholamali Saddoughi, Aspi Wadia.
Application Number | 20120195736 13/015786 |
Document ID | / |
Family ID | 45524412 |
Filed Date | 2012-08-02 |
United States Patent
Application |
20120195736 |
Kind Code |
A1 |
Jothiprasad; Giridhar ; et
al. |
August 2, 2012 |
Plasma Actuation Systems to Produce Swirling Flows
Abstract
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.
Inventors: |
Jothiprasad; Giridhar;
(Niskayuna, NY) ; Saddoughi; Seyed Gholamali;
(Clifton Park, NY) ; Bennett; Grover;
(Schenectady, NY) ; Murray; Robert; (Rotterdam,
NY) ; Wadia; Aspi; (Loveland, OH) ; Essenhigh;
Katherine; (Niskayuna, NY) |
Assignee: |
GENERAL ELECTRIC COMPANY
Schnectady
NY
|
Family ID: |
45524412 |
Appl. No.: |
13/015786 |
Filed: |
January 28, 2011 |
Current U.S.
Class: |
415/1 ;
415/148 |
Current CPC
Class: |
F01D 17/10 20130101;
F01D 11/08 20130101; F05D 2270/172 20130101; F01D 11/10 20130101;
F05D 2270/101 20130101 |
Class at
Publication: |
415/1 ;
415/148 |
International
Class: |
F04D 27/02 20060101
F04D027/02; F04D 29/56 20060101 F04D029/56 |
Claims
1. A plasma actuation system for a turbo-machine, device,
comprising: an end wall; a plurality of end wall actuators
positioned about the end wall; and a blade positioned adjacent to
the end wall; wherein the plurality of end wall actuators are
oriented to produce a swirling flow between the end wall and the
blade.
2. The plasma actuation system of claim 1, wherein the plurality of
end wall actuators comprises a plurality of circumferential
momentum end wall actuators.
3. The plasma actuation system of claim 2, wherein one or more of
the plurality of circumferential momentum end wall actuators are
positioned about the blade.
4. The plasma actuation system of claim 1, wherein the plurality of
end wall actuators comprises a plurality of intermediate momentum
end wall actuators.
5. The plasma actuation system of claim 1, further comprising a
plurality of blade actuators positioned on the blade.
6. The plasma actuation system of claim 5, wherein the plurality of
blade actuators comprises a plurality of circumferential momentum
blade actuators.
7. The plasma actuation system of claim 5, wherein the plurality of
blade actuators comprises a plurality of intermediate momentum
blade actuators.
8. The plasma actuation system of claim 5, wherein the plurality of
blade actuators are positioned about a tip of the blade.
9. The plasma actuation system of claim 5, wherein the plurality of
end wall actuators and the plurality of blade actuators comprise a
plurality of axial momentum actuators.
10. The plasma actuation system of claim 5, wherein the plurality
of end wall actuators and the plurality of blade actuators comprise
a plurality of dielectric barrier discharge plasma actuators.
11. The plasma actuation system of claim 10, wherein each of the
plurality of dielectric barrier discharge plasma actuators
comprises a pair of conductive layers to produce a plasma
therebetween.
12. The plasma actuation system of claim 1, wherein the
turbo-machinery device comprises a compressor.
13. The plasma actuation system of claim 1, wherein the
turbo-machinery device comprises a turbine.
14. A method of reducing a blockage and losses about an end wall
and a blade tip of a turbo-machinery device, comprising: actuating
a plurality 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.
15. The method of claim 14, further comprising the step of
actuating a plurality of blade actuators generating circumferential
and/or intermediate momentum in the flow therethrough.
16. The method of claim 15, further comprising the step of
generating axial momentum in the flow therethrough.
17. A plasma actuation system for a turbo-machinery device;
comprising: an end wall; the end wall comprising a plurality of
circumferential momentum end wall actuators and/or a plurality of
intermediate momentum end wall actuators; and a blade; the blade
comprising a plurality of circumferential momentum blade actuators
and/or a plurality of intermediate momentum blade actuators;
wherein the plurality of circumferential momentum end wall
actuators, the plurality of intermediate momentum end wall
actuators, the plurality of circumferential momentum blade
actuators, and the plurality of intermediate momentum blade
actuators are oriented to produce a swirling flow between the end
wall and the blade.
18. The plasma actuation system of claim 17, wherein the plurality
of circumferential momentum end wall actuators, the plurality of
intermediate momentum end wall actuators, the plurality of
circumferential momentum blade actuators, and the plurality of
intermediate momentum blade actuators comprise a plurality of
dielectric barrier discharge plasma actuators.
19. The plasma actuation system of claim 17, wherein the
turbo-machinery device comprises a compressor.
20. The plasma actuation system of claim 17, wherein the
turbo-machinery device comprises a turbine.
Description
TECHNICAL FIELD
[0001] 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.
BACKGROUND OF THE INVENTION
[0002] 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.
[0003] 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.
[0004] 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.
[0005] 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.
SUMMARY OF THE INVENTION
[0006] 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.
[0007] 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.
[0008] 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.
[0009] 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.
BRIEF DESCRIPTION OF DRAWINGS
[0010] FIG. 1 is a schematic view of a known gas turbine
engine.
[0011] FIG. 2 is a partial cross-sectional view of a
turbo-machinery device showing a blade tip and an end wall with a
flow path therethrough.
[0012] FIG. 3 is a schematic view of a portion of a turbo-machinery
device with a plasma actuation system as may be described
herein.
[0013] FIG. 4 is schematic view of a dielectric barrier discharge
plasma actuator as may be used in the plasma actuation system of
FIG. 3.
[0014] FIG. 5 a perspective view of a turbo-machinery device with a
portion of the plasma actuation system of FIG. 3.
[0015] FIG. 6 is a schematic view of the plasma actuation system of
FIG. 3 with the direction of the plasma force shown.
[0016] FIG. 7 a perspective view of a turbo-machinery device with a
portion of the plasma actuation system of FIG. 3.
[0017] FIG. 8 is a schematic view of the plasma actuation system of
FIG. 3 with the direction of the plasma force shown.
[0018] FIG. 9 is a schematic view of the plasma actuation system of
FIG. 3 with the direction of the plasma force shown.
[0019] FIG. 10 a perspective view of a turbo-machinery device with
a portion of the plasma actuation system of FIG. 3.
DETAILED DESCRIPTION
[0020] Referring now to the drawings, in which like numerals refer
to like elements throughout the several views, FIG. 1 shows a
schematic view of a rotary machine such as gas turbine engine 10.
The gas turbine engine 10 may include a compressor 15. The
compressor 15 compresses an incoming flow of air 20. The compressor
15 delivers the compressed flow of air 20 to a combustor 25. The
combustor 25 mixes the compressed flow of air 20 with a compressed
flow of fuel 30 and ignites the mixture to create a flow of
combustion gases 35. Although only a single combustor 25 is shown,
the gas turbine engine 10 may include any number of combustors 25.
The flow of combustion gases 35 is delivered in turn to a turbine
40. The flow of combustion gases 35 drives the turbine 40 so as to
produce mechanical work. The mechanical work produced in the
turbine 40 drives the compressor 15 and also may drive an external
load 45 such as an electrical generator and the like.
[0021] 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.
[0022] 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.
[0023] As is shown in FIG. 2, a tip clearance flow 75 may be driven
therethrough by a pressure difference across the blade 50 (blade
loading). The interaction of the clearance flow 75 with an incoming
main flow 80 creates a region of low-speed fluid and high losses.
These large clearance flow losses allow an interface 85 to be
formed therein. The interface 85 may be defined as a region of high
entropy gradient. The low-speed fluid region enclosed by the
interface 85 thus acts as a blockage 90 to the main flow 80 and
increases the blade loading near the tip 70. Increases in this
blockage 90 may be a precursor to stall events such as those
described above.
[0024] FIG. 3 shows one example of a plasma actuation system 100 as
may be described herein. The plasma actuation system 100 may be
used with a turbo-machinery device 105 such as the compressor 25
and/or the turbine 40. The plasma actuation system 100 may include
a number of end wall actuators 110 positioned about the end wall
60. The end wall actuators 110, in turn, may include a number of
axial momentum end wall actuators 120, a number of circumferential
momentum end wall actuators 130, and a number of intermediate
momentum end wall actuators 135. Moreover, one or more of the
blades 50 also may have a number of blade actuators 140 positioned
thereabout. The blade actuators 140 may include one or more axial
momentum blade actuators 150, one or more circumferential momentum
blade actuators 160, and one or more intermediate momentum blade
actuators. Other types of plasma actuators 110 may be used herein
in other orientations and in other locations. Not all of the
actuators 110 must be used in any given application. Any number of
plasma actuators 110 may be used herein.
[0025] FIG. 4 shows an example of dielectric barrier discharge
plasma actuator 170 as may be used as any of the actuators
described above. The actuator 170 may include a conductive or a
non-conductive substrate 180. A dielectric layer 190 may be
positioned thereon. A first thin conductive layer 200 may be
deposited on the non-conductive substrate 180 with the dielectric
layer 190 on top. If the substrate 180 is conductive, the substrate
itself acts as the first thin conductive layer 200. A second thin
conductive layer 210 then may be disposed on the dielectric layer
190. The conductive layers 200, 210 may be connected to a power
source 220 and a wave-form controller 230. The wave form controller
230 may be configured to control an input voltage level and
pulsing, variable or AC voltage frequency, duty cycle and shape,
and the like. Other types of actuators 170 also may be used herein
such as single dielectric barrier discharge actuators, surface
corona discharge actuators, and the like. Other components and
other configurations may be used herein.
[0026] 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.
[0027] FIG. 5 shows an example of the axial momentum end wall
actuator 120. As is shown, the actuators 120 may be positioned
about the end wall 60 and face the blades 50 about the tip
clearance space 65. A number of electrodes 260 may be in
communication with each actuator 120. Each actuator 120 may extend
the length of several blades 50. FIG. 6 shows the plasma 240 with
the arrows 270 indicating the direction of the plasma force
extending perpendicularly to a direction 280 of the blade rotation
so as to increase the axial momentum of the flow therethrough.
[0028] FIG. 7 shows an example of the circumferential momentum end
wall actuators 130. Likewise, the actuators 130 may be positioned
about the end wall 60. In this example, one or more actuators 130
may be used for each blade 50. FIG. 8 shows the plasma 240 with the
arrows 270 indicating the direction of the plasma force running
parallel and in the same direction 280 as the blade direction. FIG.
9 shows the force of the plasma 240 running parallel but opposite
of the direction 280 of the blade rotation (counter-swirl). Either
direction acts to alter the circumferential momentum of the flow
therethrough.
[0029] 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.
[0030] FIG. 10 shows an example of the blade actuators 140. In this
example, a number of axial momentum blade actuators 150, a number
of the circumferential momentum blade actuators 160, and a number
of intermediate momentum blade actuators 165 may be used at the tip
70. The arrows 270 show the different directions of the force of
the plasma 240 so as to alter axial, circumferential, and/or
intermediate momentum to the flow therethrough. Any number of the
actuators 150, 160, 165 may be used on a given blade 50 in any
orientation. Other components and configurations also may be used
herein.
[0031] 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.
[0032] 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.
[0033] 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.
[0034] 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.
* * * * *