Note: Descriptions are shown in the official language in which they were submitted.
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PLASMA ENHANCED COMPRESSOR DUCT
BACKGROUND OF THE INVENTION
This invention relates generally to compressors, and more specifically to a
compression system having a transition duct having plasma actuators.
In a gas turbine engine, air is pressurized in a compression module during
operation.
The air channeled through the compression module is mixed with fuel in a
combustor
and ignited, generating hot combustion gases which flow through turbine stages
that
extract energy therefrom for powering the fan and compressor rotors and
generate
engine thrust to propel an aircraft in flight or to power a load, such as an
electrical
generator.
The compressor includes a rotor assembly and a stator assembly. The rotor
assembly
includes a plurality of rotor blades extending radially outward from a disk.
More
specifically, each rotor blade extends radially between a platform adjacent
the disk, to
a tip. A gas flowpath through the rotor assembly is bound radially inward by
the rotor
blade platforms, and radially outward by a plurality of shrouds.
The stator assembly includes a plurality of circumferentially spaced apart
stator vanes
or airfoils that direct the compressed gas entering the compressor to the
rotor blades.
The stator vanes extend radially between an inner band and an outer band. A
gas
flowpath through the stator assembly is bound radially inward by the inner
bands, and
radially outward by outer bands. The rotor stages comprise rotor blades
arranged
circumferentially around a rotor hub. Each compression stage comprises a vane
stage
and a rotor stage.
Modem high by-pass ratio gas turbine engines have a booster (low pressure
compressor) and a high pressure compressor with a transition duct located in
between.
Conventional transition or gooseneck duct geometries are governed by their
levels of
endwall curvature, since excessive curvature leads to endwall boundary layer
separation and therefore high losses in efficiency. To ensure a smooth
aerodynamic
transition without flow separation, conventional transition duct designs must
have
some minimum axial length for a given change in annular flow radius. This is
not
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desirable because increased transition duct lengths translate directly to
increased
engine length, which in turn adds engine weight and reduces backbone stiffness
of the
engine. This reduction in stiffness makes it more difficult to maintain the
desired
clearances over the rotor tips, reducing the efficiency and operability range
of the
engine.
As compressor and booster rotors approach the limits of their capability to
add
work/pressure to the air, they tend to become less efficient and, if pushed
beyond this
limit, stall (fail to produce their required pressure rise, leading to
reversed flow
through the stage and a loss of engine thrust). A booster rotor that is
designed very
near to its limits in the rear stages of the booster could experience
significant
operability problems. This is a concern in conventional booster system designs
which
are limited to lower radii in the aft rotor stages. These could be corrected
by pushing
the back end of the booster outwards, as enabled by the use of plasma
actuators in the
transition duct.
Accordingly, it is would be desirable to have a shorter transition duct design
having
enhanced pressure distribution without causing flow separation in the duct. It
would
be desirable to have a booster system which has a higher radius for aft rotor
stages
without causing flow separation in the transition duct.
BRIEF DESCRIPTION OF THE INVENTION
The above-mentioned needs may be met by exemplary embodiments which provide a
compression system comprising a first compressor having a first flowpath, a
second
compressor having a second flowpath located axially aft from the first
compressor,
and a transition duct capable of flowing an airfow from the first compressor
to the
second compressor, the transition duct having at least one plasma actuator
mounted in
the transition duct.
In another aspect of the present invention, a duct comprises an inlet portion,
an exit
portion located at a distance axially aft from the inlet portion, an axially
arcuate inner
wall extending between the inlet portion and the exit portion, an axially
arcuate outer
wall extending between the inlet portion and the exit portion, an axially
arcuate
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flowpath between the inner wall and the outer wall, and at least one plasma
actuator
mounted in the duct.
In another aspect of the present invention, a gas turbine engine comprises a
duct
located between a first compressor and a second compressor, the duct having at
least
one plasma actuator mounted in the duct.
BRIEF DESCRIPTION OF THE DRAWINGS
The subject matter which is regarded as the invention is particularly pointed
out and
distinctly claimed in the concluding part of the specification. The invention,
however, may be best understood by reference to the following description
taken in
conjunction with the accompanying drawing figures in which:
Figure 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.
Figure 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.
Figure 3 is a schematic view of a gooseneck duct having plasma actuators
according
an exemplary embodiment of the present invention.
Figure 4 is an enlarged axial cross sectional view of a portion of an
exemplary duct
having a plasma actuator system in the energized mode.
Figure 5 is a schematic view of a gooseneck duct of a booster system according
to an
exemplary embodiment of the present invention.
Figure 6 is a schematic view of a booster system having plasma actuators
according to
an exemplary embodiment of the present invention superimposed with a
conventional
booster flow path for comparison.
Figure 7 is a plot of pressure distributions in a booster system according to
an
exemplary embodiment of the present invention when plasma actuators are
energized
and de-energized.
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DETAILED DESCRIPTION OF THE INVENTION
Referring to the drawings wherein identical reference numerals denote the same
elements throughout the various views, Figure 1 shows a cross-sectional view
of an
exemplary gas turbine engine assembly 10 having a longitudinal axis 11 and a
compression system 20 comprising a first compressor 21 and a second compressor
22
that is located axially aft from the first compressor 21. In the exemplary
embodiment
shown in FIG.1, the first compressor 21 is a booster 40, that is also referred
to
alternatively herein as a low-pressure compressor. The exemplary booster 40
shown
in FIGS. 1 and 2 has four rotor stages, with 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 "IGV") 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 aft 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-flow high-pressure compressor 14
("HPC"). The exemplary HPC 14 shown in FIGS. 1 and 2 has seven rotor stages,
with
each rotor stage having between 24 and 96 HPC rotor blades. The exemplary HPC
14
has a circumferential row of 40 stator vanes (alternatively 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
comprises a combustor 16, and a high-pressure turbine 18 and a low-pressure
turbine
19 that is coupled axially downstream from core gas turbine engine 12, and a
fan
assembly 13 that is coupled axially upstream from core gas turbine engine 12.
Fan
assembly 13 includes an array of fan blades 17 that extend radially outward
from a
rotor disk 29. In the exemplary embodiment shown in FIG. 1, the fan assembly
13, the
booster 40 and low-pressure turbine 19 are coupled together by a first rotor
shaft 28,
and compressor 14 and high-pressure turbine 18 are coupled together by a
second
rotor shaft 27.
In operation, air flows through fan assembly blades 17 and a portion of that
air flows
as bypass airflow 15 and a portion of the air flows as core airflow 25 into
the
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compression system 20 that includes a first compressor 21 and a second
compressor
22. In the exemplary embodiments 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 airflow 25 entering the compression system 20
is
first channeled through a first flow path 23 and is compressed in the first
compressor
21 (shown in the figures herein as booster 40). The core airflow 25 is then
channeled
through an arcuate third flowpath 33 in a duct 30 (alternatively referred to
herein as a
transition duct 30 or as a gooseneck duct 38) to a second flowpath 24 in the
second
compressor 22 (shown in the figures herein as a high-pressure compressor 14)
wherein the core airflow 25 is further compressed. Airflow exiting from the
compression system 20 is channeled to a combustor 16. Air is mixed with fuel
in the
combustor and burned. Products of combustion from combustor 16 are utilized to
drive a high pressure turbine (HPT) 18 and a low pressure turbine (LPT) 19. In
the
exemplary embodiments shown herein, the LPT 19 drives the booster 40 and fan
assembly 13 via fan rotor shaft 28 and the HPT drives the high-pressure
compressor
14 via HP rotor shaft 27. Engine 10 is operable at a range of operating
conditions
between design operating conditions and off-design operating conditions. In
the
exemplary embodiments shown in FIGS. 1 and 2, the booster 40 rotor may have
operating speeds between 1500 rpm and 2700 rpm, and the high-pressure
compressor
14 rotor may have operating speeds between 6000 rpm and 12000 rpm.
In the exemplary embodiments shown in FIGS. 1 and 2, the pitchlines of the
booster
rotor stages are located radially at a higher radius than the pitchlines of
the high-
pressure compressor rotor stages. This is especially true in the case of
modern high
bypass ratio engines. As used herein, "pitchline" of a rotor stage is defined
as an axial
line passing through the radial mid-point between the root and tip of the
leading edge
of the airfoil of a rotor blade in the rotor stage. The transition duct 30
flows the core
airflow 25 from the first flowpath 23 of the booster 40 to the second flowpath
24 of
the high-pressure compressor. FIGS. 3 and 5 show schematically 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. The duct 30 comprises an
inlet
portion 34 and an exit portion 35 that is located axially aft from the inlet
portion. The
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inlet portion 34 has an inlet end 47 having an inlet area 36 and the exit
portion 35 has
an exit end 48 having an exit area 37. The inlet portion 34 is axially located
near the
booster 40 and the exit portion is axially located near the high-pressure
compressor
14. The inlet portion 34 is located radially outward from the exit portion 35
and
centerline axis 11. The duct 38 comprises an inner wall 31 and an outer wall
32 that
form the flowpath 33 in between. The duct 38 may have an annular shape around
the
longitudinal axis 11. In the exemplary embodiments shown in FIGS. 1, 2 and 3,
struts
46 of a support frame extend radially through the third flowpath 33 of the
duct 38 at
some circumferential locations. The third flow path has a generally annular
shape
with respect to the longitudinal axis 11 in the axial direction, with the
struts 46
extending through it at certain circumferential locations in some
applications. Due to
the generally annular configuration of the duct 38 with the inlet portion 34
located
radially outward from the exit portion 35, the third flow path 33 and the duct
38 have
a gooseneck shape, such as shown, for example, in FIGS. 3 and 5. The inner
wall 31
and the outer wall 32 have an arcuate shape in the axial direction, such as
shown, for
example, in FIGS. 3 and 5.
Referring to FIG. 5, the inlet end 47 of the duct 30 is located at a higher
radius with
respect to the longitudinal axis 11 than the exit end 48. The exit end 48 is
located at
an axial aft distance 76 ("D") from the inlet end 47. In the exemplary
embodiments of
the present invention shown herein, the ratio of the inlet outer radius 71
("RI") to the
exit outer radius 72 ("RO") is about 1.8. For the same duct axial length, this
ratio is
about 1.6 or less for conventional designs. In the exemplary embodiments shown
herein, the axial distance D 76 is between about 16 inches and 18 inches. In
the
exemplary embodiments shown herein, the inlet area 36 is about 598 sq. inches
and
the exit area 37 is about 570 sq. inches. A slight reduction in the exit area
37 from the
inlet area 36 may help to further reduce flow separation in the duct 38. In
other
embodiments of the present invention, the inlet area 36 and the exit area 37
may have
other suitable values. In alternative embodiments, it may be advantageous to
have the
exit area 37 larger than the inlet area 36 to improve pressure distributions
in the duct
38 using known design methods. The present invention enables the design of
booster
systems having short duct axial lengths ("D") as compared to the inlet and
exit radii
("RI" and "RO"). In the exemplary embodiments of the present invention, the
aspect
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ratio, defined as the ratio (RI-RO)/D, is between about 0.5 and 0.8. Due to
the
geometric nature of the cross sectional shape of the third flowpath 33, such
as shown
in FIGS. 3 and 5, the transition duct 30 is alternatively referred to herein
as a
gooseneck duct 38.
As is evident from the exemplary embodiments shown herein, the inner wall 31
and
outer wall 32 have significant curvatures in the axial direction. In the
exemplary
embodiments of the present invention shown in FIGS. 1, 2, 3, 5 and 6 flow
separation
in the duct 38 is reduced by using plasma actuators 60. The terms "plasma
actuator"
and "plasma generator" as used herein have the same meaning and are used
interchangeably. The plasma actuators, such as for example, shown as items 60,
61,
62, and 63 in the figures herein, strengthen the local axial momentum of the
airflow
near the walls 31, 32 and minimize flow separation in the duct 30 in regions
having
sharp radius of curvature in the inner and outer walls 31, 32. Plasma
actuators used as
shown in the exemplary embodiments of the present invention, produce a stream
of
ions and a body force that act upon the fluid near the walls 31, 32, forcing
it to flow
closer to the walls 31, 32 in direction of the desired fluid flow with reduced
flow
separation from the walls 31, 32.
FIG. 4 schematically illustrates, in axial cross-section view, an exemplary
embodiment of plasma actuator 60 for reducing the flow separation in a
transition
duct 38 located between two compressors, such as the booster 40 and the HPC 14
shown in FIGS.1 and 2. The exemplary embodiments of the present invention
shown
herein facilitate an improvement of the pressure distribution in the duct 38
(see FIG.
7) and/or enhance the efficiency of compression systems, in a gas turbine
engine 10
such as the aircraft gas turbine engine illustrated in cross-section in FIG.
1. The
exemplary gas turbine engine plasma actuators shown in FIGS. 1-6 include
plasma
actuators, such as shown as items 60, 61, 62 or 63 located on the inner wall
31, outer
wall 32 or the hub portion 45 of the booster OGV 44. The 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
an annular groove. Alternatively, the plasma actuator 60 may be segmented
wherein a
plurality of plasma actuators 60 are located in corresponding groove segments
spaced
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circumferentially in the walls 31, 32. The exemplary embodiment shown in
Figure 4
comprises a plasma actuator 60 located in a groove 68 in the inner wall 31 of
the duct
38. Alternately, the plasma actuators 60 may be located at other locations in
the duct
38 where flow separation is likely to occur, such as, for example, locations
where the
duct 38 walls have a sharp radius of curvature in the direction of airflow.
The exemplary embodiment shown in FIG. 4 shows an annular plasma generator 60
mounted 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 within
an 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 a
range of
about 3-20 kV to the electrodes 64, 66. When the AC amplitude is large enough,
the
air ionizes in a region of largest electric potential forming a plasma 80. The
plasma
80 generally begins near an edge 67 of the first electrode 64 which is exposed
to the
air and spreads out over an area 69 projected by the second electrode 66 which
is
covered by the dielectric material 65. The plasma 80 (ionized air) in the
presence of
an electric field gradient produces a force on the airflow 25 near the wall 31
inducing
a virtual aerodynamic shape that causes a change in the pressure distribution
over the
inner wall 31 of the annular duct 38. The air near the electrodes is weakly
ionized,
and usually there is little or no heating of the air. The airflow 25 near the
wall 31
tends to remain attached to the wall 31 resulting in reduced flow separation
and
improved pressure distribution within the duct 38 due to reduced pressure loss
in the
duct 38.
FIG. 6 shows a booster system 50 according to an exemplary embodiment of the
present invention. The booster system 50 shown in FIG. 6 has a last rotor
stage 57
having a pitchline radius 54 that is larger than conventional booster systems.
This is
made possible in the present invention by the use of plasma actuators, such
as, for
example, shown as items 60, 61, 62 in FIG. 6, in a duct 38 that receives the
flow from
the last stage of the booster. A conventional flowpath 90 of a conventional
booster
system is shown by dotted line in FIG. 6 for comparison with the exemplary
embodiment of the present invention, booster system 50. There are several
benefits
associated with having the aft stages of the booster, such as the last rotor
stage 57,
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radially further outward. The rotor stage 57, having a larger pitchline radius
54, has
an increased tip speed compared to conventional designs. Since the ability of
a rotor
to do work on a fluid is directly related to its tangential velocity, the
exemplary
embodiment of the present invention shown in FIG. 6 has increased capacity to
produce pressure rise. In some applications, for a desired pressure ratio, it
is possible
to reduce the number of required stages in a booster system by using the
present
invention resulting in significantly reduced weight for the engine 10.
In the exemplary embodiment of the present invention shown in FIG.6, the
booster
system 50 has a first rotor stage 55 comprising a plurality of first rotor
blades 56
spaced circumferentially around a rotor hub 41 and having a first pitch-line
radius 53
extending from the longitudinal axis 11, a last rotor stage 57 located axially
aft from
the first rotor stage 55. 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 aft from the last rotor stage 57 and receives the airflow
25 exiting
from the last rotor stage 57. The gooseneck duct 38 has an inlet end 47, an
exit end 48
located at a distance axially aft from the inlet end 47, and has at least one
plasma
actuator mounted in the gooseneck duct 38. The geometry of the gooseneck duct
38
and the placement of the plasma actuators, such as for example, shown as items
60,
61, 62 in FIG. 6, are described previously herein. Unlike conventional booster
systems, the last rotor stage 57 has a higher pitchline radius 54 "B" as
compared to
the first rotor stage 55 pitchline radius 53 "A". In the exemplary embodiments
of the
present 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 aft end, the duct 38 having an axially arcuate inner wall 31 and an
axially arcuate
outer wall 32. The exit end 48 has an exit area 37 and the inlet end 47 has an
inlet
area 36. The geometry of the gooseneck duct (see FIG. 5) is such that the
ratio RI/RO
of the inlet outer radius 71 to the exit outer radius 72 is at least 1.6.
Plasma actuators,
such as for example, shown as items 60, 61, 62 in FIG. 6 are located in the
duct 38 as
described previously herein.
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A gas turbine engine 10 having a booster system 50 with the gooseneck duct 38
having plasma actuators as described herein, can be operated by energizing the
first
electrode 64 and second electrode 66 using the AC potential from the AC power
supply 70. By energizing the electrodes 64, 66 and creating the plasma 80,
flow
separation in the duct 38 can be reduced which results in the advantages and
improvements in pressure distributions in the booster system 50. In one
method, the
plasma actuators, such as item 60 in FIG.6, can be energized continuously
throughout
engine operation period. Alternatively, the plasma actuators can be energized
only
during selected portions of the engine operating regime. The periods and
durations of
plasma actuator energization can be determined by known engine test methods
for
determining engine operability.
FIG. 7 shows an exemplary pressure distribution within the duct 38 at the exit
end 48
determined by known fluid flow analytical methods. The horizontal axis shows
the
normalized pressure and the vertical axis shows the radial span locations
within the
duct 38. The distribution identified by numeral 91 shows the radial pressure
distribution at the exit end 48 of the duct 38 when the plasma actuator 60 is
not
energized by the AC power supply 70. The distribution identified by numeral 92
shows the radial pressure distribution at the same location (exit end 48 of
the duct 38)
when the plasma actuator 60 is energized by the AC power supply 70. It is
clear that
near the wall 32 (near the 1.0 span location) wherein the plasma actuator is
located,
the normalized pressure increases from about 0.79 to about 0.86.
As used herein, an element or step recited in the singular and proceeded with
the word
"a" or "an" should be understood as not excluding plural said elements or
steps, unless
such exclusion is explicitly recited. When introducing
elements/components/steps etc.
of designing and/or manufacturing components and systems described and/or
illustrated herein, the articles "a", "an", "the" and "said" are intended to
mean that
there are one or more of the element(s)/component(s)/etc. The terms
"comprising",
"including" and "having" are intended to be inclusive and mean that there may
be
additional element(s)/component(s)/etc. other than the listed
element(s)/component(s)/etc. Furthermore, references to "one embodiment" of
the
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present invention are not intended to be interpreted as excluding the
existence of
additional embodiments that also incorporate the recited features.
Although the methods and articles such as vanes, outer bands, inner bands and
vane
segments described herein are described in the context of a compressor used in
a
turbine engine, it is understood that the vanes and vane segments and methods
of their
manufacture or repair described herein 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, these
can be
utilized independently and separately from other components described herein.
This written description 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
intended to
be within the scope of the claims if they have structural elements that do not
differ
from the literal language of the claims, or if they include equivalent
structural
elements with insubstantial differences from the literal languages of the
claims.
While there have been described herein what are considered to be preferred and
exemplary embodiments of the present invention, other modifications of the
invention
shall be apparent to those skilled in the art from the teachings herein, and
it is,
therefore, desired to be secured in the appended claims all such modifications
as fall
within the true spirit and scope of the invention.
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