Note: Descriptions are shown in the official language in which they were submitted.
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GAS TURBINE VANE WITH A COOLED INNER SHROUD
BACKGROUND OF THE INVENTION
The present invention relates to a stationary
vane for use in the turbine section of a gas turbine. More
specifically, the present invention relates to the cooling
of the inner shroud portion of a gas turbine vane.
A gas turbine employs a plurality of stationary
vanes that are circumferentially arranged in rows in a
turbine section. Each vane is comprised of an airfoil
section formed between inner and outer shrouds. Since such
vanes are exposed to the hot gas discharging from the
combustion section, cooling of these vanes is of utmost
importance. Typically, cooling is accomplished by flowing
cooling air through radially oriented passages, such as
forward and aft passages, formed inside the vane airfoil.
A portion of the cooling air flowing through the
aft airfoil passage is typically discharged through cooling
air holes in the trailing edge of the airfoil. Another
portion of the cooling air flowing through the aft passage,
as well as the cooling air flowing through the forward
airfoil passage, is typically discharged from the vane
through the inner shroud and enters a cavity located
between adjacent rows of rotor discs. The cooling air in
the cavity serves to cool the faces of the discs.
In the past, a portion of the cooling from the
cavity between the discs has sometimes been used to cool
the inner shroud by impinging cooling air against the
shroud surface or flowing cooling air through passages in
the body of the shroud. Unfortunately, traditional schemes
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have not made optimum use of this cooling air. Although such cooling air
eventually
enters the hot gas flowing through the turbine section, little useful work is
obtained from
the cooling air, since it was not subject to heat up in the combustion
section. Thus, to
achieve high efficiency, it is crucial that the cooling air be effectively
utilized so as to
minimize the amount of cooling air used.
In FR-A-2198054 there is disclosed a gas turbine which has a chamber
formed therein and a first passage in flow communication with the chamber. The
chamber receives at least a portion of cooling air received via the first
passage.
Additionally, there are a plurality of second passages in flow communication
with the
chamber such that the second passages receive at least a portion of the
cooling air
received by the chamber. Furthermore, a shroud portion has a first passage
formed
therein which is in flow communication with a cavity and an array of pin fins.
GB-A-2104965 sliows a shroud for a vane having the above features and
in addition has a first surface exposed to the flow of hot compressed gas and
a second
surface which is disposed opposite the first surface. Furthermore there is a
chamber and
an outlet of the first passage is connected to the chamber and the inlet of
the second
passage is also connected to the chamber.
It is therefore desirable to provide a scheme for efficiently cooling the
inner shroud of a
gas turbine vane_
SUMMARY OF THE INVENTION
Accordingly, it is the general object of the current invention to provide a
scheme for
efficiently cooling the inner shroud of a gas turbine vane.
Briefly, this object, as well as other objects of the current invention, is
accomplished in a
gas turbine compnsing (i) a compressor for producing compressed air, (ii) a
combustor
for heating at least a portion of the compressed air, thereby producing a hot
compressed
gas, and (iii) a turbine for expanding the hot compressed gas so as to produce
shaft
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power. The turbine has a stationary vane disposed therein that is exposed to
the hot
compressed gas and a cavity in flow communication with the compressor, whereby
the
cavity receives a flow of cooling air formed by a portion of the compressed
air. The
stationary vane has an airfoil portion and a shroud portion disposed adjacent
to the
cavity. The shroud portion has a first passage in flow communication with the
cavity,
whereby the first passage receives a portion of the cooling air. An array of
pin fins are
disposed in the first passage.
According to one embodiment of the invention, the inner shroud further
comprises (i) a
first surface exposed to the flow of hot compressed gas and a second surface
disposed
opposite the first surface, and (ii) means for impinging cooling air against
the second
surface. The means for impinging cooling air comprises a plate attached to the
inner
shroud, the plate having a plurality of holes.
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formed therein. In this embodiment, the holes in the plate
are in flow communication with the cavity, whereby the
second portion of the cooling air flows through the holes
and forms the cooling air that impinges against the shroud
surf ace .
BRIEF DESCRIPTION OF THE DRAWINGS
Figure 1 is a longitudinal cross-section,
partially schematic, of a gas turbine incorporating the
vane of the current invention.
Figure 2 is a view of the underside -- that is,
looking radially outward, of the inner shroud portion of
the row 2 vane shown in Figure 1.
Figure 3 is a cross-section taken through line
III-III shown in Figure 2, showing a detailed view of the
portion of Figure 1 in the vicinity of the row 2 vane.
Figure 4 is a cross-section taken through line
IV-IV shown in Figure 2, showing the inlet to the pin fin
passage.
Figure 5 is a cross-section taken through line V-
V shown in Figure 2, showing the impingement chamber and
inner shroud trailing edge cooling hole.
Figure 6 is a cross-section taken through line
VI-VI shown in Figure 2, showing both impingement chambers.
Figure 7 is a cross-section taken through line
VII-VII shown in Figure 2, showing the inlet and outlet to
the pin fin passage.
Figure 8 is a cross-section taken through line
VIII-VIII shown in Figure 2, showing the cooling air hole
in the side rail of the inner shroud.
It should be noted that cross-sections taken
through Figure 2 would show the vane oriented upside down
from the manner in which it is normally viewed. Therefore,
to allow ready comprehension, Figures 3-8 have been rotated
so that the vane is oriented in its normal upright
position.
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DESCRIPTION OF THE PREFERRED EMBODIMENT
Referring to the drawings, there is shown in
Figure 1 a longitudinal cross-section through a portion of
a gas turbine. The major components of the gas turbine are
a compressor section 1, a combustion section 2, and a
turbine section 3. As can be seen, a rotor 4 is centrally
disposed and extends through the three sections. The
compressor section 1 is comprised of cylinders 7 and 8 that
enclose alternating rows of stationary vanes 12 and
rotating blades 13. The stationary vanes 12 are affixed to
the cylinder 8 and the rotating blades 13 are affixed to
discs attached to the rotor 4.
The combustion section 2 is comprised of an
approximately cylindrical shell 9 that forms a chamber 14,
together with the aft end of the cylinder 8 and a housing 6
that encircles a portion of the rotor 4. A plurality of
combustors 15 and ducts 16 are contained within the chamber
14. The ducts 16 connect the combustors 15 to the turbine
section 3. Fuel 35, which may be in liquid or gaseous form
-- such as distillate oil or natural gas -- enters each
combustor 15 through a fuel nozzle 34 and is burned therein
so as to form a hot compressed gas 30.
The turbine section 3 is comprised of an outer
cylinder 10 that encloses an inner cylinder 11. The inner
cylinder 11 encloses rows of stationary vanes and rows of
rotating blades that are circumferentially arranged around
the centerline of the-rotor 4. The stationary vanes are
affixed to the inner cylinder 11 and the rotating blades
are affixed to discs that form a portion of the turbine
section of the rotor 4.
In operation, the compressor section 1 inducts
ambient air and compresses it. A portion of the air that
enters the compressor is bled off after it has been
partially compressed and is used to cool the rows 2-4
stationary vanes within the turbine section 3, as discussed
more fully below. The remainder of the compressed air 20
is discharged from the compressor section 1 and enters the
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chamber 14. A portion of the compressed air 20 is drawn
from the chamber 14 and used to cool the first row of
stationary vanes, as well as the rotor 4 and the rotating
blades attached to the rotor. The remainder of the
5 compressed air 20 in the chamber 14 is distributed to each
of the combustors 15.
In the combustors 15, the fuel 35 is mixed with
the compressed air and burned, thereby forming the hot
compressed gas 30. The hot compressed gas 30 flows through
the ducts 16 and then through the rows of stationary vanes
and rotating blades in the turbine section 3, wherein the
gas expands and generates power that drives the rotor 4.
The expanded gas 31 is then exhausted from the turbine 3.
The current invention is directed to the cooling
of the stationary vanes and will be discussed in detail
with reference to the second row of stationary vanes 17.
As shown in Figure 1, a portion 19 of the air flowing
through the compressor 1 is extracted from an interstage
bleed manifold 21, via a pipe 32, and is directed to the
turbine section 3. In the turbine section 3, the cooling
air 19 enters a manifold 22 formed between the inner .
cylinder 11 and the outer cylinder 10. From the manifold
22, the cooling air 19 enters the second row vanes 17.
As shown in Figure 3, the vane 17 is comprised of
an airfoil portion 25 that is disposed between inner and
outer shrouds 26 and 27, respectively. Support rails 36
and 37 formed on the outer shroud 27 are used to attach the
vane 17 to the turbine inner cylinder 11. As shown in
Figure 6, the airfoil portion 25 of the vane 17 has
generally concave shaped wall 51, which forms the pressure
surface 23 of the airfoil, and a generally convex wall 52,
which forms the suction surface 24 of the airfoil. At
their upstream and downstream ends, the walls 51 and 52
form the leading and trailing edges 28 and 29,
respectively, of the airfoil 25. The airfoil 25 is
substantially hollow and a rib 40 divides the interior into
fore and aft passages 42 and 44, respectively.
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Tubular members 46 and 47 -- referred to as
"inserts" -- are attached to the outer shroud 27 and extend
into the fore and aft cavities 42 and 44, respectively. A
number of small cooling air holes 70 and 71 are formed in
the inserts 46 and 47. The cooling air holes 70 and 71
serve to impinge cooling air on the airfoil walls 51 and 52
and to distribute portions of the cooling air 19' and 19"
around the fore and rear passages 42 and 44.
The concave and convex walls 51 and 52,
respectively, form a cooling air passage 38 between
themselves in the region of the trailing edge 29 of the
airfoil 25. A number of pins 62 -- often referred to as
"pin fins" -- extend transversely through the passage 38
and serve to create turbulence that increases the heat
transfer coefficient of the cooling air 74 flowing through
the passage.
Although a substantial portion of the cooling air
19 flowing through the inserts 46 and 47 exits via the
holes 70 and 71 distributed around the walls of the
inserts, portions 75 and 76 of the cooling air 19 exit
through holes 72 and 73 formed in the bottom of the inserts
46 and 47, respectively, as shown in Figure 3. The cooling
air portions 75 and 76 exit the vane 17 through openings 68
and 69 in the inner shroud 26. From the openings 68 and 69
the cooling air 75 and 76 enters an annular cavity 45
formed between the inner shroud 26 and the discs 55 and 56
of the rotor 4. The first row of rotating blades 102 are
attached to the disc 55 and the second row of rotating
blades 103 are attached to the disc 56.
An interstage seal housing 66 is attached to the
inner shroud 26 by bolts (not shown) and carries a seal 33.
A plurality of labyrinth fins from the seal 33 extend into
an annular passage formed between the seal 33 and arms 48
and 49 that extend from the discs 55 and 56, respectively.
The seal housing 66 controls the flow of cooling air from
the cavity 45. Specifically, passages 50 in the housing
direct a portion of the cooling air 75 and 76 out of the
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cavity 45, whereupon it is split into two streams 85 and
86. The first stream 85 flows radially outward into the
hot gas 30 flowing through the turbine section 3. In so
doing, the cooling air 85 cools the rear face of the disc
55 and prevents the hot gas 30 from flowing over the disc
face. A honeycomb seal 87 formed in the vane inner shroud
26 regulates the flow of cooling air 85 into the hot gas
30.
The second stream 86 flows through the annular
labyrinth seal passage and then flows radially outward into
the hot gas 30 flowing through the turbine section 3. In
so doing, the cooling air 8-6 cools the front face of the
disc 56 and prevents the hot gas 30 from flowing over the
disc face.
Since the pressure of the hot gas 30 flowing over
the second row of rotating blades is lower than that
flowing over the first row of rotating blades, were it not
for the seal 33 substantially all of the cooling air would
flow downstream to the disc 56. The seal 33 prevents this
from happening, thereby ensuring cooling of the upstream
disc 55.
According to the current invention, another
portion of the cooling air 75 and 76 delivered to the
cavity 45 is used to cool the inner shroud 26.
The inner shroud 26 has a radially outward facing
surface -- that is, the surface that faces toward the outer
shroud 27 and that is exposed to the flow of hot gas 30.
The inner shroud 26 also has a radially inward facing
surface -- that is, the surface that is opposite the
radially outward facing surface and that faces toward the
seal housing 66. Fore and aft support lugs 81 and 82
extend radially inward from the shroud radially inward
facing surface and form a portion of the cavity 45.
As shown best in Figures 2 and 6, the radially
inward facing surface of the inner shroud 26 forms a raised
portion 95 around the passages 42 and 44 that is sometimes
referred to as a "race track." The radially inward facing
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surface also forms raised portions 96 and 97 along each of
the longitudinal edges 79 and 80 that are referred to as
"rails." In between these raised portions 95-97 are
recessed portions 98-100, as shown in Figures 4-6.
As shown in Figures 2 and 4-6, the two recessed
portions 98 and 99 are covered by impingement plates 83 and
84, respectively. The edges of the plates 83 and 84 are
attached to the raised portions 95-97 and the rear support
lug 82, for example, by welding. Chambers 77 and 78 are
formed between the impingement plates 83 and 84,
respectively, and the recessed portions 98 and 99,
respectively. Numerous small holes 101 are distributed
around each of the impingement plates 83 and 84 that cause
a portion of the cooling air 75 and 76 delivered to the
cavity 45 to form jets 59 that flow through the chambers 77
and 78 and impinge against the surfaces of the recessed
portions 98 and 99, thereby providing impingement cooling
of the inner shroud 26.
The jets of cooling air 59 that enter through the
holes 101 in the impingement plates 83 and 84 are collected
by the chamber 77 and 78. As shown in Figures 2-5, a
number of axially extending passages 92 and 92' are formed
in the portion of the inner shroud 26 adjacent the trailing
edge 43. The passage 92', which is located in the center
portion of the inner shroud 26, is connected directly to
the cavity 45, as shown in Figure 3. However, the passages
92 in the remaining portions of the inner shroud have
inlets that are connected to the chambers 77 and 78, as
shown in Figures 4 and S. Thus, the cooling air jets 59
collected by the chambers 77 and 78 flow through the
passages 92 and serve to cool the trailing edge portion of
the inner shroud 26 after they have accomplished the
impingement cooling.
As shown in Figures 2-5, a transversely
extending passage 88 is formed in the portion of the inner
shroud 26 adjacent the leading edge 42. The passage 88
preferably has a height in the radial direction of
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approximately 0.6 cm (0.25 inch) and extends across almost
the entirety of the width of the inner shroud 26.
According to an important aspect of the current invention,
the passage 88 contains numerous radially oriented pin fins
89 that connect the opposing walls of the passage. As
previously discussed, the array of pin fins 89 create
turbulence that increases the heat transfer coefficient of
the cooling air flowing through the passage 88. In the
preferred embodiment, the pin fins 89 have a diameter of
approximately 0.3 cm (0.12 inch) and are spaced
approximately 1.0 cm (0.4 inch) apart.
As shown in Figures 2, 4 and 7, a passage 90 is
formed in the forward support lug 81 adjacent the recessed
portion 100. The passage 90 forms an inlet that allows
another portion 57 of the cooling air 75 and 76 delivered
to the cavity 45 to enter the passage 88 by flowing
longitudinally upstream. After entering the passage 88,
the cooling air 57 turns approximately 90 and flows
transversely through the passage 88 and, with the aid of
the array of pin fins 89, cools the portion of the inner
shroud 26 adjacent the leading edge 42.
As shown in Figures 2, 5 and 7, a passage 91 is
formed in the forward support lug 81 adjacent the recessed
portion 98. The passage 91 forms an outlet that allows a
portion 58 of the cooling air 57 to exit the leading edge
passage 88 after it has flowed along substantially the
entirety of the length of the passage 88. After turning
approximately 90 , the cooling air 58 flows longitudinally
downstream and enters the chamber 77 formed by the
impingement plate 83. The chamber 77 collects the jets 59
of cooling air that flowed through the holes 101 in the
impingement passage 83 as well as cooling air 58 from the
leading edge passage 88 and directs it to the trailing edge
passages 92 that have their inlets connected to the chamber
77. The cooling air 60 flowing through the trailing edge
passages 92 then exits the inner shroud 26 through outlets
in the trailing edge 43.
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As shown in FIGS. 2 and 8, another portion 61 of the cooling air 57 that
entered the
leading edge passage 88 flows into passages 93 and 94. The passages 93 and 94
extend
longitudinally through the raised portions 96 and 97, respectively, along the
longitudinal
edges 79 and 80, respectively, of the inner shroud 26. Thus, the cooling air
61 serves to
5 cool the raised portions 96 and 97 as well as the leading edge portion of
the inner shroud
26.
As can be seen, according to the present invention, maximuin use is made of a
portion of
the cooling air 75 and 76 directed to the disc cavity 45 to provide
impingement cooling
10 of the radially inward facing surface of the inner shroud 26, as well as
convective cooling
of the leading edge portion, using the enhanced heat transfer provided by the
pin fins 89,
the trailing edge portion and the raised portions 96 and 97.
Although the present invention has been described with reference to the
shrouds of the
second row of stationary vanes in a gas turbine, the invention is also
applicable to other
rows of stationary vanes.
