Note: Descriptions are shown in the official language in which they were submitted.
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CONFIGURATION OF COOLING CHANNELS FOR COOLING THE TRAILING EDGE OF GAS TURBINE
VANES
BACKGROUND OF THE INVENTION
The present invention relates to an airfoil for
use in a gas turbine, such as for a stationary vane. More
specifically, the present invention relates to an aifoil
having an improved cooling air flow path.
A gas turbine employs a plurality of stationary
vanes that are circumferentially arranged in rows in a
turbine section. Since such vanes axe exposed to the hot
gas discharging from the combustion section,~cooling of
these vanes is of the utmost importance. Typically,
cooling is accomplished by flowing cooling air through
cavities formed inside the vane airfoil.
According to one approach, cooling of the vane
airfoil is accomplished by incorporating one or more
tubular inserts into each of the airfoil cavities so that
passages surrounding the inserts are formed between the
inserts and the walls of the airfoil. The inserts have a
number of holes distributed around their periphery that
distribute the cooling air around these passages.
According to another approach, each airfoil
cavity includes a number of radially extending passages,
typically three, forming a serpentine array. Cooling air,
supplied to the vane outer shroud, enters the first passage
and flows radially inward until it reaches the vane inner
shroud. A first portion of the cooling air exits 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. A second
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portion of the cooling air reverses direction and flows
radially outward through the second passage until it
reaches the outer shroud, whereupon it changes direction
again and flows radially inward through the third passage.
Cooling of the trailing edge portion of the vane
is especially difficult because of the thinness of the
trailing edge portion. In traditional open loop cooling
systems, the cooling air is discharged from the vane
internal cavity into the hot gas flow path by axially
oriented passages in the trailing edge of the airfoil. In
closed loop systems, the trailing edge portion of the vane
airfoil may be cooled by directing the cooling air through
a channel that wraps around in the trailing edge in the
chord-wise direction. However, this approach results in a
thick trailing edge, which is aerodynamically undesirable,
and increased manufacturing complexity.
In another approach, the cooling air is directed
through span-wise radial holes extending between the inner
and outer shrouds, with the air flowing either radially
outward from the inner shroud to the outer shroud or
radially inward from the outer shroud to the inner shroud.
Unfortunately, this approach suffers from several
disadvantages. First, the cooling air can become
sufficiently heated by the time it reaches the ends of the
holes that its cooling effectiveness is inadequate, thereby
resulting in over-heating of the portion of the trailing
edge adjacent to the inner or outer shroud. Also, if the
diameter of the holes is relatively small, the length of
the holes results in an undesirably high pressure drop in
the cooling air. However, reducing the pressure drop by
increasing the diameter of the holes results in undesirably
thick trailing edges..
Span-wise radial holes are also difficult to
manufacture. If the airfoil is cast, the use of long,
small diameter span-wise radial holes can result it long,
unsupported, and therefore weak, casting cores. In
addition, such long cooling holes makes it difficult to
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maintain wall thickness tolerances, and results in a long
leaching time.
It is therefore desirable to provide a cooling
scheme for cooling the trailing edge portion of an airfoil
that overcomes the problems of previous approaches,
including the minimization of both the heat up of the
cooling fluid by the time it reaches the end of the cooling
path and the pressure drop experienced by the fluid.
SUMMARY OF THE INVENTION
Accordingly, it is the general object of the
current invention to provide a cooling scheme for cooling
the trailing edge portion of an airfoil that overcomes the
problems of previous approaches,, including the minimization
of both the heat up of the cooling fluid by the time it
reaches the end of the cooling path and the pressure drop
experienced by the fluid.
Briefly, this object, as well as other objects of
the current invention, is accomplished in an airfoil for a
gas turbine, comprising (i> a leading edge and a trailing
edge, (ii) first and second ends, the first end disposed
radially outward from the second end, (iii) first and
second side walls, (iv) a first passage formed between the
first and second sidewalls, the first passage having an
inlet for receiving a flow of a cooling fluid directed to
the airfoil, (v) a plenum disposed between the first and
second ends, the plenum in flow communication with the
first passage, (vi) a plurality of second passages in flow
communication with the plenum, the second passages
extending in a substantially radial direction from the
plenum toward the first end, (v) a plurality of third
passages in flow communication with the plenum, the third
passages extending in a substantially radial direction from
the plenum toward the second end.
In a preferred embodiment of the invention, the
plenum is disposed at about mid-height adjacent the
trailing edge of the airfoil.
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In accordance with one aspect of this invention,
there is provided an airfoil for a turbomachine, comprising:
a) a leading edge and a trailing edge; b) first and second
ends, said first end disposed radially outward from said
second end; c) first and second side walls; d) a first
passage formed between said first and second sidewalls, said
first passage having an inlet for receiving a flow of a
cooling fluid directed to said airfoil; e) a plenum formed
between said side walls and disposed between said first and
second ends, said plenum in flow communication with said
first passage; f) a plurality of second passages in flow
communication with said plenum, said second passages
extending in a substantially radial direction from said
plenum toward said first end; and g) a plurality of third
passages in flow communication with said plenum, said third
passages extending in a substantially radial direction from
said plenum toward said second end.
In accordance with another aspect of this
invention, there is provided a gas turbine vane, comprising:
a) a leading edge and a trailing edge; b) first and second
sidewalk ; c) inner and outer shrouds; d) a cavity disposed
between said first and second sidewalls, said cavity having
an inlet for receiving a flow of cooling fluid directed to
said airfoil; e) a plenum disposed between said cavity and
said trailing edge approximately midway between said inner
and outer shrouds, an opening formed between said plenum and
said cavity; f) a first plurality of passages formed in an
array adjacent said trailing edge, said first plurality of
passages extending in a substantially radially outward
direction from said plenum to said outer shroud; and g) a
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second plurality of passages formed in an array adjacent
said trailing edge, said second plurality of passages
extending in a substantially radially inward direction from
t said plenum to said inner shroud.
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BRIEF DESCRIPTION OF THE DRAWINGS
Figure 1 is a longitudinal cross-section through
a gas turbine vane of the current invention.
Figure 2 is a is transverse cross-section taken
through line II-II shown in Figure 1.
Figure 3 is a is transverse cross-section taken
through line III-III shown in Figure 1.
Figure 4 is an isometric view of a portion of the
trailing edge of the vane shown in Figure 1 in the vicinity
of the plenum.
DESCRIPTION OF THE PREFERRED EMBODIMENT
Referring to the drawings, there is shown in
Figures 1-4 a vane 1 having an airfoil according to the
current invention for use in the turbine section of a gas
turbine. The vane 1 is comprised of an airfoil 6 having an
inner shroud 2 on one end and an outer shroud 4 on the
other end. As shown best in Figure 2, the airfoil portion
6 of the vane 1 is formed by opposing side walls 9 and 11
that meet to form a leading edge 8 and a trailing edge 10.
The current invention concerns an apparatus for cooling the
airfoil 6, preferably the portion of the airfoil adjacent
the trailing edge 10.
The major portion of the airfoil 6 is hollow.
Transversely extending ribs 48, 50, and 52 divide the
hollow interior of the airfoil 6 into three cooling air
passages 32, 34, and 36. The first passage 32 is a cooling
air supply passage and is formed in the portion of the
airfoil 6 adjacent the leading edge 8. The second passage
34 is also a cooling supply passage but is formed in the
vicinity.of the trailing edge 6. A passage 17 in the inner
shroud 2 connects the passages 32 and 34. The third
passage 36 is formed in the mid-chord region of the airfoil
6 and forms a cooling air discharge passage.
Referring to Figure 1, a cooling fluid supply
pipe 13 is connected to the outer shroud 4. An opening 18
in the outer shroud 4 allows the supply pipe 13 to
communicate with a passage 16 formed within the outer
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shroud. The outer shroud passage 16 is connected to
passages 32 and 34 in the airfoil 6.
As shown best in Figures 2 and 4, according to an
important aspect of the current invention, a cavity 42 is
5 formed between the side walls 9 and 11 that acts as a
plenum. The plenum 42 is preferably located at
approximately mid-height and adjacent the trailing edge 10
of the airfoil 6. An opening 40 in the rib.52 connects the
plenum 42 with the supply passage 34.
As shown best in Figures 1 and 3, a first array
of cooling fluid holes 38' extend radially outward from the
plenum 42 to a cooling fluid manifold 54 formed in the
outer shroud 4, with the inlets to the holes being at the
plenum and the outlets being at the manifold. As shown in
Figure 3, a passage 58 is formed in the outer shroud 4 that
extends generally perpendicularly to the radial direction.
The passage 58 extends from the manifold 54 around the
portion of the airfoil 6 projecting-into the outer shroud.
Openings 46 and 47 are formed in the portions of the side
walls 9 and 11, respectively, that extend into the outer
shroud 4. The openings 46 and 47 allow the passage 58 to
communicate with the discharge passage 36. As shown in
Figure 1, an outlet 30 is formed in the discharge passage
36 and is connected to a return pipe 14.
As shown best in Figures 1, 2 and 4, a second
array of cooling fluid holes 38", which are preferably
radially aligned with the cooling fluid holes 38', extend
radially inward from the plenum 42 to a cooling fluid
manifold 56 formed in the inner shroud 2, with the inlets
to the holes being at the plenum and the outlets being at
the manifold. A passage (not shown), similar to passage 58
in the outer shroud 4, is formed in the inner shroud 2 that
extends from the manifold 56 around the portion of the
airfoil 6 projecting into the inner shroud. Openings 44,
one of which is shown in Figure 1, which are similar to
openings 46 and 47 at the outer shroud 4, are formed in the
portions of the side walls 9 and 11, respectively, that
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extend into the inner shroud 2. The openings 44 allow the
passage in the inner shroud 2 to communicate with the
discharge passage 36.
It should be understood that the inner and outer
shrouds may contain cooling passages, in addition to those
connecting the trailing edge cooling fluid manifolds 54 and
56 to the discharge passage 36, that aid in the cooling of
the shrouds themselves. However, such shroud cooling is
not part of the current invention, which concerns the
cooling of the airfoil 6 and, preferably, the portion of
the airfoil adjacent the trailing edge 10.
In operation, cooling fluid, which in the
preferred embodiment is compressed air 20, typically bled
from the compressor section of the gas turbine, is directed
to the vane outer shroud 4 by the supply pipe 13, as shown
in Figure 1. According to a preferred embodiment of the
invention, the vane 1 has cooling passages that are part of
a closed loop cooling air system. Thus, essentially all of
the cooling air supplied to the vane 1 is returned to the
cooling system.
Upon flowing through the opening 1B and entering
the passage 16 in the outer shroud 4, the cooling air 20 is
divided into two streams 22 and 24. The first cooling air
stream 22 flows radially inward through the trailing edge
supply passage 34 to the plenum 42 and, in so doing, cools
a portion of the side walls 9 and 11 of the airfoil 6.
The second cooling air stream 24 flows radially
inward through the leading edge supply passage 32 and cools
the leading edge a portion of the airfoil 6. The passage
17 in the inner shroud 2 then directs the cooling air 24
from the passage 32 to the passage 34, where it flows
radially outward (that is, toward the outer shroud 4) to
the plenum 42. In the plenum 42, the cooling air streams
22 and 24 combine and are then divided into numerous small
streams by the trailing edge cooling holes 38. As shown
best in Figures 2 and 4, the plenum is tapered as it
extends in the axial direction toward the trailing edge 10
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of the airfoil 6. Such tapering provides the area
reduction necessary for uniform flow distribution among the
cooling holes 38.
A portion 28 of the combined flow of cooling air
22 and 24, flows radially outward (that is, toward the
outer shroud 4) from the plenum 42 through the holes 38' to
the manifold 54, thereby providing vigorous cooling of the
approximately upper half portion of the airfoil 6 adjacent
the trailing edge 10 that is located above the plenum 42.
In the manifold 54, the individual streams of cooling air
28 are collected and are then directed by passage 58 to the
openings 46 and 47, as shown in Figure 3. From the
openings 46 and 47, the cooling air 28 enters the discharge
passage 36 and flows radially outward to the exhaust pipe
14, as shown in Figure 1.
Similarly, a portion 26 of the combined flow of
cooling air 22 and 24, flows radially inward from the
plenum 42 through the holes 38" to the manifold 56, thereby
providing vigorous cooling of the approximately lower half
portion of the airfoil 6 adjacent the trailing edge l0
below the plenum 42. In the manifold 56, the individual
streams of cooling air 26 are collected and are then
directed by the inner shroud passage to the openings 44, as
discussed above with respect to the outer shroud 4. From
the openings 44, the cooling air 26 enters the discharge
passage 36 and flows radially outward to the exhaust pipe
14 and, in so doing, cools the mid-chord portion of the
side walls 9 and 11 of the airfoil 6. In the preferred
embodiment of the invention, the exhaust pipe 14 directs
the cooling air 29 to a cooler for recycling back to the
turbine.
The present~invention has numerous advantages
over traditional airfoil cooling schemes. First, since the
length of the cooling air passages 38 is effectively cut in
half, compared to span-wise holes that extend from the
inner shroud to the outer shroud, there is less chance of
overheating the coolant, which may be air or steam, for
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example, by the time it reaches a shroud. Also, the
pressure drop through the passages 38 is reduced, thereby
allowing the use of holes 38 of minimum diameter. Small
diameter holes permit the use of a thin trailing edge 10,
which has aerodynamic advantages. The airfoil 6 is also
easier to manufacture since long runs of cooling holes are
avoided.
Although the current invention has been discussed
in connection with the.airfoil for a stationary vane in a
gas turbine, the invention is also applicable to other
types of components. In addition, although the invention
has been discussed with reference to a closed loop cooling
system utilizing compressed air, the invention is also
applicable to more conventional open loop systems as well
as to systems using other types of cooling fluids, such as
steam. Thus, the present invention may be embodied in
other specific forms without departing from the spirit or
essential attributes thereof and, accordingly, reference
should be made to the appended claims, rather than to the
foregoing specification, as indicating the scope of the
invention.
