U.S. patent application number 12/985571 was filed with the patent office on 2012-07-12 for inner shroud cooling arrangement in a gas turbine engine.
Invention is credited to Gm Salam Azad, Zhihong Gao, Ching-Pang Lee.
Application Number | 20120177479 12/985571 |
Document ID | / |
Family ID | 46455383 |
Filed Date | 2012-07-12 |
United States Patent
Application |
20120177479 |
Kind Code |
A1 |
Azad; Gm Salam ; et
al. |
July 12, 2012 |
INNER SHROUD COOLING ARRANGEMENT IN A GAS TURBINE ENGINE
Abstract
A component in a gas turbine engine includes an airfoil and a
shroud. The shroud has an outer surface supporting an end of the
airfoil and defines a portion of an annular gas path. The shroud
includes axial edges extending between upstream and downstream
edges thereof. Each of the axial edges includes a seal slot that
receives a seal member extending between the shroud and an adjacent
shroud. A cooling air channel extends between the upstream and
downstream edges of the shroud. A cooling air supply passage
extends from a cooling air chamber at an inner surface of the
shroud to the cooling air channel. At least one cooling air exit
passage extends from the cooling air channel to one of the axial
edges. The cooling air channel is located radially between the
outer surface of the shroud and the seal slot.
Inventors: |
Azad; Gm Salam; (Oviedo,
FL) ; Lee; Ching-Pang; (Cincinnati, OH) ; Gao;
Zhihong; (Orlando, FL) |
Family ID: |
46455383 |
Appl. No.: |
12/985571 |
Filed: |
January 6, 2011 |
Current U.S.
Class: |
415/115 |
Current CPC
Class: |
F05D 2260/202 20130101;
F05D 2240/59 20130101; F05D 2260/60 20130101; F05D 2240/81
20130101; F01D 9/041 20130101; F01D 5/187 20130101 |
Class at
Publication: |
415/115 |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Claims
1. A component in a gas turbine engine, said component comprising:
an airfoil adapted to extend radially through an annular hot gas
path extending in a generally axial direction through said turbine
engine, said airfoil including a pressure side and a suction side,
an upstream leading edge and a downstream trailing edge; a shroud
having an outer surface supporting an end of said airfoil, said
shroud defining a portion of said annular gas path through said gas
turbine engine and including an upstream edge and a downstream
edge, and opposing axial edges extending between said upstream edge
and said downstream edge; each of said axial edges including a
generally axially extending seal slot adapted to receive a seal
member extending between said shroud and an adjacent shroud; a
cooling air channel extending generally axially substantially
parallel to at least one of said axial edges between said upstream
edge and said downstream edge; a cooling air supply passage
extending from a cooling air chamber at an inner surface of said
shroud to said cooling air channel; at least one cooling air exit
passage extending from said cooling air channel to said one of said
axial edges; and said cooling air channel being located radially
between said outer surface of said shroud and said seal slot at
said one of said axial edges for effecting convective cooling of a
corner defined at an intersection of said outer surface and said
one of said axial edges.
2. The component of claim 1, wherein said cooling air channel is
located on a radial plane passing through said seal slot at said
one of said axial edges.
3. The component of claim 1, wherein said cooling air exit passage
includes an exit opening located between said outer surface of said
shroud and said seal slot.
4. The component of claim 1, wherein said cooling air exit passage
comprises a purge passage having an exit opening providing a volume
of air for purging hot gas from said seal slot and said seal, said
exit opening being located at a downstream end of said cooling air
channel adjacent said downstream edge of said shroud.
5. The component of claim 4, further comprising a replenishing
cooling air supply passage extending from a mid-chord cooling air
chamber at said inner surface of said shroud to said cooling air
channel, said replenishing cooling air supply passage located
between said cooling air supply passage and said shroud downstream
edge.
6. The component of claim 5, wherein said cooling air supply
passage supplies a first portion of cooling air from a leading edge
cooling air chamber to a location adjacent an upstream end of said
cooling air channel and said replenishing cooling air supply
passage supplies a second portion of cooling air from said
mid-chord cooling air chamber to a location proximate to said purge
passage.
7. The component of claim 5, wherein said component comprises a
vane in a first row of vanes within said gas turbine engine.
8. The component of claim 1, wherein said cooling air exit passage
comprises a plurality of impingement passages having exit openings
providing a flow of cooling air impinging on an axial edge of said
adjacent shroud, said exit openings being located at an upstream
end of said cooling air channel adjacent said upstream edge of said
shroud.
9. The component of claim 8, wherein said upstream end of said
cooling air channel receives cooling air from a leading edge
cooling air chamber at said inner surface of said shroud.
10. The component of claim 9, wherein said cooling air supply
passage comprises a replenishing cooling air supply passage that
supplies cooling air from a mid-shroud impingement cavity located
at an axial midpoint between said upstream edge and said downstream
edge at said inner surface of said shroud, said replenishing
cooling air supply passage providing replenishing cooling air to
said cooling air channel.
11. The component of claim 9, wherein said component comprises a
vane in a second row of vanes within said gas turbine engine.
12. A vane in a gas turbine engine, said vane comprising: an
airfoil adapted to extend radially through an annular hot gas path
extending in a generally axial direction through said turbine
engine, said airfoil including a pressure side and a suction side,
an upstream leading edge and a downstream trailing edge; a shroud
having an outer surface supporting an end of said airfoil, said
shroud defining a portion of said annular gas path through said gas
turbine engine and including an upstream edge and a downstream
edge, and opposing axial edges extending between said upstream edge
and said downstream edge; each of said axial edges including a
generally axially extending seal slot adapted to receive a seal
member extending between said shroud and an adjacent shroud; a
cooling air channel extending generally axially substantially
parallel to at least one of said axial edges between said upstream
edge and said downstream edge; a cooling air supply passage
extending from a cooling air chamber at an inner surface of said
shroud to said cooling air channel; at least one cooling air exit
passage extending from said cooling air channel to said one of said
axial edges, said cooling air exit passage comprising a purge
passage having an exit opening providing a volume of air for
purging hot gas from said seal slot and said seal, said exit
opening being located at a downstream end of said cooling air
channel adjacent said downstream edge of said shroud; and said
cooling air channel being located radially between said outer
surface of said shroud and said seal slot at said one of said axial
edges for effecting convective cooling of a corner defined at an
intersection of said outer surface and said one of said axial
edges.
13. The vane of claim 12, further comprising a replenishing cooling
air supply passage extending from a mid-chord cooling air chamber
at said inner surface of said shroud to said cooling air channel,
said replenishing cooling air supply passage located between said
cooling air supply passage and said shroud downstream edge.
14. The vane of claim 13, wherein said cooling air supply passage
supplies a first portion of cooling air from a leading edge cooling
air chamber to a location adjacent an upstream end of said cooling
air channel and said replenishing cooling air supply passage
supplies a second portion of cooling air from said mid-chord
cooling air chamber to a location proximate to said purge
passage.
15. The vane of claim 12, wherein said cooling air channel is
without openings for discharge of air along a length of said
cooling air channel from said cooling air supply passage to within
close proximity of said replenishing cooling air supply
passage.
16. The vane of claim 12, wherein said cooling air channel is
located on a radial plane passing through said seal slot at said
one of said axial edges.
17. A vane in a gas turbine engine, said vane comprising: an
airfoil adapted to extend radially through an annular hot gas path
extending in a generally axial direction through said turbine
engine, said airfoil including a pressure side and a suction side,
an upstream leading edge and a downstream trailing edge; a shroud
having an outer surface supporting an end of said airfoil, said
shroud defining a portion of said annular gas path through said gas
turbine engine and including an upstream edge and a downstream
edge, and opposing axial edges extending between said upstream and
downstream edges; each of said axial edges including a generally
axially extending seal slot adapted to receive a seal member
extending between said shroud and an adjacent shroud; a cooling air
channel extending generally axially substantially parallel to at
least one of said axial edges between said upstream and said
downstream edges; a cooling air supply passage extending from a
cooling air chamber at an inner surface of said shroud to said
cooling air channel; a plurality of impingement passages extending
from said cooling air channel to said one of said axial edges, said
impingement passages having exit openings providing a flow of
cooling air impinging on an axial edge of said adjacent shroud,
said exit openings being located at an upstream end of said cooling
air channel and adjacent said upstream edge of said shroud; and
said cooling air channel being located radially between said outer
surface of said shroud and said seal slot at said one of said axial
edges for effecting convective cooling of a corner defined at an
intersection of said outer surface and said one of said axial
edges.
18. The vane of claim 17, wherein said upstream end of said cooling
air channel receives cooling air from a leading edge cooling air
chamber at said inner surface of said shroud.
19. The vane of claim 18, wherein said cooling air supply passage
comprises a replenishing cooling air supply passage that supplies
cooling air from a mid-shroud impingement cavity located at an
axial midpoint between said upstream edge and said downstream edge
at said inner surface of said shroud, said replenishing cooling air
supply passage providing replenishing cooling air to said cooling
air channel.
20. The vane of claim 17, wherein said cooling air channel is
located on a radial plane passing through said seal slot at said
one of said axial edges.
Description
FIELD OF THE INVENTION
[0001] The present invention relates to turbine engines and, more
particularly, to cooling arrangements for inner shrouds of vane
segments in gas turbine engines.
BACKGROUND OF THE INVENTION
[0002] In a turbomachine, such as a gas turbine engine, air is
pressurized in a compressor section then mixed with fuel and burned
in a combustor section to generate hot combustion gases. The hot
combustion gases are expanded within a turbine section of the
engine where energy is extracted to power the compressor section
and to produce useful work, such as turning a generator to produce
electricity. The hot combustion gases travel through a series of
turbine stages within the turbine section. A turbine stage may
include a row of stationary airfoils, i.e., vanes, followed by a
row of rotating airfoils, i.e., turbine blades, where the turbine
blades extract energy from the hot combustion gases for powering
the compressor section and providing output power. Because the
airfoils, i.e., vanes and turbine blades, are directly exposed to
the hot combustion gases, they are typically provided with internal
cooling channels that may feed a cooling fluid, such as compressor
bleed air, through the airfoil and through various passages formed
in structure associated with the vanes and/or blades.
[0003] One type of stationary airfoil in a turbine engine is
provided as a component of a stator vane segment. The stator vane
segment may include a radially inner shroud, a radially outer
shroud, and one or more airfoils extending between the inner and
outer shrouds. Hot combustion gases, or working gases, may be
supplied from a combustor section and pass through passages defined
between adjacent airfoils and between the inner and outer shrouds,
resulting in some of the heat of the gases being transferred to the
vane segments. As turbine engine performance has been increased
with increasing combustion gas temperature, there has been a
continuing need to improve cooling to the various portions of vane
segments in order to avoid or minimize deterioration of the
material forming the vane segments.
SUMMARY OF THE INVENTION
[0004] In accordance with a first aspect of the present invention,
a component is provided in a gas turbine engine. The component
comprises an airfoil and a shroud. The airfoil is adapted to extend
radially through an annular hot gas path extending in a generally
axial direction through the turbine engine. The airfoil includes a
pressure side and a suction side, an upstream leading edge and a
downstream trailing edge. The shroud has an outer surface
supporting an end of the airfoil and defines a portion of the
annular gas path through the gas turbine engine. The shroud
includes an upstream edge, a downstream edge, and opposing axial
edges extending between the upstream edge and the downstream edge.
Each of the axial edges includes a generally axially extending seal
slot adapted to receive a seal member extending between the shroud
and an adjacent shroud. A cooling air channel extends generally
axially substantially parallel to at least one of the axial edges
between the upstream edge and the downstream edge. A cooling air
supply passage extends from a cooling air chamber at an inner
surface of the shroud to the cooling air channel. At least one
cooling air exit passage extends from the cooling air channel to
the one of the axial edges. The cooling air channel is located
radially between the outer surface of the shroud and the seal slot
at the one of the axial edges for effecting convective cooling of a
corner defined at an intersection of the outer surface and the one
of the axial edges.
[0005] In accordance with a second aspect of the present invention,
a vane is provided in a gas turbine engine. The vane comprises an
airfoil and a shroud. The airfoil is adapted to extend radially
through an annular hot gas path extending in a generally axial
direction through the turbine engine. The airfoil includes a
pressure side and a suction side, an upstream leading edge and a
downstream trailing edge. The shroud has an outer surface
supporting an end of the airfoil and defines a portion of the
annular gas path through the gas turbine engine. The shroud
includes an upstream edge, a downstream edge, and opposing axial
edges extending between the upstream edge and the downstream edge.
Each of the axial edges includes a generally axially extending seal
slot adapted to receive a seal member extending between the shroud
and an adjacent shroud. A cooling air channel extends generally
axially substantially parallel to at least one of the axial edges
between the upstream edge and the downstream edge. A cooling air
supply passage extends from a cooling air chamber at inner surface
of the shroud to the cooling air channel. At least one cooling air
exit passage extends from the cooling air channel to the one of the
axial edges, the cooling air exit passage comprising a purge
passage having an exit opening providing a volume of air for
purging hot gas from the seal slot and the seal. The exit opening
is located at a downstream end of the cooling air channel adjacent
the downstream edge of the shroud. The cooling air channel is
located radially between the outer surface of the shroud and the
seal slot at the one of the axial edges for effecting convective
cooling of a corner defined at an intersection of the outer surface
and the one of the axial edges.
[0006] In accordance with a third aspect of the present invention,
a vane is provided in a gas turbine engine. The vane comprises an
airfoil and a shroud. The airfoil is adapted to extend radially
through an annular hot gas path extending in a generally axial
direction through the turbine engine. The airfoil includes a
pressure side and a suction side, an upstream leading edge and a
downstream trailing edge. The shroud has an outer surface
supporting an end of the airfoil and defines a portion of the
annular gas path through the gas turbine engine. The shroud
includes an upstream edge, a downstream edge, and opposing axial
edges extending between the upstream and downstream edges. Each of
the axial edges includes a generally axially extending seal slot
adapted to receive a seal member extending between the shroud and
an adjacent shroud. A cooling air channel extends generally axially
substantially parallel to at least one of the axial edges between
the upstream and the downstream edges. A cooling air supply passage
extends from a cooling air chamber at an inner surface of the
shroud to the cooling air channel. A plurality of impingement
passages extend from the cooling air channel to the one of the
axial edges, the impingement passages having exit openings
providing a flow of cooling air impinging on an axial edge of the
adjacent shroud, the exit openings being located at an upstream end
of the cooling air channel and adjacent the upstream edge of the
shroud. The cooling air channel is located radially between the
outer surface of the shroud and the seal slot at the one of the
axial edges for effecting convective cooling of a corner defined at
an intersection of the outer surface and the one of the axial
edges.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] While the specification concludes with claims particularly
pointing out and distinctly claiming the present invention, it is
believed that the present invention will be better understood from
the following description in conjunction with the accompanying
Drawing Figures, in which like reference numerals identify like
elements, and wherein:
[0008] FIG. 1 is a perspective view of a portion of a row of vane
segments including inner shroud cooling arrangements according to
an embodiment of the invention;
[0009] FIG. 2 is a perspective view of a radially inner side of one
of the vane segments of FIG. 1 and including cut-away portions
illustrating an inner shroud cooling arrangement;
[0010] FIG. 3 is a diagrammatic top perspective view looking in a
radially inward direction of the vane segment shown in FIG. 2,
wherein the outer shroud and airfoil of the vane segment have been
removed for clarity;
[0011] FIG. 4 is an end view looking in an axial direction
illustrating mating edges of two vane segments of the row of vane
segments illustrated in FIG. 1;
[0012] FIG. 5 is a top view looking in a radially inward direction
illustrating mating edges of two vane segments of the row of vane
segments illustrated in FIG. 1;
[0013] FIG. 6 is a diagrammatic top perspective view looking in a
radially inward direction of a vane segment including an inner
shroud cooling arrangement in a row of vane segments according to
another embodiment of the invention, wherein the outer shroud and
airfoils of the vane segment have been removed for clarity;
[0014] FIG. 7 is an end view looking in an axial direction
illustrating mating edges of the vane segment illustrated in FIG. 6
and an adjacent vane segment; and
[0015] FIG. 8 is a top view looking in a radially inward direction
illustrating mating edges of the vane segment illustrated in FIG. 6
and an adjacent vane segment.
DETAILED DESCRIPTION OF THE INVENTION
[0016] In the following detailed description of the preferred
embodiments, reference is made to the accompanying drawings that
form a part hereof, and in which is shown by way of illustration,
and not by way of limitation, specific preferred embodiments in
which the invention may be practiced. It is to be understood that
other embodiments may be utilized and that changes may be made
without departing from the spirit and scope of the present
invention.
[0017] Referring to FIG. 1, a portion of a row 10 of vanes or
stator vane segments 12 is illustrated, such as may be incorporated
in a turbine section 14 of a turbine engine. The row 10 of vane
segments 12 shown in FIG. 1 comprises a first row 10 of vane
segments 12, also referred to as row one vane segments, of the
turbine section 14. The vane segments 12 each include at least one
airfoil 16, a radially inner shroud 18 rigidly connected to a
radially inner end of the airfoil 16, and a radially outer shroud
20 rigidly connected to a radially outer end of the airfoil 16.
Each airfoil 16 comprises a generally concave pressure side 22 and
a generally convex suction side 24. The pressure and suction sides
22, 24 are connected at an upstream leading edge 26 and at a
chordally spaced downstream trailing edge 28 (FIG. 3).
[0018] Referring to FIG. 1, the outer shroud 20 is suspended
radially inwardly from a casing structure (not shown) of the
turbine section, and the row 10 of vane segments 12 comprises a
plurality of vane segments 12 suspended in side-by-side relation
extending circumferentially about a turbine rotor (not shown)
within the turbine engine. Although the vane segments 12 are
illustrated as each including one airfoil 16, it should be
understood that the vane segments 12 may be constructed with two or
more airfoils 16. A hot working gas H.sub.G created in a
conventional combustor assembly (not shown) is discharged into the
turbine section 14 and passes between adjacent airfoils 16 of the
vane segments 12. The vane segments 12 direct the hot working gas
H.sub.G toward rows of blades (not shown) in the turbine section,
which blades are caused to rotate and cause corresponding rotation
of the turbine rotor.
[0019] As noted above, the vane segments 12 are suspended in a
circumferential row 10 about the turbine rotor, such that the
airfoils 16 are spaced apart and define flow passages 30
therebetween for channeling the hot working gas H.sub.G through the
turbine section 14 during engine operation. Each flow passage 30
forms a portion of an annular path for the hot working gas H.sub.G
and is bounded by the pressure sidewall 22 of one airfoil 16 and
the suction sidewall 24 of an adjacent airfoil 16. The flow passage
30 is also defined between the inner shroud 18 and the outer shroud
20 and extends in a flow direction from an upstream edge 32 to a
downstream edge 34 of the inner shroud 18 (FIGS. 2 and 3) and from
an upstream edge 36 to a downstream edge 38 of the outer shroud 20
(FIGS. 1 and 2).
[0020] Each vane segment 12 includes a first generally axially
extending mating edge 40 extending between the upstream edge 32 and
the downstream edge 34 of the inner shroud 18, and an opposing
second generally axially extending mating edge 42 extending
generally parallel to the first mating edge 40 between the upstream
edge 32 and the downstream edge 34 of the inner shroud 18. Each of
the mating edges 40, 42 includes a generally axially extending seal
slot 44 adapted to receive an axially extending seal member 46
(FIG. 1). The seal member 46 spans across a gap formed between
adjacent mating edges 40, 42 of adjacent vane segments 12 to
separate the hot working gas H.sub.G from a region of cool air
located radially inwardly from the inner shroud 18.
[0021] Referring to FIGS. 2-5, respective primary cooling air
channels 48 extend within the inner shroud 18 generally axially
adjacent to and substantially parallel to each of the mating edges
40, 42. As illustrated in FIG. 3, in the embodiment shown the
channels 48 extend from upstream portions 48A thereof adjacent to
the upstream edge 32 of the inner shroud 18 to outlets 49 located
at downstream portions 48B thereof adjacent to the downstream edge
34 of the inner shroud 18. Each channel 48 is located radially
between an outer surface 50 defined by an outer wall 51 of the
inner shroud 18 and the seal slot 44, see FIG. 4. Further, each
channel 48 is preferably aligned circumferentially with a
respective seal slot 44, such that the channels 48 are located on,
or intersected by, respective radial planes 54 passing radially
through the corresponding seal slots 44, see FIG. 4. Preferably,
the entirety of a cross-sectional area of the cooling air channel
48, as viewed in the axial direction, is located radially outwardly
from, i.e., substantially overlies, a respective seal slot 44, as
shown in FIG. 4. The cooling air channels 48 effect convective
cooling of a corner 56 at an intersection of the outer surface 50
of the outer wall 51 and one of the mating edges 40, 42 of the
inner shroud 18, as will be described herein. It is noted that the
spacing between the outer surface 50 of the outer wall 51 and the
seal slots 44 is large enough to accommodate the respective cooling
air channels 48 therebetween while still maintaining the structural
rigidity of the inner shroud 18 at the respective mating edges 40,
42.
[0022] Referring now to FIGS. 2 and 3, an inner surface 52 (FIGS. 2
and 4) of the outer wall 51 defines a radially outer boundary for a
leading edge cooling air chamber 53 associated with the inner
shroud 18. The chamber 53 receives cooling air, i.e., compressor
discharge air, for cooling the inner shroud 18 proximate to the
upstream edge 32 thereof. In the embodiment shown, cooling air
enters the chamber 53 from a first inner cavity 55 (FIGS. 1 and 2)
located radially inwardly from the inner shroud 18 proximate to the
upstream edge 32 thereof. The cooling air enters the chamber 53
through impingement holes 57A formed in an impingement plate 57,
which impingement plate 57 defines a radially inner boundary for
the chamber 53, see FIG. 2. The cooling air passing through the
holes 57A impinges on the inner surface 52 of the outer wall 51,
and then a portion of the cooling air flows through film cooling
holes 59 (FIGS. 1 and 3) formed in the outer wall 51. The cooling
air passing through the film cooling holes 59 provides film cooling
for the outer surface 50 of the outer wall 51 and may then be swept
up and carried through the annular path with the hot working gas
H.sub.G.
[0023] As shown in FIGS. 2 and 3, primary cooling air supply
passages 58 provide fluid communication between the leading edge
cooling air chamber 53 and the upstream portions 48A of the
respective cooling air channels 48 for supplying a first portion of
cooling air from the chamber 53 to the channels 48. The passages 58
may include a component in the radial direction, since the channels
48 in the embodiment shown may be located in the outer wall 51 of
the inner shroud 18 radially outwardly from the chamber 53.
Further, the passages 58 may be angled toward the downstream edge
34 of the inner shroud 18 to promote a flow toward the downstream
portions 48B of the cooling air channels 48.
[0024] The inner surface 52 of the outer wall 51 further defines a
radially outer boundary for a mid-chord cooling air chamber 60
(FIGS. 2-4).
[0025] The mid-chord cooling air chamber 60 receives cooling air,
i.e., compressor discharge air, for cooling a mid-chord portion 18A
and a trailing edge portion 18B of the inner shroud 18, see FIGS. 2
and 3. In the embodiment shown, cooling air enters the chamber 60
from a second inner cavity 62 (FIG. 2) located radially inwardly
from the inner shroud 18 proximate to the mid-chord portion 18A
thereof. The cooling air enters the chamber 60 through impingement
holes 63A formed in a plurality of impingement plates 63 associated
with the mid-chord portion 18A, which impingement plates 63 define
a radially inner boundary for the mid-chord cooling air chamber 60,
see FIG. 2. It is noted that additional or fewer impingement plates
63 may be used than as shown in FIG. 2 to define additional or
fewer sub-chambers of the mid-chord cooling air chamber 60. For
example, the mid-chord cooling air chamber 60 illustrated
diagrammatically in FIG. 3 has been consolidated into a single
chamber, whereas the mid-chord cooling air chamber 60 illustrated
in FIG. 2 comprises a plurality of sub-chambers that collectively
form the mid-chord cooling air chamber 60. A portion of the cooling
air passing through the holes 63A impinges on the inner surface 52
of the outer wall 51, and then flows through film cooling holes 64
(FIGS. 1 and 3) formed in the outer wall 51. The cooling air
passing through the film cooling holes 64 provides film cooling for
the outer surface 50 of the outer wall 51 and may then be swept up
and carried through the annular path with the hot working gas
H.sub.G.
[0026] The portion of the mid-chord cooling air chamber 60
associated with the trailing edge portion 18B of the inner shroud
18 is associated with a cover plate 66 (FIG. 2) that covers and
defines a radially inner boundary for the trailing edge portion 18B
of the mid-chord cooling air chamber 60. The cooling air passing
into the trailing edge portion 18B of the mid-chord cooling air
chamber 60 passes over a bridge 65 that spans between the mating
edges 40, 42, i.e., between the outer wall 51 and the bridge 65.
The cooling air in the trailing edge portion 18B of the mid-chord
cooling air chamber 60 provides convective cooling for the trailing
edge portion 18B of the inner shroud 18, and then flows through
film cooling holes 67 (FIG. 3) formed in the outer wall 51. The
cooling air passing through the film cooling holes 67 provides film
cooling for the outer surface 50 of the outer wall 51 at the
trailing edge portion 18B of the inner shroud 18 and may then be
swept up and carried through the annular path with the hot working
gas H.sub.G.
[0027] Referring to FIGS. 2-5, replenishing cooling air supply
passages 68 provide fluid communication between the mid-chord
cooling air chamber 60 and the respective cooling air channels 48
at a location between the upstream and downstream portions 48A and
48B thereof, i.e., between the primary cooling air supply passages
58 and the downstream edge 34 of the inner shroud 18, for supplying
a second portion of cooling air from the mid-chord cooling air
chamber 60 to the channels 48. As shown in FIG. 4, the passages 68
may include a component in the radial direction, since the channels
48 in the embodiment shown may be located in the outer wall 51 of
the inner shroud 18 radially outwardly from the mid-chord cooling
air chamber 60. Further, the passages 68 may be angled toward the
downstream edge 34 of the inner shroud 18 to promote a flow toward
the downstream portions 48B of the cooling air channels 48, as
shown in FIG. 5.
[0028] As shown in FIGS. 2-5, one or more cooling air exit passages
69 extend from each cooling air channel 48 to the respective axial
mating edge 40, 42. The cooling air exit passages 69 according to
this embodiment comprise purge passages having a diameter D.sub.1
(see FIG. 4) that is large enough, e.g., 3-4 mm, for providing a
sufficient volume of cooling air for purging hot gas from the seal
slot 44 and the seal member 46 and for creating a barrier of cool
air in the gap between adjacent inner shrouds 18.
[0029] Each cooling air exit passage 69 is in fluid communication
with the cooling air channel 48 and includes an exit opening 70
located between the outer surface 50 of the inner shroud 18 and the
seal slot 44 (see FIGS. 2 and 4). As shown in FIGS. 2 and 3, the
cooling air exit passages 69 according to this embodiment are
located along a downstream region 71 of the mating edges 40, 42,
and are preferably located toward the downstream portions 48B of
the cooling air channels 48, i.e., closer to the downstream edge 34
of the inner shroud 18 than to the upstream edge 32 thereof,
although the cooling air exit passages 69 could be located
elsewhere along the respective mating edges 40, 42 as desired.
Further, at least one of the cooling air exit passages 69
associated with each cooling air channel 48 is located near a
respective replenishing cooling air supply passage 68. As
illustrated in FIG. 5, the passages 69 may be angled toward the
downstream edge 34 of the inner shroud 18 to promote a flow having
a velocity component in the same direction as the hot working gas
H.sub.G. Moreover, the exit openings 70 of the passages 69 may be
axially offset with respect to exit openings 70 of adjacent inner
shrouds 18, as shown in FIG. 5.
[0030] During operation of the engine, the row 10 of vane segments
12, which, as noted above, is a first row 10 of vane segments 12
according to this embodiment, is exposed to the high temperature
hot working gas H.sub.G entering the turbine section 14 from the
combustor assembly. A region of the vane segment 12 along the inner
shroud 18 adjacent to and downstream from the trailing edge 28 is
believed to be particularly susceptible to exposure to high
temperature gases that may lead to elevated temperatures of the
surfaces of the inner shroud 18. In particular, the mating edges
40, 42 may be exposed to elevated temperatures in the region 71,
which may generally be defined as extending from a location at or
near an axial location of the trailing edge 28, as identified by
line 73 in FIG. 3, to the location of the downstream edge 34. The
replenishing cooling air supply passages 69, air exit passages 69
and exit openings 70 are preferably located in or near the region
71.
[0031] Cooling air, e.g., compressor discharge air, enters the
leading edge cooling air chamber 53 from the first cavity 55
through the impingement holes 57A formed in the impingement plate
57. The cooling air entering the chamber 53 through the impingement
holes 57A contacts and provides impingement cooling to the inner
surface 52 of the outer wall 51.
[0032] A first portion of the cooling air in the leading edge
cooling air chamber 53 passes into the cooling air channels 48
through the respective primary cooling air supply passages 58.
Another portion of the cooling air in the chamber 53 passes through
the film cooling holes 59 in the wall 51 and provides film cooling
for the outer surface 50 of the wall 51 as discussed above.
[0033] The cooling air flowing through the cooling air channels 48
provides convective cooling for the inner shroud 18 adjacent to the
mating edges 40, 42. Since the cooling air channels 48 are located
close to the mating edges 40, 42 and to the outer surface 50 of the
wall 51, the corners 56 of the inner shroud 18 are convectively
cooled by the cooling air flowing through the cooling air channels
48.
[0034] As a result of the cooling air in the cooling air channels
48 convectively cooling the inner shroud 18 adjacent to the mating
edges 40, 42 by removing heat from the inner shroud 18, the cooling
air flowing through the cooling air channels 48 is heated as it
flows downstream toward the downstream portions 48B of the cooling
air channels 48.
[0035] The second portion of cooling air, which is provided into
the cooling air channels 48 from the mid-chord cooling air chamber
60 via the replenishing cooling air supply passages 68, is added to
and cools the first portion of cooling air. Hence, the cooling air
that is available for convective cooling in the downstream portions
48B of the cooling air channels 48, i.e., for cooling the mating
edges 40, 42 and the corners 56, is cool enough to sufficiently
cool the inner shroud 18 adjacent to the downstream edge 34
thereof. That is, the cooling air exiting the cooling air exit
passages 69 through the exit openings 70 is cool enough, and
comprises an adequate volume, to sufficiently cool the seal member
46 and the mating edges 40, 42 and corners 56 of the inner shroud
18, and to provide a cool air barrier for the hot working gases
H.sub.G. Further, since the exit openings 70 of the respective
adjacent inner shrouds 18 are axially offset from one another, the
cooling air exiting the respective exit openings 70 provides a
substantially even distribution of cooling air between the mating
edges 40, 42 of the respective inner shrouds 18. It is noted that
any cooling air in the channels 48 that does not exit through the
cooling air exit passages 69 may exit the inner shroud 18 through
the outlets 49 of the channels 48.
[0036] Referring now to FIGS. 6-8, a vane segment 110 belonging to
a row (not shown in this embodiment) of vane segments 110 according
to another embodiment of the invention is shown. The vane segment
110 according to this embodiment belongs to a second row of vane
segments 110, also referred to as a row two vane segment. The vane
segment 110 illustrated in FIG. 6 includes two airfoils 112 and a
radially inner shroud 114 rigidly connected to a radially inner end
of each airfoil 112. The vane segment 110 also includes one or more
radially outer shrouds (not shown in this embodiment) rigidly
connected to radially outer ends of the airfoils 112. It is noted
that the vane segment 110 may be constructed with additional or
fewer airfoils 112.
[0037] The vane segment 110 includes a first generally axially
extending mating edge 116 extending between an upstream edge 118
and a downstream edge 120 of the inner shroud 114, and an opposing
second generally axially extending mating edge 122 extending
generally parallel to the first mating edge 116 between the
upstream edge 118 and the downstream edge 120 of the inner shroud
114. Each of the mating edges 116, 122 includes a generally axially
extending seal slot 124 adapted to receive an axially extending
seal member 126 (FIG. 7). The seal member 126 spans across a gap
formed between adjacent mating edges 116, 122 of adjacent vane
segments 110 to separate the hot working gas (discussed above with
reference to FIGS. 1-5) from a region of cool air located radially
inwardly from the inner shroud 114.
[0038] As shown in FIG. 6-8, respective primary cooling air
channels 130 extend within the inner shroud 114 generally axially
adjacent to and substantially parallel to each of the mating edges
116, 122. As illustrated in FIG. 6, in the embodiment shown the
channels 130 extend from upstream portions 130A thereof adjacent to
the upstream edge 118 of the inner shroud 114 to outlets 131
located at downstream portions 1308 thereof adjacent to the
downstream edge 120 of the inner shroud 114. As shown in FIG. 7,
each channel 130 is located radially between an outer surface 132
defined by an outer wall 134 of the inner shroud 114 and the seal
slot 124. Further, each channel 130 is preferably aligned
circumferentially with a respective seal slot 124, such that the
channels 130 are located on, or intersected by, respective radial
planes 136 passing radially through the corresponding seal slots
124, see FIG. 7. Preferably, the entirety of a cross-sectional area
of the cooling air channel 130, as viewed in the axial direction,
is located radially outwardly from, i.e., substantially overlies, a
respective seal slot 124, as shown in FIG. 7. The cooling air
channels 130 effect convective cooling of a corner 138 (FIG. 7) at
an intersection of the outer surface 132 of the outer wall 134 and
respective ones of the mating edges 116, 122 of the inner shroud
114, as will be described herein. It is noted that the spacing
between the outer surface 132 of the outer wall 134 and the seal
slots 124 is large enough to accommodate the respective cooling air
channels 130 therebetween while still maintaining the structural
rigidity of the inner shroud 114 at the respective mating edges
116, 122.
[0039] Referring now to FIGS. 6 and 7, an inner surface 140 (FIG.
7) of the outer wall 134 defines a radially outer boundary for a
mid-chord cooling air chamber 142 associated with the inner shroud
114. The chamber 142 receives cooling air, i.e., compressor
discharge air, for cooling the inner shroud 114. In the embodiment
shown, cooling air, i.e., compressor discharge air, may enter the
chamber 142 from internal cooling passageways 144 (FIG. 6) that
extend through the airfoils 112. It is noted that, while film
cooling holes (such as the film cooling holes 64 described above
with reference to FIGS. 1 and 3) are not illustrated in the outer
wall 134 of the inner shroud 114 according to this embodiment, such
film cooling holes could be provided in the outer wall 134 if film
cooling of the outer surface 132 of the outer wall 134 is
desired.
[0040] As shown in FIG. 6, cooling air is provided from the
mid-chord cooling air chamber 142 via passageways 148 into a
leading edge cooling air chamber 150 located adjacent to the
upstream edge 118 of the inner shroud 114. The cooling air in the
leading edge cooling air chamber 150 provides convective cooling to
the inner shroud 114 adjacent to the upstream edge 118 thereof. It
is noted that the cooling air chamber 150 may be formed by one or
more sub-chambers, two such sub-chambers define the leading edge
cooling air chamber 150 in the embodiment shown in FIG. 6.
[0041] A first portion of cooling air is provided into the cooling
air channels 130 via respective primary cooling air supply passages
152 that extend from the leading edge cooling air chamber 150 to
the upstream portions 130A of the respective cooling air channels
130, see FIG. 6. The passages 152 may include a component in the
radial direction, as the channels 130 in the embodiment shown may
be located in the outer wall 134 of the inner shroud 114 radially
outwardly from the leading edge cooling air chamber 150.
[0042] Referring to FIGS. 6-8, replenishing cooling air supply
passages 156 provide fluid communication between the mid-chord
cooling air chamber 142 and the respective cooling air channels 130
between the upstream and downstream portions 130A and 130B thereof,
i.e., between the primary cooling air supply passages 152 and the
downstream edge 120 of the inner shroud 114. The replenishing
cooling air supply passages 156 supply a second portion of cooling
air from the mid-chord cooling air chamber 142 to the channels 130.
As shown in FIG. 7, the passages 156 may include a component in the
radial direction, as the channels 130 in the embodiment shown may
be located in the outer wall 134 of the inner shroud 114 radially
outwardly from the chamber 142. Further, the passages 156 may be
angled toward the downstream edge 120 of the inner shroud 114 to
promote a flow toward the downstream portions 130B of the cooling
air channels 130, as shown in FIGS. 6 and 8.
[0043] As shown in FIGS. 6-8, a plurality of cooling air exit
passages 158 extend from each cooling air channel 130 to the
respective axial mating edges 116, 122. The cooling air exit
passages 158 according to this embodiment comprise impingement
passages having a diameter D.sub.2 (see FIG. 7) that is large
enough, e.g., 1-2 mm, for providing impingement cooling to a mating
edge 116, 122 of an adjacent vane segment 110, see FIGS. 7 and
8.
[0044] Each cooling air exit passage 158 is in fluid communication
with a respective cooling air channel 130 and includes an exit
opening 160 located between the outer surface 132 of the inner
shroud 114 and the seal slot 124 (see FIG. 7). As shown in FIG. 6,
the cooling air exit passages 158 according to this embodiment are
preferably located near the upstream portions 130A of the channels
130, i.e., from the upstream edge 118 of the inner shroud 114 to a
location about 1/3 of the distance between the upstream and
downstream edges 118 and 120 of the inner shroud 114, although it
is understood that the cooling air exit passages 158 could be
located elsewhere at the mating edges 116, 122 as desired. As
illustrated in FIGS. 6 and 8, the passages 158 may be angled toward
the downstream edge 120 of the inner shroud 114 to promote a flow
including a velocity component in the same direction as the hot
working gas. Moreover, the exit openings 160 of the passages 158
may be axially offset with respect to the exit openings 160 of
passages 158 of the adjacent inner shroud 114, as shown in FIG.
8.
[0045] During operation of the engine, the row of vane segments
110, which, as noted above, is a second row of vane segments 110
according to this embodiment, is exposed to high temperature hot
working gas entering the turbine section from the combustor
assembly as described above with reference to FIGS. 1-5. A region
of the vane segment 110 along the inner shroud 114 adjacent to and
downstream from the upstream edge 118, i.e., in a region extending
approximately 1/3 of the distance between the upstream and
downstream edges 118 and 120, is particularly susceptible to high
temperatures.
[0046] Cooling air, e.g., compressor discharge air, enters the
mid-chord cooling air chamber 142 from the internal cooling
passageways 144 that extend through the airfoils 112. The cooling
air entering the mid-chord cooling air chamber 142 provides
convective cooling to the inner shroud 114 around the chamber
142.
[0047] As noted above, cooling air is provided from the mid-chord
chamber cooling air chamber 142 into the leading edge cooling air
chamber 150 through the passageways 148. The cooling air in the
leading edge cooling air chamber 150 provides convective cooling to
the inner shroud 114 adjacent to the upstream edge 118 thereof. A
first portion of the cooling air in the leading edge cooling air
chamber 150 passes into the cooling air channels 130 through the
respective primary cooling air supply passages 152.
[0048] The cooling air flowing through the cooling air channels 130
provides convective cooling for the inner shroud 114 adjacent to
the mating edges 116, 122. Since the cooling air channels 130 are
located close to the mating edges 116, 122 and to the outer surface
132 of the wall 134, the corners 138 of the inner shroud 114 are
convectively cooled by the cooling air flowing through the cooling
air channels 130. As a result of the cooling air in the channels
130 convectively cooling the inner shroud 114 adjacent to the
mating edges 116, 122, the cooling air flowing through the cooling
air channels 130 is heated as it flows downstream toward the
downstream portions 130B of the cooling air channels 130. Some of
the first portion of cooling air exits the channels 130 through the
cooling air exit passages 158 near the upstream portions 130A of
the channels 130.
[0049] The second portion of cooling air, which is provided to the
cooling air channels 130 directly from the mid-chord cooling air
chamber 142 via the replenishing cooling air supply passages 156,
is added to and cools the remaining portion of the first portion of
cooling air. Hence, the cooling air that is available for cooling
within the cooling air channels 130 is cool enough to sufficiently
cool the inner shroud 114, i.e., the corners 138 and the mating
edges 116, 122, and also to cool the seal members 126.
Additionally, since the exit openings 160 of the adjacent inner
shrouds 114 are axially offset from one another, the cooling air
exiting the respective exit openings 160 provides a substantially
even distribution of cooling air between the mating edges 116, 122
of the respective inner shrouds 114 for providing impingement
cooling for the opposed mating edges 116, 122. It is noted that any
cooling air in the channels 130 that does not exit through the
cooling air exit passages 158 may exit the inner shroud 114 through
the outlets 131 of the channels 130.
[0050] While particular embodiments of the present invention have
been illustrated and described, it would be obvious to those
skilled in the art that various other changes and modifications can
be made without departing from the spirit and scope of the
invention. It is therefore intended to cover in the appended claims
all such changes and modifications that are within the scope of
this invention.
* * * * *