U.S. patent application number 13/924178 was filed with the patent office on 2014-12-25 for nozzle film cooling with alternating compound angles.
The applicant listed for this patent is Solar Turbines Incorporated. Invention is credited to Hee Koo Moon, Juan Yin, Luzeng Zhang.
Application Number | 20140377054 13/924178 |
Document ID | / |
Family ID | 52105298 |
Filed Date | 2014-12-25 |
United States Patent
Application |
20140377054 |
Kind Code |
A1 |
Zhang; Luzeng ; et
al. |
December 25, 2014 |
NOZZLE FILM COOLING WITH ALTERNATING COMPOUND ANGLES
Abstract
A nozzle segment for a nozzle ring of a gas turbine engine is
disclosed. The nozzle segment includes a first endwall, a second
endwall, and an airfoil extending between the first endwall and the
second endwall. The airfoil includes a multiple groups of cooling
apertures spaced apart and alternating in directionality such that
a first grouping of cooling apertures is angled toward the first
endwall, a second grouping of cooling apertures is angled toward
the second endwall and spaced apart from the first grouping of
cooling apertures, and a third grouping of cooling apertures are
angled toward the first endwall and spaced apart from the second
grouping of cooling apertures.
Inventors: |
Zhang; Luzeng; (San Diego,
CA) ; Yin; Juan; (San Diego, CA) ; Moon; Hee
Koo; (San Diego, CA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Solar Turbines Incorporated |
San Diego |
CA |
US |
|
|
Family ID: |
52105298 |
Appl. No.: |
13/924178 |
Filed: |
June 21, 2013 |
Current U.S.
Class: |
415/115 |
Current CPC
Class: |
F01D 5/186 20130101;
F01D 9/065 20130101 |
Class at
Publication: |
415/115 |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Claims
1. A nozzle segment for a nozzle ring of a gas turbine engine, the
nozzle segment comprising: a first endwall; a second endwall; and
an airfoil extending between the first endwall and the second
endwall, the airfoil including a leading edge extending radially
from the first endwall to the second endwall, a trailing edge
extending radially from the first endwall to the second endwall
axially distal to the leading edge, a pressure side wall extending
from the leading edge to the trailing edge, a suction side wall
extending from the leading edge to the trailing edge, a plurality
of showerhead cooling apertures spanning along the leading edge, a
plurality of forward cooling apertures grouped together in the
pressure side wall proximate the plurality of showerhead cooling
apertures, and a plurality of intermediate cooling apertures
grouped together in the pressure side wall between the trailing
edge and the plurality of forward cooling apertures, the plurality
of showerhead cooling apertures, the plurality of forward cooling
apertures, and the plurality of intermediate cooling apertures
alternating in directionality such that the plurality of showerhead
cooling apertures are angled toward the first endwall, the
plurality of forward cooling apertures are angled toward the second
endwall, and the plurality of intermediate cooling apertures are
angled toward the first endwall.
2. The nozzle segment of claim 1, wherein each showerhead cooling
aperture is angled toward the first endwall as each showerhead
cooling aperture extends through a wall of the airfoil, each
forward cooling aperture includes a forward compound angle from
fifteen to forty-five degrees towards the second endwall relative
to a flow direction of air through the nozzle segment during
operation, and each intermediate cooling aperture including an
intermediate compound angle from fifteen to forty-five degrees
towards the first endwall relative to a flow direction of air
through the nozzle segment during operation.
3. The nozzle segment of claim 2, wherein the airfoil further
includes a plurality of aft cooling apertures grouped together in
the pressure side wall proximate the trailing edge, each aft
cooling aperture including an aft compound angle from fifteen to
forty-five degrees towards the second endwall relative to a flow
direction of air through the nozzle segment during operation.
4. The nozzle segment of claim 1, wherein the plurality of forward
cooling apertures are arranged in a single column and the plurality
of intermediate cooling apertures are arranged in a single
column.
5. The nozzle segment of claim 1, wherein the first endwall is a
lower endwall and the second endwall is an upper endwall located
radially outward from the lower endwall.
6. The nozzle segment of claim 1, wherein each forward cooling
aperture is spaced apart from an adjacent forward cooling aperture
from 3 to 4 pitch over diameter and each intermediate cooling
aperture is spaced apart from an adjacent intermediate cooling
aperture from 3 to 4 pitch over diameter.
7. The nozzle segment of claim 1, wherein the plurality of
showerhead cooling apertures is configured to direct air to film
cool the leading edge and cool the first endwall, the plurality of
forward cooling apertures is configured to direct air to film cool
a pressure side surface of the pressure side wall and cool the
second endwall, and the plurality of intermediate cooling apertures
is configured to direct air to film cool the pressure side surface
and cool the first endwall.
8. A gas turbine engine including the nozzle segment of claim 1,
wherein the nozzle segment is located in a first stage turbine
nozzle of the gas turbine engine.
9. A nozzle segment for a nozzle ring of a gas turbine engine, the
nozzle segment comprising: a first endwall; a second endwall; an
airfoil extending between the first endwall and the second endwall,
the airfoil including a leading edge extending radially from the
first endwall to the second endwall, a trailing edge extending
radially from the first endwall to the second endwall, a pressure
side wall extending from the leading edge to the trailing edge, the
pressure side wall including a pressure side surface with a concave
shape, the pressure side surface being the outer surface of the
pressure side wall, a suction side wall extending from the leading
edge to the trailing edge, a cooling cavity located between the
leading edge, the trailing edge, the pressure side wall, and the
suction side wall, a plurality of showerhead cooling apertures
spanning along the leading edge, each showerhead cooling aperture
including a showerhead inlet end adjacent the cooling cavity and a
showerhead outlet end at the outer surface of the leading edge, the
showerhead inlet end being radially closer to the second endwall
than the showerhead outlet end and the showerhead outlet end being
radially closer to the first endwall than the showerhead inlet end,
a plurality of forward cooling apertures in the pressure side wall
grouped together and spaced apart from the plurality of showerhead
cooling apertures at least 1/8 the length of the pressure side
wall, each forward cooling aperture including a forward inlet end
adjacent the cooling cavity and a forward outlet end adjacent the
pressure side surface, the forward inlet end being radially closer
to the first endwall and axially closer to the leading edge than
the forward outlet end, and the forward outlet end being radially
closer to the second endwall and axially closer to the trailing
edge than the forward inlet end, and a plurality of intermediate
cooling apertures in the pressure side wall grouped together and
spaced apart from the plurality of forward cooling apertures at
least 1/8 the length of the pressure side wall, each intermediate
cooling aperture including an intermediate inlet end adjacent the
cooling cavity and an intermediate outlet end adjacent the pressure
side surface, the intermediate inlet end being radially closer to
the second endwall and axially closer to the leading edge than the
intermediate outlet end, and the intermediate outlet end being
radially closer to the first endwall and axially closer to the
trailing edge than the intermediate inlet end.
10. The nozzle segment of claim 9, wherein the airfoil further
includes a plurality of aft cooling apertures in the pressure side
wall grouped together and spaced apart from the plurality of
intermediate cooling apertures at least 1/8 the length of the
pressure side wall, each aft cooling aperture including an aft
inlet end adjacent the cooling cavity and an aft outlet end
adjacent the pressure side surface, the aft inlet end being
radially closer to the first endwall and axially closer to the
leading edge than the aft outlet end, and the aft outlet end being
radially closer to the second endwall and axially closer to the
trailing edge than the aft inlet end.
11. The nozzle segment of claim 9, wherein the first endwall is a
lower endwall of a lower shroud and the second endwall is an upper
endwall of an upper shroud, the lower endwall being located
radially inward from the upper endwall.
12. The nozzle segment of claim 9, wherein each forward cooling
aperture is spaced apart from an adjacent forward cooling aperture
from 3 to 4 pitch over diameter and each intermediate cooling
aperture is spaced apart from an adjacent intermediate cooling
aperture from 3 to 4 pitch over diameter.
13. The nozzle segment of claim 10, wherein each forward cooling
aperture is spaced apart from an adjacent forward cooling aperture
from 3 to 4 pitch over diameter, each intermediate cooling aperture
is spaced apart from an adjacent intermediate cooling aperture from
3 to 4 pitch over diameter, and each aft cooling aperture is spaced
apart from an adjacent aft cooling aperture from 3 to 4 pitch over
diameter.
14. A nozzle segment for a nozzle ring of a gas turbine engine, the
nozzle segment comprising: an upper shroud including an upper
endwall, the upper endwall being the shape of a sector of a toroid;
a lower shroud including a lower endwall located radially inward
from the upper endwall, the lower endwall being the shape of a
sector of a toroid; an airfoil extending between the upper endwall
and the lower endwall, the airfoil including a leading edge
extending radially from the upper endwall to the lower endwall, a
trailing edge extending radially from the upper endwall to the
lower endwall axially distal to the leading edge, a pressure side
wall extending from the leading edge to the trailing edge, the
pressure side wall including a concave shape, a suction side wall
extending from the leading edge to the trailing edge, the suction
side wall including a convex shape, a cooling cavity located
between the leading edge, the trailing edge, the pressure side
wall, and the suction side wall, a plurality of showerhead cooling
apertures arranged in four to seven columns spanning along the
leading edge, each showerhead cooling aperture being angled toward
the lower endwall as each showerhead cooling aperture extends
through a wall of the airfoil from the cooling cavity, a plurality
of forward cooling apertures arranged in a column extending
radially between the upper endwall and the lower endwall and
located in the third of the pressure side wall adjacent the leading
edge, each forward cooling aperture extending through the pressure
side wall from the cooling cavity and including a forward compound
angle from fifteen to forty-five degrees towards the upper endwall
and the trailing edge relative to a reference line in the plane of
a pressure side surface of the pressure side wall, the reference
line being defined as an intersection between the pressure side
surface and a plane perpendicular to a radial extending from an
axis of the upper shroud along the pressure side surface, a
plurality of intermediate cooling apertures arranged in a column
extending radially between the upper endwall and the lower endwall
and located in the middle third of the pressure side wall between
the leading edge and the trailing edge, each intermediate cooling
aperture extending through the pressure side wall from the cooling
cavity and including an intermediate compound angle from fifteen to
forty-five degrees towards the lower endwall and the trailing edge
relative to the reference line, and a plurality of aft cooling
apertures arranged in a column extending radially between the upper
endwall and the lower endwall and located in the third of the
pressure side wall adjacent the trailing edge, each aft cooling
aperture extending through the pressure side wall from the cooling
cavity and including an aft compound angle from fifteen to
forty-five degrees towards the upper endwall and the trailing edge
relative to the reference line.
15. The nozzle segment of claim 14, further comprising: a second
airfoil extending between the upper endwall and the lower endwall,
the second airfoil including a second leading edge extending
radially from the upper endwall to the lower endwall, a second
trailing edge extending radially from the upper endwall to the
lower endwall distal to the second leading edge, a second pressure
side wall extending from the second leading edge to the second
trailing edge, the second pressure side wall including a second
concave shape, a second suction side wall extending from the second
leading edge to the second trailing edge, the second suction side
wall including a second convex shape, a second cooling cavity
located between the second leading edge, the second trailing edge,
the second pressure side wall, and the second suction side wall, a
plurality of second showerhead cooling apertures arranged in four
to seven columns spanning along the second leading edge, each
second showerhead cooling aperture being angled toward the lower
endwall as each second showerhead cooling aperture extends through
a second wall of the second airfoil from one of the second cooling
cavity, a plurality of second forward cooling apertures arranged in
a column extending radially between the upper endwall and the lower
endwall and located in the third of the second pressure side wall
adjacent the second leading edge, each second forward cooling
aperture extending through the second pressure side wall from the
second cooling cavity and including a second forward compound angle
from fifteen to forty-five degrees towards the upper endwall and
the second trailing edge relative to a second reference line in the
plane of a second pressure side surface of the second pressure side
wall, the second reference line being defined as an intersection
between the second pressure side surface and a second plane
perpendicular to a radial extending from the axis of the upper
shroud along the second pressure side surface, a plurality of
second intermediate cooling apertures arranged in a column
extending radially between the upper endwall and the lower endwall
and located in the middle third of the second pressure side wall
between the second leading edge and the second trailing edge, each
second intermediate cooling aperture extending through the second
pressure side wall from the second cooling cavity and including a
second intermediate compound angle from fifteen to forty-five
degrees towards the lower endwall and the second trailing edge
relative to the second reference line, and a plurality of second
aft cooling apertures arranged in a column extending radially
between the upper endwall and the lower endwall and located in the
third of the second pressure side wall adjacent the second trailing
edge, each second aft cooling aperture extending through the second
pressure side wall from the second cooling cavity and including a
second aft compound angle from fifteen to forty-five degrees
towards the upper endwall and the second trailing edge relative to
the second reference line.
16. The nozzle segment of claim 14, wherein each forward cooling
aperture is spaced apart from an adjacent forward cooling aperture
from 3 to 4 pitch over diameter, each intermediate cooling aperture
is spaced apart from an adjacent intermediate cooling aperture from
3 to 4 pitch over diameter, and each aft cooling aperture is spaced
apart from an adjacent aft cooling aperture from 3 to 4 pitch over
diameter.
17. The nozzle segment of claim 14, wherein each forward cooling
aperture is spaced apart from an adjacent forward cooling aperture
at 3.5 pitch over diameter, each intermediate cooling aperture is
spaced apart from an adjacent intermediate cooling aperture at 3.5
pitch over diameter, and each aft cooling aperture is spaced apart
from an adjacent aft cooling aperture at 3.5 pitch over
diameter.
18. The nozzle segment of claim 14, wherein the plurality of
forward cooling apertures are spaced apart from the plurality of
showerhead cooling apertures from 1/8 to 1/4 the length of the
pressure side wall, the plurality of intermediate cooling apertures
are spaced apart from the plurality of forward cooling apertures
from 1/4 to 3/8 the length of the pressure side wall, and the
plurality of aft cooling apertures are spaced apart from the
plurality of intermediate cooling apertures from 1/4 to 3/8 of the
length of the pressure side wall.
19. The nozzle segment of claim 14, wherein the forward compound
angle is within a predetermined tolerance of thirty degrees, the
intermediate compound angle is within a predetermined tolerance of
thirty degrees, and the aft compound angle is within a
predetermined tolerance of thirty degrees.
20. A gas turbine engine including the nozzle segment of claim 14,
wherein the nozzle segment is located in a first stage turbine
nozzle of the gas turbine engine.
Description
TECHNICAL FIELD
[0001] The present disclosure generally pertains to gas turbine
engines, and is more particularly directed toward nozzle segments
including film cooling holes with alternating compound angles.
BACKGROUND
[0002] Gas turbine engines include compressor, combustor, and
turbine sections. The turbine section is subject to high
temperatures. In particular, the first stages of the turbine
section are subject to such high temperatures that the first stages
are often cooled with air directed from the compressor and into,
inter alia, the nozzle segments and turbine blades.
[0003] A portion of the air directed into the nozzle segments may
be directed through the walls of the nozzle segment airfoils and
along the pressure side surface of the walls to film cool the
walls. U.S. Pat. No. 7,377,743 to D. Flodman discloses a turbine
nozzle that includes a mid vane mounted between a pair of end vanes
in outer and inner bands. The mid vane includes a first pattern of
film cooling holes configured to discharge more cooling air than
each of the two end vanes having respective second patterns of film
cooling holes.
[0004] The present disclosure is directed toward overcoming one or
more of the problems discovered by the inventors or that is known
in the art.
SUMMARY OF THE DISCLOSURE
[0005] A nozzle segment for a nozzle ring of a gas turbine engine
is disclosed. The nozzle segment includes a first endwall, a second
endwall, and an airfoil extending between the first endwall and the
second endwall. The airfoil includes a leading edge, a trailing
edge, a pressure side wall, and a suction side wall. The leading
edge extends radially from the first endwall to the second endwall.
The trailing edge extends radially from the first endwall to the
second endwall axially distal to the leading edge. The pressure
side wall extends from the leading edge to the trailing edge. The
suction side wall also extends from the leading edge to the
trailing edge. The airfoil also includes a plurality of showerhead
cooling apertures, a plurality of forward cooling apertures, and a
plurality of intermediate cooling apertures. The plurality of
showerhead cooling apertures span along the leading edge. The
plurality of forward cooling apertures are grouped together
proximate the plurality of showerhead cooling apertures. The
plurality of intermediate cooling apertures are grouped together in
the pressure side wall between the trailing edge and the plurality
of forward cooling apertures. The plurality of showerhead cooling
apertures, the plurality of forward cooling apertures, and the
plurality of intermediate cooling apertures alternate in
directionality such that the plurality of showerhead cooling
apertures are angled toward the first endwall, the plurality of
forward cooling apertures are angled toward the second endwall, and
the plurality of intermediate cooling apertures are angled toward
the first endwall.
BRIEF DESCRIPTION OF THE DRAWINGS
[0006] FIG. 1 is a schematic illustration of an exemplary gas
turbine engine.
[0007] FIG. 2 is a perspective view of a nozzle segment for the gas
turbine engine of FIG. 1.
[0008] FIG. 3 is a cross-sectional view of a portion of the nozzle
segment of FIG. 2 showing the showerhead cooling apertures.
[0009] FIG. 4 is a detailed view of the forward cooling apertures
of FIG. 2.
[0010] FIG. 5 is a detailed view of the intermediate cooling
apertures of FIG. 2.
[0011] FIG. 6 is a detailed view of the aft cooling apertures of
FIG. 2.
[0012] FIG. 7 is a cross-section of the airfoil of FIG. 2.
DETAILED DESCRIPTION
[0013] The systems and methods disclosed herein include a nozzle
segment for a nozzle ring of a gas turbine engine. In embodiments,
the nozzle segment includes an upper endwall, an inner endwall, and
one or more airfoils there between. Each airfoil includes spaced
apart groups of cooling apertures through the pressure side wall of
the airfoil. One group is angled toward the lower endwall and the
next group is angled towards the upper endwall in an alternating
pattern for subsequent groups of cooling holes. Alternating the
directionality of the groups of cooling apertures towards the lower
endwall and the upper endwall may reduce the temperatures of the
lower endwall and the upper endwall, and may reduce the amount of
cooling air needed to effectively cool the nozzle segment.
[0014] FIG. 1 is a schematic illustration of an exemplary gas
turbine engine 100. Some of the surfaces have been left out or
exaggerated (here and in other figures) for clarity and ease of
explanation. Also, the disclosure may reference a forward and an
aft direction. Generally, all references to "forward" and "aft" are
associated with the flow direction of primary air (i.e., air used
in the combustion process), unless specified otherwise. For
example, forward is "upstream" relative to primary air flow, and
aft is "downstream" relative to primary air flow.
[0015] In addition, the disclosure may generally reference a center
axis 95 of rotation of the gas turbine engine, which may be
generally defined by the longitudinal axis of its shaft 120
(supported by a plurality of bearing assemblies 150). The center
axis 95 may be common to or shared with various other engine
concentric components. All references to radial, axial, and
circumferential directions and measures refer to center axis 95,
unless specified otherwise, and terms such as "inner" and "outer"
generally indicate a lesser or greater radial distance from,
wherein a radial 96 may be in any direction perpendicular and
radiating outward from center axis 95.
[0016] A gas turbine engine 100 includes an inlet 110, a shaft 120,
a compressor 200, a combustor 300, a turbine 400, an exhaust 500,
and a power output coupling 600. The gas turbine engine 100 may
have a single shaft or a dual shaft configuration.
[0017] The compressor 200 includes a compressor rotor assembly 210,
compressor stationary vanes (stators) 250, and inlet guide vanes
255. The compressor rotor assembly 210 mechanically couples to
shaft 120. As illustrated, the compressor rotor assembly 210 is an
axial flow rotor assembly. The compressor rotor assembly 210
includes one or more compressor disk assemblies 220. Each
compressor disk assembly 220 includes a compressor rotor disk that
is circumferentially populated with compressor rotor blades.
Stators 250 axially follow each of the compressor disk assemblies
220. Each compressor disk assembly 220 paired with the adjacent
stators 250 that follow the compressor disk assembly 220 is
considered a compressor stage. Compressor 200 includes multiple
compressor stages. Inlet guide vanes 255 axially precede the
compressor stages.
[0018] The combustor 300 includes one or more fuel injectors 310
and includes one or more combustion chambers 390.
[0019] The turbine 400 includes a turbine rotor assembly 410 and
turbine nozzles 450. The turbine rotor assembly 410 mechanically
couples to the shaft 120. As illustrated, the turbine rotor
assembly 410 is an axial flow rotor assembly. The turbine rotor
assembly 410 includes one or more turbine disk assemblies 420. Each
turbine disk assembly 420 includes a turbine disk that is
circumferentially populated with turbine blades. A turbine nozzle
450 or nozzle ring axially precedes each of the turbine disk
assemblies 420. Each turbine nozzle 450 includes multiple nozzle
segments 451 grouped together to form a ring. Each turbine disk
assembly 420 paired with the adjacent turbine nozzle 450 that
precede the turbine disk assembly 420 is considered a turbine
stage. Turbine 400 includes multiple turbine stages.
[0020] The turbine 400 may also include a turbine housing 430 and
turbine diaphragms 440. Turbine housing 430 may be located radially
outward from turbine rotor assembly 410 and turbine nozzles 450.
Turbine housing 430 may include one or more cylindrical shapes.
Each nozzle segment 451 may be configured to attach, couple to, or
hang from turbine housing 430. Each turbine diaphragm 440 may
axially precede each turbine disk assembly 420 and may be adjacent
a turbine disk. Each turbine diaphragm 440 may also be located
radially inward from a turbine nozzle 450. Each nozzle segment 451
may also be configured to attach or couple to a turbine diaphragm
440.
[0021] The exhaust 500 includes an exhaust diffuser 520 and an
exhaust collector 550. The power output coupling 600 may be located
at an end of shaft 120.
[0022] FIG. 2 is a perspective view of a nozzle segment 451 for the
gas turbine engine 100 of FIG. 1. Nozzle segment 451 includes upper
shroud 452, lower shroud 456, airfoil 460, and second airfoil 470.
In other embodiments, nozzle segment 451 can include more or fewer
airfoils. Upper shroud 452 may be located adjacent and radially
inward from turbine housing 430 when nozzle segment 451 is
installed in gas turbine engine 100. Upper shroud 452 includes
upper endwall 453. Upper endwall 453 may be a portion or a sector
of an annular shape, such as a sector of a toroid or a sector of a
hollow cylinder. The toroidal shape may be defined by a
cross-section with an inner edge including a convex shape. Multiple
upper endwalls 453 are arranged to form the annular shape and to
define the radially outer surface of the flow path through a
turbine nozzle 450. Upper endwall 453 may be coaxial to center axis
95 when installed in the gas turbine engine 100.
[0023] Upper shroud 452 may also include upper forward rail 454 and
upper aft rail 455. Upper forward rail 454 extends radially outward
from upper endwall 453. In the embodiment illustrated in FIG. 2,
upper forward rail 454 extends from upper endwall 453 at an axial
end of upper endwall 453. In other embodiments, upper forward rail
454 extends from upper endwall 453 near or adjacent to an axial end
of upper endwall 453. Upper forward rail 454 may include a lip,
protrusion or other features that may be used to secure nozzle
segment 451 to turbine housing 430.
[0024] Upper aft rail 455 may also extend radially outward from
upper endwall 453. In the embodiment illustrated in FIG. 2, upper
aft rail 455 is `L` shaped, with a first portion extending radially
outward from the axial end of upper endwall 453 opposite the
location of upper forward rail 454, and a second portion extending
in the direction opposite the location of upper forward rail 454
extending axially beyond upper endwall 453. In other embodiments,
upper aft rail 455 includes other shapes and may be located near or
adjacent to the axial end of upper endwall 453 opposite the
location of upper forward rail 454. Upper aft rail 455 may also
include other features that may be used to secure nozzle segment
451 to turbine housing 430.
[0025] Lower shroud 456 is located radially inward from upper
shroud 452. Lower shroud 456 may also be located adjacent and
radially outward from turbine diaphragm 440 when nozzle segment 451
is installed in gas turbine engine 100. Lower shroud 456 includes
lower endwall 457. Lower endwall 457 may be a portion or a sector
of an annular shape, such as a toroid or a hollow cylinder. The
toroidal shape may be defined by a cross-section with an outer edge
including a convex shape. Multiple lower endwalls 457 are arranged
to form the annular shape and to define the radially inner surface
of the flow path through a turbine nozzle 450. Lower endwall 457
may be coaxial to upper endwall 453 and center axis 95 when
installed in the gas turbine engine 100.
[0026] Lower shroud 456 may also include lower forward rail 458 and
lower aft rail 459. Lower forward rail 458 extends radially inward
from lower endwall 457. In the embodiment illustrated in FIG. 2,
lower forward rail 458 extends from lower endwall 457 at an axial
end of lower endwall 457. In other embodiments, lower forward rail
458 extends from lower endwall 457 near or adjacent to an axial end
of lower endwall 457. Lower forward rail 458 may include a lip,
protrusion or other features that may be used to secure nozzle
segment 451 to turbine diaphragm 440.
[0027] Lower aft rail 459 may also extend radially inward from
lower endwall 457. In the embodiment illustrated in FIG. 2, lower
aft rail 459 extends from lower endwall 457 near or adjacent to the
axial end of lower endwall 457 opposite the location of lower
forward rail 458. In other embodiments, lower aft rail 459 extends
from the axial end of lower endwall 457 opposite the location of
lower forward rail 458. Lower aft rail 459 may also include a lip,
protrusion or other features that may be used to secure nozzle
segment 451 to turbine diaphragm 440.
[0028] Airfoil 460 extends between upper endwall 453 and lower
endwall 457. Airfoil 460 includes leading edge 461, trailing edge
462, pressure side wall 463, and suction side wall 464. Leading
edge 461 extends from upper endwall 453 adjacent an axial end of
upper endwall 453 to lower endwall 457 adjacent an axial end of
lower endwall 457. Leading edge 461 may be located near upper
forward rail 454 and lower forward rail 458. Trailing edge 462
extends from upper endwall 453 distal to leading edge 461, adjacent
the axial end of upper endwall 453 opposite the location of leading
edge 461 and from lower endwall 457 adjacent the axial end of upper
endwall 453 opposite or axial distal to the location of leading
edge 461. When nozzle segment 451 is installed in gas turbine
engine 100, leading edge 461, upper forward rail 454, and lower
forward rail 458 may be located axially forward and upstream of
trailing edge 462, upper aft rail 455, and lower aft rail 459.
Leading edge 461 may be the point at the upstream end of airfoil
460 with the maximum curvature and trailing edge 462 may be the
point at the downstream end of airfoil 460 with maximum curvature.
In the embodiment illustrated in FIG. 1, nozzle segment 451 is part
of the first stage turbine nozzle adjacent combustion chamber 390.
In other embodiments, nozzle segment 451 is located within a
turbine nozzle 450 of another stage.
[0029] Pressure side wall 463 may span from leading edge 461 to
trailing edge 462 between upper endwall 453 and lower endwall 457.
Pressure side wall 463 may include a concave shape. Pressure side
wall 463 may also include a pressure side surface 469, the outer
surface of pressure side wall 463, with a concave shape. Suction
side wall 464 may also span from leading edge 461 to trailing edge
462 between upper endwall 453 and lower endwall 457. Suction side
wall 464 may include a convex shape. Leading edge 461, trailing
edge 462, pressure side wall 463 and suction side wall 464 may form
a cooling cavity 485 (illustrated in FIGS. 3 and 6) or cooling
cavities there between. Upper endwall 453, lower endwall 457, or
both may include a hole, holes, or pathways for cooling air (not
shown) to enter the cooling cavity 485.
[0030] Airfoil 460 may also include multiple groupings of film
cooling holes or apertures. Each cooling hole or aperture may be a
channel extending through a wall of the airfoil, such as the
pressure side wall 463. In the embodiment illustrated in FIG. 2,
airfoil 460 includes showerhead cooling apertures 465, forward
cooling apertures 466, aft cooling apertures 467, and intermediate
cooling apertures 468. Showerhead cooling apertures 465 are located
at leading edge 461 and may be grouped together along leading edge
461. Showerhead cooling apertures 465 may be arranged in columns.
In the embodiment shown in FIG. 2, showerhead cooling apertures 465
are arranged in six columns, each column extending in the radial
direction between upper endwall 453 and lower endwall 457. In other
embodiments, showerhead cooling apertures 465 may be arranged in
four to seven columns or may be arranged in other configurations.
The portions of pressure side wall 463 and suction side wall 464
adjacent leading edge 461 may include showerhead cooling apertures
465.
[0031] Forward cooling apertures 466 may be grouped together and
located within the third of pressure side wall 463 that is adjacent
leading edge 461. Forward cooling apertures 466 may be proximate
showerhead cooling apertures 465. In embodiments, forward cooling
apertures 466 are located from 1/8 to 1/4 of the length of pressure
side wall 463 from showerhead cooling apertures 465. In other
embodiments, forward cooling apertures 466 are located 1/6 of the
length of pressure side wall 463 from showerhead cooling apertures
465. In yet other embodiments, forward cooling apertures 466 are
located at least 1/8 of the length of pressure side wall 463 from
showerhead cooling apertures 465. Forward cooling apertures 466 may
be grouped together between upper endwall 453 and lower endwall
457. In the embodiment illustrated in FIG. 2, forward cooling
apertures 466 are arranged in a single radial column and spaced
apart radially at 3.5 pitch over diameter, the distance between the
centers of adjacent apertures over the diameter of the apertures.
In other embodiments, forward cooling apertures 466 are spaced
apart radially from 3 to 4 pitch over diameter. Forward cooling
apertures 466 may overlap with an adjacent forward cooling aperture
466 rather than align in the radial direction along the surface of
pressure side wall 463.
[0032] Aft cooling apertures 467 may be grouped together and
located within the third of pressure side wall 463 that is adjacent
to trailing edge 462. Aft cooling apertures 467 may be proximate
trailing edge 462. In embodiments, aft cooling apertures are
located from 1/8 to 1/4 of the length of pressure side wall 463
from trailing edge 462. In other embodiments, aft cooling apertures
467 are located 1/6 of the length of pressure side wall 463 from
trailing edge 462. In yet other embodiments, aft cooling apertures
467 are located at least 1/8 of the length of pressure side wall
463 from trailing edge 462. Aft cooling apertures 467 may be
arranged radially between upper endwall 453 and lower endwall 457.
In the embodiment illustrated in FIG. 2, aft cooling apertures 467
are arranged in a single radial column and spaced apart radially at
3.5 pitch over diameter. In other embodiments, aft cooling
apertures 467 are spaced apart radially from 3 to 4 pitch over
diameter. Aft cooling apertures 467 may overlap with an adjacent
aft cooling aperture 467 rather than align in the radial direction
along the surface of pressure side wall 463.
[0033] Intermediate cooling apertures 468 may be grouped together
and located within the middle third of pressure side wall 463.
Intermediate cooling apertures 468 may be between forward cooling
apertures 466 and trailing edge 462. Intermediate cooling apertures
468 may also be between forward cooling apertures 466 and aft
cooling apertures 467. In some embodiments, intermediate cooling
apertures 468 are located from 1/4 to 3/8 of the length of pressure
side wall 463 from forward cooling apertures 466 and 1/4 to 3/8 of
the length of pressure side wall 463 from aft cooling apertures
467. In other embodiments, intermediate cooling apertures 468 are
located 1/3 of the length of pressure side wall 463 from forward
cooling apertures 466 and 1/3 of the length of pressure side wall
463 from aft cooling apertures 467. In yet other embodiments,
intermediate cooling apertures 468 are located at least 1/8 of the
length of pressure side wall 463 from forward cooling apertures 466
and at least 1/8 of the length of pressure side wall 463 from aft
cooling apertures 467. Intermediate cooling apertures 468 may be
arranged radially between upper endwall 453 and lower endwall 457.
In the embodiment illustrated in FIG. 2, intermediate cooling
apertures 468 are arranged in a single radial column and spaced
apart radially at 3.5 pitch over diameter. In other embodiments,
intermediate cooling apertures 468 are spaced apart radially from 3
to 4 pitch over diameter. Intermediate cooling apertures 468 may
overlap with an adjacent intermediate cooling aperture 468 rather
than being aligned in the radial direction along the surface of
pressure side wall 463.
[0034] While the embodiment illustrated in FIG. 2 includes forward
cooling apertures 466, aft cooling apertures 467, and intermediate
cooling apertures 468, some embodiments do not include aft cooling
apertures 467 and other embodiments include second intermediate
cooling apertures located between intermediate cooling apertures
468 and aft cooling apertures 467. The second intermediate cooling
apertures may be arranged similar to the arrangements of forward
cooling apertures 466, aft cooling apertures 467, and intermediate
cooling apertures 468. Other groups or columns of cooling apertures
may also be included. The spacing between groups or columns of
cooling apertures may be dependent on the number of groups or
columns of cooling apertures located along pressure side wall
463.
[0035] Airfoil 460 may further include slots 483. Slots 483 may be
located on pressure side wall 463 and may be adjacent trailing edge
462. Slots 483 may be rectangular and may be aligned in the radial
direction between upper endwall 453 and lower endwall 457. Slots
483 may extend from cooling cavity 485 to trailing edge 462.
[0036] In the embodiment illustrated in FIG. 2, nozzle segment 451
includes second airfoil 470. Second airfoil 470 may include the
same or similar features as airfoil 460 including second leading
edge 471, second trailing edge (not shown), second pressure side
wall 473, and second suction side wall 474. Second airfoil 470 may
further include second showerhead cooling apertures 475, second
forward cooling apertures 476, second aft cooling apertures 477,
second intermediate cooling apertures 478, and second slots (not
shown). The description of second leading edge 471, the second
trailing edge, second pressure side wall 473, second suction side
wall 474, second showerhead cooling apertures 475, second forward
cooling apertures 476, second aft cooling apertures 477, second
intermediate cooling apertures 478, and the second slots may be
oriented in the same or a similar manner as leading edge 461,
trailing edge 462, pressure side wall 463, suction side wall 464,
showerhead cooling apertures 465, forward cooling apertures 466,
aft cooling apertures 467, intermediate cooling apertures 468, and
slots 483 respectively. In other embodiments, nozzle segment 451
only includes airfoil 460 and not second airfoil 470.
[0037] The various components of nozzle segment 451 including upper
shroud 452, lower shroud 456, airfoil 460, and second airfoil 470
may be integrally cast or metalurgically bonded to form a unitary
or one piece assembly thereof.
[0038] In accordance with embodiments of this invention, the spaced
apart groups of cooling apertures, showerhead cooling apertures
465, forward cooling apertures 466, intermediate cooling apertures
468, and aft cooling apertures 467, alternate in directionality,
being angled or partially angled at lower endwall 457 or upper
endwall 453. The directionality or angle of the apertures directs
cooling air in a selected direction. In the embodiment illustrated
in FIGS. 2-6, showerhead cooling apertures 465 are angled toward
lower endwall 457, forward cooling apertures 466, the next grouping
of cooling holes along pressure side wall 463, are angled toward
upper endwall 453, intermediate cooling apertures 468, the
following grouping of cooling holes along pressure side wall 463,
are also angled at lower endwall 457, and aft cooling apertures
467, the last grouping of cooling holes, are also angled at upper
endwall 453. In other embodiments, showerhead cooling apertures 465
are angled toward upper endwall 453, forward cooling apertures 466
are angled toward lower endwall 457, intermediate cooling apertures
468 are also angled at upper endwall 453, and aft cooling apertures
467 are also angled at lower endwall 457.
[0039] In embodiments that include the second intermediate cooling
apertures and showerhead cooling apertures 465 angled toward lower
endwall 457, second intermediate cooling apertures would be the
grouping after intermediate cooling apertures 468 and would be
angled towards upper endwall 453 and aft cooling apertures 467
would be the grouping after the second intermediate cooling
apertures and would be angled toward lower endwall 457.
[0040] FIG. 3 is a cross-sectional view of a portion of the nozzle
segment 451 of FIG. 2 showing the showerhead cooling apertures 465.
In the embodiment illustrated in FIG. 3, showerhead cooling
apertures 465 may extend through a wall 444 of airfoil 460 a
cooling cavity 485 towards lower endwall 457. Wall 444 may be a
part of leading edge 461, pressure side wall 463, or suction side
wall 464. Showerhead cooling apertures 465 may be angled relative
to a reference plane 480. Reference plane 480 may be defined as a
plane perpendicular to a radial extending from the nozzle axis, the
axis of upper shroud 452 and lower shroud 456 and along leading
edge 461. In one embodiment, showerhead angle 481, the angle of
showerhead cooling apertures 465 relative to reference plane 480 is
from twenty to forty-five degrees towards the lower endwall 457. In
another embodiment, showerhead angle 481 is thirty degrees or
within a predetermined tolerance of thirty degrees. The
predetermined tolerance may be the engineering tolerance or the
manufacturing tolerance. While showerhead angle 481 is directed
toward lower endwall 457 in the embodiment illustrated, showerhead
angle 481 may be directed toward upper endwall 453 in other
embodiments.
[0041] Each showerhead cooling aperture 465 may also be angled
towards the lower endwall 457 or the upper endwall 453 relative to
the direction normal to leading edge 461 at the location where the
showerhead cooling aperture 465 is located.
[0042] As illustrated in FIG. 3, each showerhead cooling aperture
465 may include showerhead inlet end 491 adjacent cooling cavity
485 and showerhead outlet end 492 located at the surface of leading
edge 461. Showerhead inlet end 491 may be radially closer to the
upper endwall 453 than showerhead outlet end 492 and showerhead
outlet end 492 may be radially closer to the lower endwall 457 than
showerhead inlet end 491.
[0043] FIG. 4 is a detailed view of the forward cooling apertures
466 of FIG. 2. FIG. 5 is a detailed view of the intermediate
cooling apertures 468 of FIG. 2. FIG. 6 is a detailed view of the
aft cooling apertures 467 of FIG. 2. Referring to FIGS. 4, 5, and
6, forward cooling apertures 466, aft cooling apertures 467, and
intermediate cooling apertures 468 may be angled relative to the
flow direction of the air traveling through turbine nozzle 450
along pressure side surface 469 during operation of gas turbine
engine 100. Reference line 482 illustrates the flow direction.
Reference line 482 may also be defined as the intersection between
pressure side surface 469 and a plane perpendicular to a radial
extending from the turbine nozzle axis, the axis of upper shroud
452 and lower shroud 456, along the pressure side surface 469.
[0044] Referring to FIGS. 2 and 4, forward cooling apertures 466
may be angled at a forward compound angle 486. Forward compound
angle 486 may be the component of the angle of forward cooling
apertures 466 in the plane of pressure side surface 469. As
illustrated, forward compound angle 486 is angled toward upper
endwall 453 relative to the flow direction or reference line 482.
In one embodiment, forward compound angle 486 is from fifteen to
forty-five degrees. In another embodiment, forward compound angle
486 is thirty degrees or within a predetermined tolerance of thirty
degrees. The predetermined tolerance may be the engineering
tolerance or the manufacturing tolerance. Zero degrees may be the
flow direction of the direction along reference line 482 traveling
from leading edge 461 to trailing edge 462. While forward compound
angle 486 is directed towards upper endwall 453 in the embodiment
illustrated, forward compound angle 486 is directed towards lower
endwall 457 in embodiments where showerhead angle 481 is directed
towards upper endwall 453.
[0045] Referring to FIGS. 2 and 5, intermediate cooling apertures
468 may be angled at an intermediate compound angle 488.
Intermediate compound angle 488 may be the component of the angle
of intermediate cooling apertures 468 in the plane of pressure side
surface 469. As illustrated, intermediate compound angle 488 is
angled toward the lower endwall 457 relative to the flow direction
or reference line 482. In one embodiment, intermediate compound
angle 488 is from fifteen to forty-five degrees. In another
embodiment, intermediate compound angle 488 is thirty degrees or
within a predetermined tolerance of thirty degrees. The
predetermined tolerance may be the engineering tolerance or the
manufacturing tolerance. While intermediate compound angle 488 is
directed towards lower endwall 457 in the embodiment illustrated,
intermediate compound angle 488 is directed towards upper endwall
453 in embodiments where forward compound angle 486 is directed
towards lower endwall 457.
[0046] Referring to FIGS. 2 and 6, aft cooling apertures 467 may be
angled at an aft compound angle 487. Aft compound angle 487 may be
the component of the angle of aft cooling apertures 467 in the
plane of pressure side surface 469. As illustrated, aft compound
angle 487 is angled toward the upper endwall 453 relative to the
flow direction or reference line 482 and may be similar or equal to
forward compound angle 486. In one embodiment, aft compound angle
487 is from fifteen to forty-five degrees. In another embodiment,
aft compound angle 487 is thirty degrees or within a predetermined
tolerance of thirty degrees. The predetermined tolerance may be the
engineering tolerance or the manufacturing tolerance. While aft
compound angle 487 is directed towards upper endwall 453 in the
embodiment illustrated, aft compound angle 487 may be directed
towards lower endwall 457 in embodiments where intermediate
compound angle 488 is directed towards upper endwall 453.
[0047] In embodiments not including intermediate cooling apertures
468 or embodiments including second intermediate cooling apertures,
aft compound angle 487 may be angled toward lower endwall 457
relative to the flow direction or reference line 482. In one of
these embodiments, aft compound angle 487 is from fifteen to
forty-five degrees. In another of these embodiments, aft compound
angle 487 is approximately thirty degrees.
[0048] FIG. 7 is a cross-section of the airfoil 460 of FIG. 2.
Referring to FIG. 7, forward cooling apertures 466 may also include
a forward injection angle 441. Forward injection angle 441 may be
the component of the angle of forward cooling apertures 466 in the
plane perpendicular to pressure side surface 469. Forward injection
angle 441 may be measured relative to a line extending toward
trailing edge 462 and tangent to pressure side surface 469 at the
location of each forward cooling aperture 466. In one embodiment,
forward injection angle 441 is from fifteen to fifty degrees. In
another embodiment, forward injection angle 441 is approximately
thirty degrees.
[0049] Aft cooling apertures 467 may also include an aft injection
angle 442. Aft injection angle 442 may be the component of the
angle of aft cooling apertures 467 in the plane perpendicular to
pressure side surface 469. Aft injection angle 442 may be measured
relative to a line extending toward trailing edge 462 and tangent
to pressure side surface 469 at the location of each aft cooling
aperture 467. In one embodiment, aft injection angle 442 is from
fifteen to fifty degrees. In another embodiment, aft injection
angle 442 is approximately thirty degrees.
[0050] Intermediate cooling apertures 468 may also include an
intermediate injection angle 443. Intermediate injection angle 443
may be the component of the angle of intermediate cooling apertures
468 in the plane perpendicular to pressure side surface 469.
Intermediate injection angle 443 may be measured relative to a line
extending toward trailing edge 462 and tangent to pressure side
surface 469 at the location of each intermediate cooling aperture
468. In one embodiment, intermediate injection angle 443 is from
fifteen to fifty degrees. In another embodiment, intermediate
injection angle 443 is approximately thirty degrees.
[0051] Cooling cavity 485 may be a single cavity or may be
subdivided into multiple cavities. In the embodiment illustrated in
FIG. 7, cooling cavity 485 is subdivided into two cooling
cavities.
[0052] Each forward cooling aperture 466 may include forward inlet
end 493 adjacent cooling cavity 485 and forward outlet end 494
adjacent or at pressure side surface 469. Each intermediate cooling
aperture 468 may include intermediate inlet end 497 adjacent
cooling cavity 485 and intermediate outlet end 498 adjacent or at
pressure side surface 469. Each aft cooling aperture 467 may
include aft inlet end 495 adjacent cooling cavity 485 and aft
outlet end 496 adjacent or at pressure side surface 469.
[0053] The compound angles may be determined by the positions of
the inlet ends and the outlet ends of the apertures relative to
lower endwall 457 and upper endwall 453, while the injection angle
may be determined by the positions of the inlet ends and the outlet
ends relative to leading edge 461 and trailing edge 462.
[0054] In the embodiment illustrated in FIG. 4, forward inlet end
493 is radially closer to lower endwall 457 and axially closer to
leading edge 461 than forward outlet end 494, and forward outlet
end 494 is radially closer to upper endwall 453 and axially closer
to trailing edge 462 than forward inlet end 493. In other
embodiments, forward inlet end 493 is radially closer to upper
endwall 453 and axially closer to leading edge 461 than forward
outlet end 494, and forward outlet end 494 is radially closer to
lower endwall 457 and axially closer to trailing edge 462 than
forward inlet end 493.
[0055] In the embodiment illustrated in FIG. 5, intermediate inlet
end 497 is radially closer to upper endwall 453 and axially closer
to leading edge 461 than intermediate outlet end 498, and
intermediate outlet end 498 is radially closer to lower endwall 457
and axially closer to trailing edge 462 than intermediate inlet end
497. In other embodiments, intermediate inlet end 497 is radially
closer to lower endwall 457 and axially closer to leading edge 461
than intermediate outlet end 498, and intermediate outlet end 498
is radially closer to upper endwall 453 and axially closer to
trailing edge 462 than intermediate inlet end 497.
[0056] In the embodiment illustrated in FIG. 6, aft inlet end 495
is radially closer to lower endwall 457 and axially closer to
leading edge 461 than aft outlet end 496, and aft outlet end 496 is
radially closer to upper endwall 453 and axially closer to trailing
edge 462 than aft inlet end 495. In other embodiments, aft inlet
end 495 is radially closer to upper endwall 453 and axially closer
to leading edge 461 than aft outlet end 496, and aft outlet end 496
is radially closer to lower endwall 457 and axially closer to
trailing edge 462 than aft inlet end 495.
[0057] One or more of the above components (or their subcomponents)
may be made from stainless steel and/or durable, high temperature
materials known as "superalloys". A superalloy, or high-performance
alloy, is an alloy that exhibits excellent mechanical strength and
creep resistance at high temperatures, good surface stability, and
corrosion and oxidation resistance. Superalloys may include
materials such as HASTELLOY, alloy x, INCONEL, WASPALOY, RENE
alloys, HAYNES alloys, alloy 188, alloy 230, INCOLOY, MP98T, TMS
alloys, and CMSX single crystal alloys.
INDUSTRIAL APPLICABILITY
[0058] Gas turbine engines may be suited for any number of
industrial applications such as various aspects of the oil and gas
industry (including transmission, gathering, storage, withdrawal,
and lifting of oil and natural gas), the power generation industry,
cogeneration, aerospace, and other transportation industries.
[0059] Referring to FIG. 1, a gas (typically air 10) enters the
inlet 110 as a "working fluid", and is compressed by the compressor
200. In the compressor 200, the working fluid is compressed in an
annular flow path 115 by the series of compressor disk assemblies
220. In particular, the air 10 is compressed in numbered "stages",
the stages being associated with each compressor disk assembly 220.
For example, "4th stage air" may be associated with the 4th
compressor disk assembly 220 in the downstream or "aft" direction,
going from the inlet 110 towards the exhaust 500). Likewise, each
turbine disk assembly 420 may be associated with a numbered
stage.
[0060] Once compressed air 10 leaves the compressor 200, it enters
the combustor 300, where it is diffused and fuel is added. Air 10
and fuel are injected into the combustion chamber 390 via fuel
injector 310 and combusted. Energy is extracted from the combustion
reaction via the turbine 400 by each stage of the series of turbine
disk assemblies 420. Exhaust gas 90 may then be diffused in exhaust
diffuser 520, collected and redirected. Exhaust gas 90 exits the
system via an exhaust collector 550 and may be further processed
(e.g., to reduce harmful emissions, and/or to recover heat from the
exhaust gas 90).
[0061] Operating efficiency of a gas turbine engine generally
increases with a higher combustion temperature. Thus, there is a
trend in gas turbine engines to increase the temperatures. Gas
reaching forward stages of a turbine from a combustion chamber 390
may be 1000 degrees Fahrenheit or more. To operate at such high
temperatures a portion of the compressed air from the compressor
200, cooling air, may be diverted through internal passages or
chambers to cool various components of a turbine including turbine
nozzle segments such as nozzle segment 451. However, the use of
cooling air may reduce the operating efficiency of the gas turbine
engine.
[0062] Alternating the direction of groupings of cooling apertures
such as showerhead cooling apertures 465, forward cooling apertures
466, intermediate cooling apertures 468, and aft cooling apertures
467, to direct cooling air towards upper endwall 453 of upper
shroud 452 and lower endwall 457 of lower shroud 456 may reduce the
temperatures of upper endwall 453 and lower endwall 457, which may
improve the operating life of nozzle segment 451.
[0063] The first order cooling or initial use of the cooling air
exiting showerhead cooling apertures 465, forward cooling apertures
466, intermediate cooling apertures 468, and aft cooling apertures
467 may be to film cool pressure side wall 463. The second order
cooling or second use of the cooling air may be to reduce the
temperatures of upper endwall 453 and lower endwall 457.
[0064] The cooling air may be directed through turbine housing 430,
turbine diaphragm 440, or both and into cooling cavity 485. The
cooling air may then be directed through the cooling apertures
including showerhead cooling apertures 465, forward cooling
apertures 466, intermediate cooling apertures 468, and aft cooling
apertures 467. The cooling air may also be used for cooling airfoil
460 internally prior to passing through the cooling apertures. The
multiple uses of the cooling air that may include the first order
film cooling, the second order endwall cooling, and the internal
cooling may reduce the amount of air needed to effectively cool
nozzle segment 451. Reducing the amount of air needed to cool
nozzle segment 451 may improve or increase the efficiency of gas
turbine engine 100.
[0065] The cooling apertures of second airfoil 470 may be used in
the same or a similar manner as the cooling apertures of airfoil
460 resulting in a further reduction of the temperatures of upper
endwall 453 and lower endwall 457, as well as the reduction in the
amount of cooling air needed to effectively cool each nozzle
segment 451.
[0066] The preceding detailed description is merely exemplary in
nature and is not intended to limit the invention or the
application and uses of the invention. The described embodiments
are not limited to use in conjunction with a particular type of gas
turbine engine. Hence, although the present disclosure, for
convenience of explanation, depicts and describes a particular
nozzle segment, it will be appreciated that the nozzle segment in
accordance with this disclosure can be implemented in various other
configurations, can be used with various other types of gas turbine
engines, and can be used in other types of machines. Furthermore,
there is no intention to be bound by any theory presented in the
preceding background or detailed description. It is also understood
that the illustrations may include exaggerated dimensions to better
illustrate the referenced items shown, and are not consider
limiting unless expressly stated as such.
* * * * *