U.S. patent application number 11/967190 was filed with the patent office on 2009-07-02 for cooled turbine nozzle segment.
Invention is credited to Michael Scott Cole, James Herbert Deines, Ching-Pang Lee.
Application Number | 20090169361 11/967190 |
Document ID | / |
Family ID | 40690918 |
Filed Date | 2009-07-02 |
United States Patent
Application |
20090169361 |
Kind Code |
A1 |
Cole; Michael Scott ; et
al. |
July 2, 2009 |
COOLED TURBINE NOZZLE SEGMENT
Abstract
A turbine nozzle segment may have a band having a flange
extending radially from a non-flowpath side and an aft end. A
plurality of airfoils may extend radially from a flowpath side of
the band and may have trailing edges. A plurality of cooling holes
may be disposed in the flange and directed at the aft end between
the trailing edges.
Inventors: |
Cole; Michael Scott; (Mason,
OH) ; Deines; James Herbert; (Mason, OH) ;
Lee; Ching-Pang; (Cincinnati, OH) |
Correspondence
Address: |
GENERAL ELECTRIC COMPANY
GE AVIATION, ONE NEUMANN WAY MD H17
CINCINNATI
OH
45215
US
|
Family ID: |
40690918 |
Appl. No.: |
11/967190 |
Filed: |
December 29, 2007 |
Current U.S.
Class: |
415/115 ;
415/177 |
Current CPC
Class: |
Y02T 50/60 20130101;
F05B 2260/301 20130101; Y02T 50/6765 20180501; Y02T 50/67 20130101;
F01D 5/082 20130101; F05D 2230/90 20130101; Y02T 50/676 20130101;
F01D 9/042 20130101; F05D 2230/64 20130101 |
Class at
Publication: |
415/115 ;
415/177 |
International
Class: |
F01D 5/08 20060101
F01D005/08; F01D 5/14 20060101 F01D005/14 |
Claims
1. A turbine nozzle segment comprising: a band having a flowpath
side, a non-flowpath side, a flange extending radially from said
non-flowpath side and an aft end; a plurality of airfoils extending
radially from said flowpath side, said airfoils having trailing
edges; and a plurality of cooling holes disposed in said flange,
said cooling holes directed at said aft end of said non-flowpath
side of said band between said trailing edges.
2. The turbine nozzle segment of claim 1 wherein said cooling holes
are directed so as to impinge upon said aft end of said
non-flowpath side of said band.
3. The turbine nozzle segment of claim 1 wherein said cooling holes
have a compound angle relative to a line parallel to the engine
centerline.
4. The turbine nozzle segment of claim 3 wherein said cooling holes
have a first angle measured in the radial plane relative to a line
parallel to the engine centerline between about 10 degrees and
about 75 degrees.
5. The turbine nozzle segment of claim 4 wherein said cooling holes
have a second angle measured in the circumferential plane relative
to a line parallel to the engine centerline between about 10
degrees and about 80 degrees.
6. The turbine nozzle segment of claim 1 wherein said cooling holes
have a first angle measured in the radial plane relative to a line
parallel to the engine centerline between about 10 degrees and
about 75 degrees.
7. The turbine nozzle segment of claim 1 wherein said cooling holes
have a second angle measured in the circumferential plane relative
to a line parallel to the engine centerline between about 10
degrees and about 80 degrees.
8. The turbine nozzle segment of claim 1 further comprising a
thermal barrier coating applied to said aft end of said flowpath
side of said band between said airfoil trailing edges.
9. The turbine nozzle segment of claim 1 wherein said flange is
located near said aft end.
10. The turbine nozzle segment of claim 1 further comprising: a
plenum on said non-flowpath side of said band for providing cooling
air to said plurality of cooling holes.
11. A turbine nozzle assembly comprising: a plurality of arcuate
turbine nozzle segments joined together to form an annular ring;
said plurality of arcuate segments each comprising: a band having a
flowpath side, a non-flowpath side, a flange extending radially
from said non-flowpath side and an aft end; a plurality of airfoils
extending radially from said flowpath side, said airfoils having
trailing edges; and a plurality of cooling holes disposed in said
flange, said cooling holes directed at said aft end of said
non-flowpath side of said band between said trailing edges.
12. The turbine nozzle assembly of claim 11 wherein said cooling
holes are directed so as to impinge upon said aft end of said
non-flowpath side of said band.
13. The turbine nozzle assembly of claim 11 wherein said cooling
holes have a compound angle relative to a line parallel to the
engine centerline.
14. The turbine nozzle assembly of claim 13 wherein said cooling
holes have a first angle measured in the radial plane relative to a
line parallel to the engine centerline between about 10 degrees and
about 75 degrees.
15. The turbine nozzle assembly of claim 14 wherein said cooling
holes have a second angle measured in the circumferential plane
relative to a line parallel to the engine centerline between about
10 degrees and about 80 degrees.
16. The turbine nozzle assembly of claim 11 wherein said cooling
holes have a first angle measured in the radial plane relative to a
line parallel to the engine centerline between about 10 degrees and
about 75 degrees.
17. The turbine nozzle assembly of claim 11 wherein said cooling
holes have a second angle measured in the circumferential plane
relative to a line parallel to the engine centerline between about
10 degrees and about 80 degrees.
18. The turbine nozzle assembly of claim 11 further comprising a
thermal barrier coating applied to said aft end of said flowpath
side of said band between said airfoil trailing edges.
19. The turbine nozzle assembly of claim 11 wherein said flange is
located near said aft end.
20. The turbine nozzle assembly of claim 11 further comprising: a
plenum on said non-flowpath side of said band for providing cooling
air to said plurality of cooling holes.
Description
BACKGROUND OF THE INVENTION
[0001] The exemplary embodiments relate generally to gas turbine
engine components and more particularly to turbine nozzle segments
having improved cooling.
[0002] Gas turbine engines typically include a compressor, a
combustor, and at least one turbine. The compressor may compress
air, which may be mixed with fuel and channeled to the combustor.
The mixture may then be ignited for generating hot combustion
gases, and the combustion gases may be channeled to the turbine.
The turbine may extract energy from the combustion gases for
powering the compressor, as well as producing useful work to propel
an aircraft in flight or to power a load, such as an electrical
generator.
[0003] The turbine may include a stator assembly and a rotor
assembly. The stator assembly may include a stationary nozzle
assembly having a plurality of circumferentially spaced apart
airfoils extending radially between inner and outer bands, which
define a flow path for channeling combustion gases therethrough.
Typically the airfoils and bands are formed into a plurality of
segments, which may include one or two spaced apart airfoils
radially extending between an inner and an outer band. The segments
are joined together to form the nozzle assembly. The band may
include one or more flanges for attaching the nozzle assembly to
other components of the gas turbine engine.
[0004] The rotor assembly may be downstream of the stator assembly
and may include a plurality of blades extending radially outward
from a disk. Each rotor blade may include an airfoil, which may
extend between a platform and a tip. Each rotor blade may also
include a root that may extend below the platform and be received
in a corresponding slot in the disk. Alternatively, the disk may be
a blisk or bladed disk, which may alleviate the need for a root and
the airfoil may extend directly from the disk. The rotor assembly
may be bounded radially at the tip by a stationary annular shroud.
The shrouds and platforms (or disk, in the case of a blisk) define
a flow path for channeling the combustion gases therethrough.
[0005] As gas temperatures rise due to the demand for increased
performance, components may not be able to withstand the increased
temperatures. Higher gas temperatures lead to higher metal
temperatures, which is a primary contributor to distress. Distress
may cause cracking or holes to form within these areas, leading to
decreased performance and higher repair costs. Higher pressure and
temperature areas suffer the greatest distress. As shown in FIG. 1,
one such higher temperature and pressure area 80 is between the
trailing edges of the airfoils in a nozzle segment. In this area,
the pressure and temperature combination is highest and is the most
susceptible to damage.
BRIEF DESCRIPTION OF THE INVENTION
[0006] In one exemplary embodiment, a turbine nozzle segment may
have a band having a flowpath side, a non-flowpath side, a flange
extending radially from the non-flowpath side and an aft end. The
nozzle segment may further include a plurality of airfoils having
trailing edges and extending radially from the flowpath side. The
nozzle segment may also have a plurality of cooling holes disposed
in the flange, the cooling holes directed at the aft end of the
non-flowpath side of the band between the trailing edges.
[0007] In another exemplary embodiment, a turbine nozzle assembly
may include a plurality of arcuate turbine nozzle segments joined
together to form an annular ring, each of the plurality of arcuate
segments having a band having a flowpath side, a non-flowpath side,
a flange extending radially from the non-flowpath side and an aft
end. The nozzle segment may further include a plurality of airfoils
having trailing edges and extending radially from the flowpath
side. The nozzle segment may also have a plurality of cooling holes
disposed in the flange, the cooling holes directed at the aft end
of the non-flowpath side of the band between the trailing
edges.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] FIG. 1 is a schematic diagram illustrating the pressures and
temperatures of a typical turbine nozzle segment.
[0009] FIG. 2 is a cross-sectional view of an exemplary gas turbine
engine.
[0010] FIG. 3 is a cross-sectional view of an exemplary embodiment
of a turbine nozzle assembly.
[0011] FIG. 4 is a close-up cross-sectional view of the outer band
area of an exemplary embodiment of a turbine nozzle assembly.
[0012] FIG. 5 is a perspective view of an exemplary embodiment of a
turbine nozzle segment.
[0013] FIG. 6 is a top plan view of an exemplary embodiment of a
turbine nozzle segment.
[0014] FIG. 7 is a perspective view of an exemplary embodiment of a
turbine nozzle segment.
DETAILED DESCRIPTION OF THE INVENTION
[0015] FIG. 2 illustrates a cross-sectional schematic view of an
exemplary gas turbine engine 100. The gas turbine engine 100 may
include a low-pressure compressor 102, a high-pressure compressor
104, a combustor 106, a high-pressure turbine 108, and a
low-pressure turbine 110. The low-pressure compressor may be
coupled to the low-pressure turbine through a shaft 112. The
high-pressure compressor 104 may be coupled to the high-pressure
turbine 108 through a shaft 114. In operation, air flows through
the low-pressure compressor 102 and high-pressure compressor 104.
The highly compressed air is delivered to the combustor 106, where
it is mixed with a fuel and ignited to generate combustion gases.
The combustion gases are channeled from the combustor 106 to drive
the turbines 108 and 110. The turbine 110 drives the low-pressure
compressor 102 by way of shaft 112. The turbine 108 drives the
high-pressure compressor 104 by way of shaft 114.
[0016] As shown in FIGS. 3-7, the high-pressure turbine 108 may
include a turbine nozzle assembly 116. The turbine nozzle assembly
116 may be downstream of the combustor 106 or a row of turbine
blades. The turbine nozzle assembly 116 includes an annular array
of turbine nozzle segments 118. A plurality of arcuate turbine
nozzle segments 118 may be joined together to form the annular
turbine nozzle assembly 116. The turbine nozzle segments 118 may
have an inner band 120 and an outer band 122, which radially bound
the flow of combustion gases through the turbine nozzle assembly
116. The inner band 120 may have a flowpath side 124 and a
non-flowpath side 126 and the outer band 122 may have a flowpath
side 128 and a non-flowpath side 130. One or more flanges 132 may
extend from the non-flowpath sides 126 and 130 of the inner band
120 and outer band 122. For example, as shown in FIG. 3, flange 134
extends radially from said the outer band 122 and may be used to
attach the turbine nozzle assembly 116 to other components of the
gas turbine engine 100.
[0017] Airfoils 136 extend radially between the inner band 120 and
outer band 122 for directing the flow of combustion gases through
the turbine nozzle assembly 116. The airfoils 136 have a leading
edge 138 on the forward side of the turbine nozzle segment 118 and
a trailing edge 140 on the aft side of the turbine nozzle segment
118. The airfoils 136 may be formed of solid or hollow
construction. Hollow airfoils may include one or more internal
cooling passages for cooling the airfoil and providing film cooling
to the airfoil surfaces. Other hollow airfoils may include one or
more cavities for receiving a cooling insert. The cooling insert
may have a plurality of cooling holes for impinging on the interior
surface of the hollow airfoil before exiting as film cooling
through holes in the airfoil. Any configuration of airfoil known in
the art may be used.
[0018] Band, as used below, may mean the inner band 120, the outer
band 122 or each of the inner band 120 and outer band 122. The band
may have one or more flanges 132 extending radially from the
non-flowpath side 126, 130. At least one of the flanges 132 may be
located near the aft side of the nozzle segment 118, such as, but
not limited to, flange 134 in FIG. 3. Upstream of the flange 134,
may be a plenum 142. The plenum 142 may receive cooling air from
another part of the engine, such as, the high-pressure compressor
104. The cooling air may be provided to the plenum 142 through any
means known in the art.
[0019] A plurality of cooling holes 144 may be disposed within the
flange 134. The cooling holes 144 may have an inlet 146 at the
plenum 142 on the upstream side of the flange 134 and an outlet 148
on the downstream side of the flange 134. The inlet 146 may receive
cooling air from the plenum 142 and flow the cooling air through to
the outlet 148. The cooling hole 144 and outlet 148 may be arranged
so that the outlet 148 is directed at the aft end 150 of the band,
so as to impinge on the aft end 150. The outlets 148 may have any
shaped known in the art. Further, the holes 144 may be formed in
any manner known in the art, such as, but not limited to,
electrodischarge machining, electrochemical machining, laser
drilling, mechanical drilling, or any other similar manner.
[0020] In one exemplary embodiment, as shown in FIGS. 3, 4 and 6,
the cooling holes 144 may have a compound angle. The cooling holes
144 may have a first angle .beta. measured in the radial plane (the
X-Y plane) relative to a line parallel to the engine centerline 152
so that the outlet is directed at the aft end 150. The cooling
holes 144 may have a second angle .alpha. measured in the
circumferential plane (the X-Z plane) relative to a line parallel
to the engine centerline 152 so that the cooling holes 144 are
directed generally in the direction of flow exiting the nozzle
segment as directed by the airfoil trailing edges 140. The first
angle .beta. may be between about 10 degrees and about 75 degrees.
The second angle .alpha. may be between about 10 degrees and about
80 degrees. The cooling holes 144 may be positioned such that they
are directed at an area of high pressure and temperature. In one
exemplary embodiment, the cooling holes may be directed at an area
158 on the aft end 150 of the band on the non-flowpath side 126,
130 between the trailing edges 140 of the airfoils 136. In another
exemplary embodiment, the cooling holes 144 may be directed at the
aft end 150 in a single plane, such that the holes 144 have one
angle .beta. measured in the radial plane (the X-Y plane) relative
to a line parallel to the engine centerline 152. In this exemplary
embodiment, all other angles would be zero.
[0021] In one exemplary embodiment, a thermal barrier coating (TBC)
160 may be applied to the band flowpath surface 124, 128. The TBC
may be between about 5 mils and about 25 mils thick. Any TBC known
in the art may be used. In one exemplary embodiment, the TBC may be
a three layer TBC having a MCrAlY first layer, where M is selected
from the group of Ni and Co, an aluminide second layer, and a
yttria-stablized zirconia (YSZ) third layer. In another exemplary
embodiment, a two layer TBC may be used where platinum aluminide or
aluminide may be used in place of the MCrAlY first layer and the
aluminide second layer.
[0022] By providing cooling holes in these areas and in particular
by impinging cooling air in these areas, the metal temperature may
be reduced, leading to less distress and less likelihood of forming
a crack or hole. As such, the turbine nozzle segment will last
longer leading to less repairs and/or replacements over time for
the gas turbine engine.
[0023] This written description discloses exemplary embodiments,
including the best mode, to enable any person skilled in the art to
make and use the exemplary embodiments. The patentable scope is
defined by the claims, and may include other examples that occur to
those skilled in the art. Such other examples are intended to be
within the scope of the claims if they have structural elements
that do not differ from the literal language of the claims, or if
they include equivalent structural elements with insubstantial
differences from the literal languages of the claims.
* * * * *