U.S. patent application number 10/039719 was filed with the patent office on 2003-07-10 for methods and apparatus for cooling gas turbine nozzles.
Invention is credited to Manning, Robert F., Roach, Christopher, Rydbeck, Randall B..
Application Number | 20030129054 10/039719 |
Document ID | / |
Family ID | 21907013 |
Filed Date | 2003-07-10 |
United States Patent
Application |
20030129054 |
Kind Code |
A1 |
Manning, Robert F. ; et
al. |
July 10, 2003 |
METHODS AND APPARATUS FOR COOLING GAS TURBINE NOZZLES
Abstract
A method for fabricating a nozzle for a gas turbine engine
facilitates extending a useful life of the nozzles. The nozzle
includes an airfoil. The method includes forming the airfoil to
include a suction side and a pressure side connected at a leading
edge and a trailing edge, forming a plurality of slots in the
pressure side of the airfoil extending towards the trailing edge,
and extending a plurality of pins arranged in rows between the
airfoil suction and pressure sides, such that at each of a first
row of pins has a substantially elliptical cross-sectional area,
wherein the first row of pins is between the remaining rows of pins
and the airfoil pressure side slots.
Inventors: |
Manning, Robert F.;
(Newburyport, MA) ; Rydbeck, Randall B.;
(Hamilton, MA) ; Roach, Christopher; (Wollaston,
MA) |
Correspondence
Address: |
John S. Beulick
Armstrong Teasdale LLP
Suite 2600
One Metropolitan Sq.
St. Louis
MO
63102
US
|
Family ID: |
21907013 |
Appl. No.: |
10/039719 |
Filed: |
January 4, 2002 |
Current U.S.
Class: |
415/115 ;
416/97R |
Current CPC
Class: |
F01D 5/187 20130101;
Y02T 50/676 20130101; Y02T 50/60 20130101; F05D 2260/22141
20130101; F05D 2240/10 20130101 |
Class at
Publication: |
415/115 ;
416/97.00R |
International
Class: |
F01D 009/06 |
Claims
What is claimed is:
1. A method for fabricating a nozzle for a gas turbine engine, the
nozzle including an airfoil, said method comprising: forming the
airfoil to include a suction side and a pressure side connected at
a leading edge and a trailing edge; forming a plurality of slots in
the pressure side of the airfoil extending towards the trailing
edge; and extending a plurality of pins arranged in rows between
the airfoil suction and pressure sides, such that at each of a
first row of pins has a substantially elliptical cross-sectional
area, wherein the first row of pins is between the remaining rows
of pins and the airfoil pressure side slots.
2. A method in accordance with claim I wherein forming a plurality
of slots further comprises forming a plurality of slots to extend
radially between an airfoil root and an airfoil tip, such that a
first slot is adjacent the airfoil root, a second slot is adjacent
the airfoil root, and a plurality of identical third slots extend
between the first and second slots.
3. A method in accordance with claim 2 wherein forming a plurality
of slots in the pressure side further comprises defining the slots
with adjacent land areas such that a radial height of each of the
third slots is less than a radial height of the first and second
slots.
4. A method in accordance with claim 3 wherein defining the slots
further comprises defining the slots such that the first and second
slots are identical and have a radial height that is approximately
twice that of each of the plurality of third slots.
5. A method in accordance with claim 3 wherein defining the slots
further comprises defining the slots with adjacent land areas such
that each land area has a substantially semi-circular-shaped
upstream portion and a substantially triangular-shaped downstream
portion.
6. A method in accordance with claim 1 wherein forming the airfoil
further comprises casting the airfoil.
7. A turbine nozzle for a gas turbine engine, said nozzle
comprising an airfoil vane comprising a first wall, a second wall,
and a plurality of pins extending therebetween, said first and
second walls connected at a leading edge and a trailing edge, said
first wall comprising a plurality of slots extending towards said
trailing edge, said plurality of pins comprising at least a first
row of pins having a substantially elliptical cross-sectional area,
said first row of pins between said remaining plurality of pins and
said first wall slots.
8. A turbine nozzle in accordance with claim 7 wherein said airfoil
first row of pins is tapered such that a downstream side of each
said pin has a width that is less than an upstream side of each
said pin.
9. A turbine nozzle in accordance with claim 7 wherein said airfoil
further comprises a root, a tip, and a plurality of land areas,
said first and second walls extending radially between said root
and said tip, said plurality of land areas separate said slots,
each said slot having a radial height extending between adjacent
said land areas.
10. A turbine nozzle in accordance with claim 9 wherein said
airfoil plurality of slots further comprise a first slot adjacent
said airfoil root, a second slot adjacent said airfoil tip, and a
plurality of identical third slots extending between said first and
second slots, each said third slot having a radial height less than
a radial height of said first and said second slots.
11. A turbine nozzle in accordance with claim 10 wherein said
airfoil first and second slots are identical and have a radial
height that is approximately twice each said third slot.
12. A turbine nozzle in accordance with claim 9 wherein each said
slot comprises a centerline, each said first row pin comprises a
centerline, each said first row pin centerline substantially
parallel to each said slot centerline.
13. A turbine nozzle in accordance with claim 9 wherein at least
one of said airfoil land areas between adjacent said slots
comprises a substantially semicircular-shaped upstream side and a
substantially triangular-shaped downstream side.
14. A turbine nozzle in accordance with claim 9 wherein said first
wall defines a pressure side of said nozzle, said first row of pins
and said first wall slots configured to provide cooling air to said
airfoil to facilitate extending a useful life of said nozzle.
15. An airfoil for a gas turbine engine nozzle, said airfoil
comprising a root, a tip, a plurality of pins, a convex sidewall
and a concave sidewall connected at a trailing edge, each said
sidewall extending between said root and tip, said convex sidewall
defining a pressure side of said airfoil and comprising a plurality
of slots extending towards said trailing edge, said plurality of
pins comprising at least a first row of pins and a second row of
pins, said first row of pins concentrically aligned radially, each
said first row pin having a substantially elliptical cross
sectional profile and tapered such that an upstream side of each
said first row pin has a width that is greater than a downstream
side of each said first row pin, said slots adjacent to and
downstream from said first row of pins.
16. An airfoil in accordance with claim 15 wherein said plurality
of slots are defined by a plurality of land areas such that each
said slot comprises a radial height, said plurality of slots
comprising a first slot adjacent said airfoil root, a second slot
adjacent said airfoil tip, and a plurality of third slots extending
between said first and second slots.
17. An airfoil in accordance with claim 16 wherein each said land
area defining a third slot comprises a substantially
semi-circular-shaped upstream portion and a substantially
triangular-shaped downstream portion extending from said upstream
portion.
18. An airfoil in accordance with claim 16 wherein each of said
plurality of slots has a radial height, each of said plurality of
third slots having a radial height less than each of said first and
second slots.
19. An airfoil in accordance with claim 16 wherein each of said
plurality of slots has a radial height, said first and second slots
identical and having a radial height that is approximately twice
that of said plurality of third slots.
20. An airfoil in accordance with claim 16 wherein each said slot
comprises a centerline, each said first row pin comprises a
centerline, each said first row pin centerline substantially
parallel with each said slot area centerline.
Description
BACKGROUND OF THE INVENTION
[0001] This invention relates generally to gas turbine engine
nozzles and more particularly, to methods and apparatus for cooling
gas turbine engine nozzles.
[0002] Gas turbine engines include combustors which ignite fuel-air
mixtures which are then channeled through a turbine nozzle assembly
towards a turbine. At least some known turbine nozzle assemblies
include a plurality of nozzles arranged circumferentially within
the engine downstream from the combustors. Each nozzle includes an
airfoil vane that extends between integrally-formed inner and outer
band platforms. The nozzles are cooled by a combination of internal
convective cooling and gas side film cooling.
[0003] Each nozzle includes a pair of sidewalls that are connected
at a leading edge and a trailing edge. The metal temperature
distribution of a typical vane airfoil is such that the trailing
edge is significantly hotter than a temperature of the bulk of the
airfoil. The temperature gradient created results in high
compressive stress at the vane trailing edge, and the combination
of high stresses and high temperatures generally results in the
vane trailing edge being the life limiting location of the nozzle.
Accordingly, within at least some known nozzles, the airfoil vane
trailing edge is cooled by a film of cooling air discharged from an
internally-defined vane cavity. More specifically, the film of
cooling air is discharged through trailing edge slots formed on the
airfoil vane pressure side, and upstream from the airfoil vane
trailing edge.
[0004] The amount of air supplied to each nozzle vane is attempted
to be optimized to lessen the effect on engine performance
decrement that may be associated with cooling flow extraction.
Generally, the slots are formed with a length that facilitates
optimizing an amount of cooling flow supplied to the trailing edge.
Because of the slot length, such slots are typically manufactured
using an electrical discharge machining (EDM) process. However,
such a manufacturing process may increase manufacturing costs and
times, and because of the complexity of the task may cause airfoil
vanes to be reworked. A nozzle design including an internal cooling
geometry that is comparable with the investment casting process
generally is less expensive to manufacture relative to a nozzle
design that requires the EDM process to produce the slots.
BRIEF SUMMARY OF THE INVENTION
[0005] In one aspect, a method for fabricating a nozzle for a gas
turbine engine is provided. The nozzle includes an airfoil. The
method comprises forming the airfoil to include a suction side and
a pressure side connected at a leading edge and a trailing edge,
forming a plurality of slots in the pressure side of the airfoil
extending towards the trailing edge, and extending a plurality of
pins arranged in rows between the airfoil suction and pressure
sides, such that at each of a first row of pins has a substantially
elliptical cross-sectional area, wherein the first row of pins is
between the remaining rows of pins and the airfoil pressure side
slots.
[0006] In another aspect of the invention, a turbine nozzle for a
gas turbine engine is provided. The nozzle includes an airfoil vane
including a first wall, a second wall, and a plurality of pins
extending therebetween. The first and second walls are connected at
a leading edge and a trailing edge. The first wall includes a
plurality of slots that extend towards the trailing edge. The
plurality of pins include at least a first row of pins which have a
substantially elliptical cross-sectional area. The first row of
pins is positioned between the remaining plurality of pins and the
first wall slots.
[0007] In a further aspect, an airfoil for a gas turbine engine
nozzle is provided. The airfoil includes a root, a tip, a plurality
of pins, a convex sidewall and a concave sidewall connected at a
trailing edge. Each of the sidewalls extends between the root and
tip. The convex sidewall defines a pressure side of the airfoil and
includes a plurality of slots that extend towards the trailing
edge. The plurality of pins include at least a first row of pins
and a second row of pins. The first row of pins are concentrically
aligned radially and each of the first row pins has a substantially
elliptical cross sectional profile and is tapered such that an
upstream side of each first row pin has a width that is greater
than a downstream side of each first row pin. The slots are
adjacent to and downstream from the first row of pins.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] FIG. 1 is a schematic illustration of a gas turbine
engine;
[0009] FIG. 2 is a perspective view of a turbine nozzle assembly
that may be used with the gas turbine engine shown in FIG. 1;
and
[0010] FIG. 3 is an enlarged schematic cross-sectional view of a
turbine nozzle shown in FIG. 2.
DETAILED DESCRIPTION OF THE INVENTION
[0011] FIG. 1 is a schematic illustration of a gas turbine engine
10 including a fan assembly 12, a high-pressure compressor 14, and
a combustor 16. Engine 10 also includes a high-pressure turbine 18
and a low-pressure turbine 20. Engine 10 has an intake, or
upstream, side 28 and an exhaust, or downstream, side 30. In one
embodiment, engine 10 is a CF6-80 engine commercially available
from General Electric Aircraft Engines, Cincinnati, Ohio.
[0012] In operation, air flows through fan assembly 12 and
compressed air is supplied to high-pressure compressor 14. The
highly compressed air is delivered to combustor 16. Airflow from
combustor 16 is discharged through a turbine nozzle assembly (not
shown in FIG. 1) that includes a plurality of nozzles (not shown in
FIG. 1) and used to drive turbines 18 and 20. Turbine 20, in turn,
drives fan assembly 12, and turbine 18 drives high-pressure
compressor 14.
[0013] FIG. 2 is a perspective view of turbine nozzle assembly 50
that may be used with a gas turbine engine, such as engine 10
(shown in FIG. 1). Nozzle assembly 50 includes a plurality of
nozzles 51 which extend circumferentially within engine 10. FIG. 3
is an enlarged schematic cross-sectional view of a portion of a
turbine nozzle 51 used with nozzle assembly 50. Each nozzle 51
includes an airfoil vane 52 that extends between a radially outer
band or platform 54 and a radially inner band or platform 56. More
specifically, in the exemplary embodiment, each band 54 and 56 is
formed integrally with each airfoil vane 52.
[0014] Each airfoil vane 52 includes a first sidewall 60 and a
second sidewall 62. First sidewall 60 is convex and defines a
suction side of airfoil vane 52, and second sidewall 62 is concave
and defines a pressure side of airfoil vane 52. Sidewalls 60 and 62
are joined at a leading edge 64 and at an axially-spaced trailing
edge 66 of airfoil vane 52.
[0015] First and second sidewalls 60 and 62, respectively, extend
longitudinally, or radially outwardly, in span from radially inner
band 56 to radially outer band 54. An airfoil root 70 is defined as
being adjacent inner band 56, and an airfoil tip 72 is defined as
being adjacent outer band 54. Additionally, first and second
sidewalls 60 and 62, respectively, define a cooling chamber (not
shown) within airfoil vane 52. More specifically, the cooling
chamber is bounded by an inner surface (not shown) of each
respective sidewall 60 and 62.
[0016] Second sidewall 62 defines a pressure side of airfoil vane
52 and includes a plurality of cooling openings 80 extending
therethrough and in flow communication with the cooling chamber. In
one embodiment, the nozzle cooling chamber receives cooling air
discharged from a compressor, such as compressor 14 (shown in FIG.
1). Openings 80 discharge cooling air onto an external surface 81
of airfoil vane 52. More specifically, film cooling openings 80
include trailing edge diffusion slots 82 and a plurality sidewall
film cooling openings 86. It should be noted that the arrangement
and number of openings 80 is variable.
[0017] Diffusion slots 82 extend through second sidewall 62 and
extend towards airfoil trailing edge 66, and include a pair of
identical first slots 90 and 92 adjacent each respective platform
band 54 and 56, and a plurality of identical body slots 94 spaced
radially between slots 90 and 92. Adjacent diffusion slots 82 are
separated by land areas 98, such that a slot radial height is
defined for each slot 82 between adjacent land areas 98.
Accordingly, because slots 94 are identical, a radial height 100 of
each slot 94 is the same, and because slots 90 and 92 are
identical, a radial height 102 of each slot 90 and 92 is the same,
and is larger than slot radial height 100. More specifically, each
slot 90 and 92 has a radial height 102 that is approximately twice
that of a radial height 100 of each slot 94. In the exemplary
embodiment, each slot 90 and 92 has a radial height 102 equal
approximately 0.090 inches, and each slot 94 has a radial height
100 equal approximately 0.045 inches. The difference in radial
heights 102 and 100 facilitates relieving thermal strain at airfoil
trailing edge 66 due to a mismatch in thermal growth between
airfoil vane 52 and bands 54 and 56. Furthermore, the slot geometry
facilitates enhancing cooling in airfoil vane 52 adjacent each band
54 and 56.
[0018] Each land area 98 is divergent and includes an upstream
portion 120 and a downstream portion 122. Upstream portion 120 is
curved and extends between a leading edge 124 and downstream
portion 122. In the exemplary embodiment, upstream portion 120 is
substantially semi-circular. Downstream portion 122 extends from
upstream portion 120 to an apex 126 at a trailing edge 128.
Downstream portion 122 is substantially triangular-shaped and
defines a half angle of diffusion 129 for each diffusion slot 82.
In the exemplary embodiment, each slot 82 has a half angle of
diffusion 129 that is equal approximately seven degrees.
[0019] Each land area 98 also has a length 130 measured between
each land area leading edge 124 and each respective land area
trailing edge 126. In the exemplary embodiment, each land area has
a length 130 that is approximately equal 0.155 inches. Land areas
98 are also equi-spaced and as such, each slot 82 has a centerline
136. Slot centerlines 136 are substantially parallel.
[0020] Airfoil vane 52 also includes a pin bank 144 that includes a
plurality of pins 150 that extend through the cooling chamber
between sidewalls 60 and 62. Pin bank 144 facilitates inducing
turbidity into cooling air flowing through the cooling chamber.
More specifically, in the exemplary embodiment, each vane 52
includes at least a row of first pins 152 that are immediately
upstream from diffusion slots 82, between remaining pins 150 and
slots 82. Pins 152 extend radially between platform bands 54 and
56, and are radially-aligned.
[0021] Each pin 150 has a substantially elliptical cross-sectional
area and is tapered towards airfoil vane trailing edge 66. Tapering
each pin 150 facilitates enhancing a local velocity of cooling air
flowing through the cooling chamber towards diffusion slots 82, and
thus, facilitates increasing cooling effectiveness within airfoil
vane 52 between pins 150 and airfoil vane trailing edge 66. More
specifically, each pin 150 has a radial width 154 at a pin leading
edge 156 that is greater than a radial width 158 at a pin trailing
edge 160. Tapering pins 150 facilitates reducing pressure losses of
cooling air flowing through the cooling chamber. Furthermore, the
geometry configuration within nozzle vane 52 enables each land area
98 to be fabricated with a shorter length 130 than other known
nozzles that includes trailing edge diffusion slots. The shorter
length 130 of each land area 98 facilitates each airfoil vane 52
being manufactured with an investment casting process.
[0022] Each pin 150 also has a centerline 170 that extends between
leading and trailing edges 156 and 160, respectively. Pins 150
extending immediately upstream from body diffusion slots 94 are
positioned such that each pin centerline 170 is substantially
co-linear with each slot centerline 136, and are thus parallel with
centerlines 136 extending through slots 90 and 92.
[0023] Airfoil vanes 52 are fabricated by casting a core (not
shown). The core is fabricated by injecting a liquid ceramic and
graphite slurry into a core die (not shown), and the slurry is
heated to form a solid ceramic airfoil core. The airfoil core is
suspended in an airfoil die (not shown) and hot wax is injected
into the airfoil vane die to surround the ceramic airfoil core. The
hot wax solidifies and forms a wax airfoil vane with the ceramic
core suspended in the airfoil vane.
[0024] The wax airfoil vane with the ceramic core is then inserted
into the airfoil vane die and molten metal is injected into the
die. The molten metal melts, and takes the place of, the wax
airfoil vane, and forms a metal airfoil vane with the ceramic core
remaining in place. The airfoil vane is then cooled, and the
ceramic core removed. Because nozzles 51 can be fabricated with an
investment casting process, manufacturing costs associated with
nozzles 51 are reduced in comparison to other known nozzles.
[0025] In operation, cooling air is supplied to each into each
airfoil vane cooling chamber from a high pressure compressor. The
cooling air is channeled through selectively turbulated cavities
(not shown) and through a pin bank cooling array 144 upstream from
trailing edge diffusion slots 82. The cooling air is also
discharged through the airfoil cooling openings 80 to provide a
buffer layer of cooling air to airfoil external surface 81. This
buffer layer, or film layer, is substantially cooler than combustor
discharge air passing through nozzle 51. Airfoil trailing edge 66
and the surrounding trailing edge region, are cooled by cooling air
flowing through pin bank 144 and through diffusion slots 82. The
combination of pins 150 and slots 90, 92, and 94 facilitate
providing balanced airfoil cooling while minimizing a flow level of
cooling air passing through airfoil vane 52. More specifically,
pins 150 and slots 82 facilitate reducing nozzle metal temperatures
to extend a useful life of nozzles 50.
[0026] The above-described turbine nozzle include a geometry
configuration that facilitates extending a useful life of the
turbine nozzle. Each nozzle airfoil vane includes a row of
diffusion slots that include an oversized slot adjacent each
platform band. The oversized slots facilitate cooling the regions
of the airfoil vane adjacent the platforms. Furthermore, each
nozzle airfoil vane also includes a row of tapered pins that are
immediately upstream and axially-aligned with respect to each
respective slot. The pins facilitate enhanced cooling and flow
regulation with improved manufacturing yeilds. As a result, the
geometry configuration of the nozzles facilitate extending a useful
life of the nozzles in a cost-effective and reliable manner.
[0027] While the invention has been described in terms of various
specific embodiments, those skilled in the art will recognize that
the invention can be practiced with modification within the spirit
and scope of the claims.
* * * * *