U.S. patent application number 14/728568 was filed with the patent office on 2015-12-10 for gas turbine engine airfoil platform cooling.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Wieslaw A. Chlus, Seth J. Thomen.
Application Number | 20150354369 14/728568 |
Document ID | / |
Family ID | 53284103 |
Filed Date | 2015-12-10 |
United States Patent
Application |
20150354369 |
Kind Code |
A1 |
Chlus; Wieslaw A. ; et
al. |
December 10, 2015 |
GAS TURBINE ENGINE AIRFOIL PLATFORM COOLING
Abstract
An airfoil for a gas turbine engine includes an airfoil that
extends from a platform that has first and second circumferential
sides that respectively extend to first and second circumferential
edges. The first circumferential side has a tapered surface at a
first angle relative to a flow path surface. The second
circumferential surface has a cooling hole that extends toward the
second lateral edge at a second angle relative to the flow path
surface. The tapered surface and the cooling hole are axially
aligned with one another.
Inventors: |
Chlus; Wieslaw A.;
(Wethersfield, CT) ; Thomen; Seth J.; (Colchester,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Hartford |
CT |
US |
|
|
Family ID: |
53284103 |
Appl. No.: |
14/728568 |
Filed: |
June 2, 2015 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
62008599 |
Jun 6, 2014 |
|
|
|
Current U.S.
Class: |
416/96R |
Current CPC
Class: |
F05D 2250/312 20130101;
F05D 2250/314 20130101; Y02T 50/672 20130101; F05D 2250/292
20130101; F05D 2250/322 20130101; F05D 2260/202 20130101; F01D
5/143 20130101; Y02T 50/676 20130101; Y02T 50/60 20130101; F01D
5/186 20130101; F01D 5/187 20130101; Y02T 50/673 20130101 |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Claims
1. An airfoil for a gas turbine engine comprising: an airfoil
extending from a platform that has first and second circumferential
sides respectively extending to first and second circumferential
edges, the first circumferential side having a tapered surface at a
first angle relative to a flow path surface, and the second
circumferential surface having a cooling hole extending toward the
second lateral edge at a second angle relative to the flow path
surface, the tapered surface and the cooling hole axially aligned
with one another.
2. The airfoil according to claim 1, wherein the airfoil includes a
cooling passage, the cooling hole in fluid communication with the
cooling hole.
3. The airfoil according to claim 2, wherein the second angle is
5-40.degree..
4. The airfoil according to claim 3, wherein the second angle is
15-30.degree..
5. The airfoil according to claim 2, comprising multiple cooling
holes arranged in a cluster, the cluster arranged near a trailing
edge of the airfoil on a pressure side.
6. The airfoil according to claim 5, wherein the cluster is
arranged within about three inches (76.2 mm) of an aft edge of the
platform.
7. The airfoil according to claim 5, wherein the cluster is within
about 0.6 inch (15.2 mm) of the second lateral edge.
8. The airfoil according to claim 5, wherein the cooling holes each
have a diameter equivalent of 0.010-0.050 inch (0.25-1.27 mm).
9. The airfoil according to claim 1, wherein the first angle is
1-20.degree..
10. The airfoil according to claim 9, wherein the first angle is
2-15.degree..
11. The airfoil according to claim 9, wherein the tapered surface
extends within about three inches (76.2 mm) of the first
circumferential edge to an aft edge of the platform.
12. The airfoil according to claim 11, wherein the tapered surface
extends less that 0.7 inch (17.78 mm) from the first
circumferential edge.
13. The airfoil according to claim 1, wherein the airfoil is a
turbine blade.
14. An array of airfoils for a gas turbine engine comprising:
adjacent airfoils, each airfoil extending from a platform that has
first and second circumferential sides respectively extending to
first and second circumferential edges, the first circumferential
side having a tapered surface at a first angle relative to a flow
path surface, and the second circumferential surface having a
cooling hole extending toward the second lateral edge at a second
angle relative to the flow path surface, the tapered surface and
the cooling hole axially aligned with one another.
15. The array of airfoils according to claim 14, wherein the
airfoils includes a cooling passage, the cooling hole in fluid
communication with the cooling passage.
16. The array of airfoils according to claim 15, wherein the second
angle is 5-40.degree., and comprising multiple cooling holes
arranged in a cluster, the cluster arranged near a trailing edge of
the airfoil on a pressure side.
17. The array of airfoils according to claim 16, wherein the
cluster is arranged within about three inches (76.2 mm) of an aft
edge of the platform, the cluster is within about 0.6 inch (15.2
mm) of the second lateral edge, wherein the cooling holes each have
a diameter equivalent of 0.010-0.050 inch (0.25-1.27 mm).
18. The array of airfoils according to claim 16, wherein the first
angle is 1-20.degree..
19. The array of airfoils according to claim 18, wherein the
tapered surface extends within about three inches (76.2 mm) of the
first circumferential edge to an aft edge of the platform, wherein
the tapered surface extends less that 0.7 inch (17.78 mm) from the
first circumferential edge.
20. The array of airfoils according to claim 14, wherein the
airfoil is a turbine blade.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application claims priority to U.S. Provisional
Application No. 62/008,599 which was filed on Jun. 6, 2014 and is
incorporated herein by reference.
BACKGROUND
[0002] This disclosure relates to a gas turbine engine airfoil, and
more particularly, a platform cooling arrangement.
[0003] Industrial gas turbine engines include a compressor section,
a combustor section and a turbine section. Air entering the
compressor section is compressed and delivered into the combustor
section where it is mixed with fuel and ignited to generate a
high-speed exhaust gas flow. The high-speed exhaust gas flow
expands through the turbine section to drive the compressor and a
power turbine. The compressor and turbine sections each included
multiple circumferential arrays of blades and vanes.
[0004] The turbine section in particular is subject to high
temperatures that may exceed the melting temperature of the
components. To this end, these components are cooled by one or more
cooling mechanisms. Airfoils extend from a platform, and in the
case of a blade, an inner platform supported by a root section. The
airfoil and platform typically included cooling holes to supply
cooling fluid to the hotter areas of the blade.
[0005] As gas turbine engines are pushed to higher temperatures to
increase power output and efficiency, distress of the airfoil
platforms increasingly becomes the service life limiting area. A
typical solution to this includes decreasing the platform metal
temperatures with cooling air. One approach includes providing
platform cooling holes supplied with "wheel-space air," which
corresponds to fluid provided between adjacent turbine blade
shanks. This approach may supply insufficient pressure needed for
adequate film cooling to the local and adjacent platform. Another
approach includes providing cooling to the platform mate faces, or
facing edges, supplied by cooling air from the airfoil core and/or
"wheel-space air." This approach may not provide film cooling and
subjects the area to cracking due to reduced wall thickness.
SUMMARY
[0006] In one exemplary embodiment, an airfoil for a gas turbine
engine includes an airfoil that extends from a platform that has
first and second circumferential sides that respectively extend to
first and second circumferential edges. The first circumferential
side has a tapered surface at a first angle relative to a flow path
surface. The second circumferential surface has a cooling hole that
extends toward the second lateral edge at a second angle relative
to the flow path surface. The tapered surface and the cooling hole
are axially aligned with one another.
[0007] In a further embodiment of the above, the airfoil includes a
cooling passage. The cooling hole is in fluid communication with
the cooling hole.
[0008] In a further embodiment of any of the above, the second
angle is 5-40.degree..
[0009] In a further embodiment of any of the above, the second
angle is 15-30.degree..
[0010] In a further embodiment of any of the above, multiple
cooling holes are arranged in a cluster. The cluster is arranged
near a trailing edge of the airfoil on a pressure side.
[0011] In a further embodiment of any of the above, the cluster is
arranged within about three inches (76.2 mm) of an aft edge of the
platform.
[0012] In a further embodiment of any of the above, the cluster is
within about 0.6 inch (15.2 mm) of the second lateral edge.
[0013] In a further embodiment of any of the above, the cooling
holes each have a diameter equivalent of 0.010-0.050 inch
(0.25-1.27 mm).
[0014] In a further embodiment of any of the above, the first angle
is 1-20.degree..
[0015] In a further embodiment of any of the above, the first angle
is 2-15.degree..
[0016] In a further embodiment of any of the above, the tapered
surface extends within about three inches (76.2 mm) of the first
circumferential edge to an aft edge of the platform.
[0017] In a further embodiment of any of the above, the tapered
surface extends less that 0.7 inch (17.78 mm) from the first
circumferential edge.
[0018] In a further embodiment of any of the above, the airfoil is
a turbine blade.
[0019] In another exemplary embodiment, an array of airfoils for a
gas turbine engine includes adjacent airfoils. Each airfoil extends
from a platform that has first and second circumferential sides
respectively that extend to first and second circumferential edges.
The first circumferential side has a tapered surface at a first
angle relative to a flow path surface. The second circumferential
surface has a cooling hole that extends toward the second lateral
edge at a second angle relative to the flow path surface. The
tapered surface and the cooling hole are axially aligned with one
another.
[0020] In a further embodiment of the above, the airfoils include a
cooling passage. The cooling hole is in fluid communication with
the cooling passage.
[0021] In a further embodiment of any of the above, the second
angle is 5-40.degree. and comprises multiple cooling holes that are
arranged in a cluster. The cluster is arranged near a trailing edge
of the airfoil on a pressure side.
[0022] In a further embodiment of any of the above, the cluster is
arranged within about three inches (76.2 mm) of an aft edge of the
platform. The cluster is within about 0.6 inch (15.2 mm) of the
second lateral edge. The cooling holes each have a diameter
equivalent of 0.010-0.050 inch (0.25-1.27 mm).
[0023] In a further embodiment of any of the above, the first angle
is 1-20.degree..
[0024] In a further embodiment of any of the above, the tapered
surface extends within about three inches (76.2 mm) of the first
circumferential edge to an aft edge of the platform. The tapered
surface extends less that 0.7 inch (17.78 mm) from the first
circumferential edge.
[0025] In a further embodiment of any of the above, the airfoil is
a turbine blade.
BRIEF DESCRIPTION OF THE DRAWINGS
[0026] The disclosure can be further understood by reference to the
following detailed description when considered in connection with
the accompanying drawings wherein:
[0027] FIG. 1 is a schematic cross-sectional view of an example
industrial gas turbine engine.
[0028] FIG. 2 schematically illustrates a section of the gas
turbine engine, such as a turbine section.
[0029] FIGS. 3A and 3B are perspective and elevational views
respectively of adjacent blades.
[0030] FIG. 4 is an elevational view of the blade shown in FIGS. 3A
and 3B.
[0031] FIG. 5 is an enlarged cross-sectional view of the adjacent
blades.
[0032] The embodiments, examples and alternatives of the preceding
paragraphs, the claims, or the following description and drawings,
including any of their various aspects or respective individual
features, may be taken independently or in any combination.
Features described in connection with one embodiment are applicable
to all embodiments, unless such features are incompatible.
DETAILED DESCRIPTION
[0033] A schematic view of an industrial gas turbine engine 10 is
illustrated in FIG. 1. The engine 10 includes a compressor section
12 and a turbine section 14 interconnected to one another by a
shaft 16 rotatable about an axis X. A combustor 18 is arranged
between the compressor and turbine sections 12, 14. A generator 22
is rotationally driven by a shaft coupled to the turbine or
uncoupled via a power turbine 20, which is connected to a power
grid 23. It should be understood that the illustrated engine 10 is
highly schematic, and may vary from the configuration illustrated.
Moreover, the disclosed airfoil may be used in commercial and
military aircraft engines as well as industrial gas turbine
engines.
[0034] The turbine section 14 includes multiple turbine blades, one
of which is illustrated at 64 in FIG. 2. In the example turbine
section 14, first and second arrays of circumferentially spaced
fixed vanes 60, 62 are axially spaced apart from one another. A
first stage array of circumferentially spaced turbine blades 64,
mounted to a rotor disk 68, is arranged axially between the first
and second fixed vane arrays. A second stage array of
circumferentially spaced turbine blades 66 is arranged aft of the
second array of fixed vanes 62. It should be understood that any
number of stages may be used. Moreover, the disclosed airfoil may
be used in a compressor section, turbine section and/or fixed or
rotating stages.
[0035] The turbine blades each include a tip 80 adjacent to a blade
outer air seal 70 of a case structure 72, which provides an outer
flow path. The first and second stage arrays of turbine vanes and
first and second stage arrays of turbine blades are arranged within
a core flow path C and are operatively connected to the shaft 16,
for example.
[0036] Each blade 64 includes an inner platform 76 respectively
defining an inner flow path. The platform inner platform 76
supports an airfoil 78 that extends in a radial direction R. It
should be understood that the turbine blades may be discrete from
one another or arranged in integrated clusters. The airfoil 78
provides leading and trailing edges 82, 84.
[0037] The airfoil 78 is provided between pressure (typically
concave) and suction (typically convex) sides in circumferential
direction Y (FIG. 4) provided between the leading and trailing
edges 82, 84. The turbine blades 64 are constructed from a high
strength, heat resistant material such as a nickel-based or
cobalt-based superalloy, or of a high temperature, stress resistant
ceramic or composite material. In cooled configurations, internal
fluid passages and external cooling apertures provide for a
combination of impingement and film cooling. Other cooling
approaches may be used such as trip strips, pedestals or other
convective cooling techniques. In addition, one or more thermal
barrier coatings, abrasion-resistant coatings or other protective
coatings may be applied to the turbine vane 64.
[0038] The airfoil 78 extends from the platform 76 and provides
first and second circumferential sides, which corresponds to
pressure and suction sides 86, 88, as shown in Figure 4. With
continued reference to FIG. 4, the first and second circumferential
sides 86, 88 include first and second circumferential edges 92, 94,
respectively. This first circumferential side 86 has a tapered
surface 90 at a first angle 100 relative to the flowpath surface
provided by the platform 76. The first angle is 1-20 degrees, and
in another example, 2-15 degrees. In still another example, the
first angle is 2-12 degrees.
[0039] The tapered surface 90 extends within about three inches
(76.2 mm) of the first circumferential edge 92 to an aft edge 120
of the platform 76. The tapered surface 90 extends a width 108 less
that 0.7 inch (17.78 mm) from the first edge 92.
[0040] The second circumferential surface 88 has at least one
cooling hole 98, for example, a cluster of cooling holes, extending
toward the second circumferential edge 92 at a second angle 96
relative to the flowpath surface. As shown in FIG. 5, the airfoil
78 includes a cooling passage 99 in fluid communication with the
cooling hole 98. In one example, the second angle 96 is 5-40
degrees, in another example, the second angle 96 is 15-30
degrees.
[0041] The tapered surface 90 and the cooling hole 98 are axially
aligned with one another such that cooling fluid from the cooling
hole 98 is directed toward the tapered surface 90. The cooling
arrangement provides for a more effective platform cooling. The
relationship between these features and adjacent blades is shown in
FIGS. 3A-3B.
[0042] The cluster of cooling holes 98 is arranged within about 3
inches (76.2 mm) of an aft edge 120 of the platform 76. The cluster
is within about 0.60 inch (15.2 mm) of the second lateral edge 90.
A lateral offset 110 is about 0.60 inch (15.2 mm), and a lateral
width 112 is at least 0.010 inch (0.25 mm). Each cooling hole 98
has an effective diameter, or diameter equivalent of 0.010-0.050
inch (0.25 -1.27 mm). The cluster includes an axial offset 114 from
the aft edge 120 of about 1 inch (25.4 mm) and may extend a
distance 116, which is less than three inches (76.2 mm). The
cooling holes may be round or shaped. The holes may have a uniform
cross-section or may be shaped as diffusers.
[0043] A seal 118 may be provided between adjacent blades 64 to
obstruct the gap provided between the facing circumferential edges
92, 94. The seal 118 is schematically illustrated and may be
provided using any suitable arrangement.
[0044] It should also be understood that although a particular
component arrangement is disclosed in the illustrated embodiment,
other arrangements will benefit herefrom. Although particular step
sequences are shown, described, and claimed, it should be
understood that steps may be performed in any order, separated or
combined unless otherwise indicated and will still benefit from the
present invention.
[0045] Although the different examples have specific components
shown in the illustrations, embodiments of this invention are not
limited to those particular combinations. It is possible to use
some of the components or features from one of the examples in
combination with features or components from another one of the
examples.
[0046] Although an example embodiment has been disclosed, a worker
of ordinary skill in this art would recognize that certain
modifications would come within the scope of the claims. For that
reason, the following claims should be studied to determine their
true scope and content.
* * * * *