U.S. patent application number 12/486939 was filed with the patent office on 2010-12-23 for turbine blade having platform cooling holes.
Invention is credited to Michel P. Arnal, Gregory M. Nadvit, Andrew D. Williams.
Application Number | 20100322767 12/486939 |
Document ID | / |
Family ID | 43354549 |
Filed Date | 2010-12-23 |
United States Patent
Application |
20100322767 |
Kind Code |
A1 |
Nadvit; Gregory M. ; et
al. |
December 23, 2010 |
Turbine Blade Having Platform Cooling Holes
Abstract
A turbine blade having a plurality of cooling holes which extend
from an outside edge of the platform to a cooling passage formed
within the turbine blade and a method of limiting the formation of
cracks in the platform of the blade are provided. The plurality of
cooling holes in the platform are formed at an approximate angle of
45.degree. to the outside edge of the platform and are formed at
the approximate mid-point of the thickness of the platform. The
cooling holes are generally cylindrical in shape and have a
diameter of approximately 50% of the platform thickness.
Inventors: |
Nadvit; Gregory M.; (Grand
Prairie, TX) ; Williams; Andrew D.; (Balcraig,
GB) ; Arnal; Michel P.; (Turgi, CH) |
Correspondence
Address: |
Baker Botts L.L.P
910 Louisiana Street, One Shell Plaza
HOUSTON
TX
77002
US
|
Family ID: |
43354549 |
Appl. No.: |
12/486939 |
Filed: |
June 18, 2009 |
Current U.S.
Class: |
416/1 ;
416/97R |
Current CPC
Class: |
F05D 2240/81 20130101;
F01D 5/186 20130101; F05B 2240/801 20130101 |
Class at
Publication: |
416/1 ;
416/97.R |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Claims
1. A turbine blade having an airfoil connected to a platform in a
root region, the airfoil and platform having a common cooling
passage formed therein, the platform comprising at least one
cooling hole which extends from an outside edge of the platform to
the common cooling passage.
2. The turbine blade according to claim 1, wherein a plurality of
cooling holes are formed in the platform which extend from the
outside edge of the platform to the common cooling passage.
3. The turbine blade according to claim 2, wherein four cooling
holes are formed in the platform which extend from the outside edge
of the platform to the common cooling passage.
4. The turbine blade according to claim 3, wherein the plurality of
cooling holes are formed at an angle to the outside edge of the
platform which is less than 90.degree..
5. The turbine blade according to claim 4, wherein the angle at
which the plurality of cooling holes are formed to the outside edge
of the platform is approximately 45.degree..
6. The turbine blade according to claim 5, wherein the common
cooling passage includes a serpentine cooling circuit.
7. The turbine blade according to claim 6, wherein each of the
cooling holes extends to a distinct pathway within the serpentine
cooling circuit.
8. The turbine blade according to claim 5, wherein the common
cooling passage includes a plurality of generally parallel cooling
veins extending through the platform and airfoil.
9. The turbine blade according to claim 8, wherein each of the
cooling holes extends to a distinct parallel cooling vein.
10. The turbine blade according to claim 1, wherein the platform
has a defined thickness and the at least one cooling hole is formed
at the approximate mid-point of the thickness.
11. The turbine blade according to claim 1, wherein the at least
one cooling hole is generally cylindrical in shape.
12. The turbine blade according to claim 1, wherein the platform
has a defined thickness, and the at least one cooling hole has a
diameter of approximately 50% of the platform thickness.
13. A method of limiting damage to a platform of a turbine blade
having an airfoil connected to the platform in a root region of the
airfoil blade, the airfoil and platform having a common cooling
passage formed therein, the method comprising the step of forming
at least one cooling hole in the platform which extends from an
outside edge of the platform to the common cooling passage.
14. The method according to claim 13, wherein a plurality of
cooling holes are formed in the platform which extend from the
outside edge of the platform to the common cooling passage.
15. The method according to claim 14, wherein four cooling holes
are formed in the platform which extend from the outside edge of
the platform to the common cooling passage.
16. The method according to claim 15, wherein the plurality of
cooling holes are formed at an angle to the outside edge of the
platform which is less than 90.degree..
17. The method according to claim 16, wherein the angle at which
the plurality of cooling holes are formed to the outside edge of
the platform is approximately 45.degree..
18. The method according to claim 17, wherein the cooling passage
includes a serpentine cooling circuit.
19. The method according to claim 18, wherein each of the cooling
holes extends to a distinct pathway within the serpentine cooling
circuit.
20. The method according to claim 17, wherein the cooling passage
includes a plurality of generally parallel cooling veins extending
through the platform and airfoil.
21. The method according to claim 18, wherein each of the cooling
holes extends to a distinct parallel cooling vein.
22. The method according to claim 13, wherein the platform has a
defined thickness and the at least one cooling hole is formed at
the approximate mid-point of the thickness.
23. The method according to claim 13, wherein the platform has a
defined thickness and the at least one cooling hole is generally
cylindrical in shape and has a diameter of approximately 50% of the
platform thickness.
24. The method according to claim 13, wherein the at least one
cooling hole is formed by an EDM process.
25. The method according to claim 13, wherein the step of forming
at least one cooling hole is performed in the course of forming a
new turbine blade.
26. The method according to claim 13, wherein the step of forming
at least one cooling hole is performed in the course of repairing a
turbine blade that has previously been in service.
27. The method according to claim 25, wherein the step of forming
at least one cooling hole is performed without removing any TBC
layer formed on the turbine blade.
Description
FIELD OF THE INVENTION
[0001] The present invention relates generally to techniques for
reducing cracks in platforms of gas turbine blades and more
specifically to a turbine blade having a plurality of cooling holes
formed in the platform.
BACKGROUND
[0002] The turbine section of gas turbine engines typically
comprises multiple sets or stages of stationary airfoils, known as
nozzles or vanes, and moving airfoils, known as rotor blades or
buckets. FIG. 1 illustrates a typical turbine rotor blade or
turbine blade 100 found in the first stage of the turbine section,
which is the section immediately adjacent to the combustion section
of the gas turbine and thus is in the region of the turbine section
that is exposed to the highest temperatures. Known problems with
such blades 100 are the formation of multiple cracks and, when
present, the delamination of a thermal barrier coating (TBC) in the
platform region 104 due to the heat stresses in this region of the
blade. In some cases the cracking is so severe that it results in
breakage and separation of a substantial portion of the platform
104 on the pressure side of the blade 100, leading to the early
retirement of the blade. In order to prevent this early retirement
and to extend blade operational lifetime, various approaches have
been proposed.
[0003] In one such solution, an undercut is machined into the blade
platform along the pressure side of the blade. This proposed
solution purports to reduce the total stress level in this region
of high stress. This approach has been implemented on both turbine
and compressor blades as both a field repair and a design
modification. If a stress reduction is achieved in the platform
region, the concern is whether the undercut results in a high
stress within the grooved region where material is removed. In
other words, the success of the strategy depends on whether a
stress reduction in an existing high-stress region can be achieved
without creating a new area of high stress within the blade.
[0004] There are two primary concerns raised with platform
undercuts. First, whether the undercut will be effective in
reducing the stress in the platform. Second, whether the stress
concentration occurring in the undercut will be so high that it
offsets the benefit of the undercut to the platform region.
Undercut solutions have had difficulty striking a balance between
these two concerns. It is desired to have a solution which
effectively reduces the stress in the platform and, thereby, the
potential for formation of cracks, TBC delamination, and, in the
worst case, breakage and separation of significant portions of the
platform altogether, and which does not add additional stresses to
the blade. The present invention seeks to solve this problem.
SUMMARY
[0005] In one embodiment, the present invention is directed to a
turbine blade and limits platform cracking. The turbine blade has
of an airfoil connected to a platform in the blade root region. The
airfoil and the platform share a common cooling passage, which may
include one or more cooling channels or paths. The turbine blade
configuration limits platform damage, including but not limited to
cracking, removal of the TBC layer, and breakage and loss of blade
material, through the influence of at least one cooling hole which
extends from an outside edge of the platform, through the platform
and to the common cooling passage. In one embodiment, multiple
cooling holes are formed in the platform at an angle to the outside
edge of the platform which is approximately 90.degree. and in
another embodiment at approximately 45.degree.. In one embodiment,
the common cooling passage has one or more serpentine cooling
circuits, and each of the cooling holes extends to a distinct
channel within the serpentine cooling circuit. In another
embodiment, the common cooling passage includes a plurality of
generally parallel cooling veins extending through the platform and
airfoil to the airfoil tip, which may be formed by a Shaped-Tube
Electolytic Machining (STEM) drilling process for example. In this
case, each of the platform cooling holes extends through the
platform from the platform edge to a distinct parallel cooling vein
in the blade. In one embodiment, the platform cooling holes are
formed at the approximate midpoint of the defined thickness of the
platform. The cooling holes may be generally cylindrical in shape
with a diameter approximately 50% of the platform thickness.
[0006] In another embodiment, the present invention is directed to
a method of limiting damage to the platform of a turbine blade
having an airfoil connected to the platform in a blade root region.
The method includes the step of forming at least one cooling hole
in the platform which extends from an outside edge of the platform
to a cooling passage within the platform which passes through and
is shared with the airfoil. In one embodiment, multiple cooling
holes are formed in the platform. In a more specific embodiment,
there may be four cooling holes formed in the platform. The cooling
holes can be formed at the angles, having the location and
geometric dimensions as described in the immediately preceding
paragraph. The cooling holes can be formed by a number of known
processes, but in at least one embodiment is formed by an
electro-discharge machining (EDM) process.
[0007] The present invention has application in both the
manufacture of a new turbine blade as well as in the repair of an
existing turbine blade not having cooling holes formed in the
platform region thereof. In the latter case, the one or more
cooling holes can be formed without having to remove any existing
TBC layer that may have been applied on the blade.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] The following drawings form part of the present
specification and are included to further demonstrate certain
aspects of the present invention. The present invention may be
better understood by reference to one or more of these drawings in
combination with the description of embodiments presented herein.
However, the present invention is not intended to be limited by the
drawings.
[0009] FIG. 1 is a perspective view of a prior art turbine blade
having a portion of its platform eroded.
[0010] FIG. 2 is a perspective view of a turbine blade in
accordance with the present invention illustrating a plurality of
cooling holes formed in its platform.
[0011] FIG. 3 is a cross-sectional view of the platform of the
turbine blade taken across line 3-3 shown in FIG. 2 illustrating
the platform cooling holes of the present invention communicating
with the corresponding distinct cooling pathways of a serpentine
cooling circuit.
[0012] FIG. 4 is a cross-sectional view of the platform of the
turbine blade taken across the same line 3-3 shown in FIG. 2
illustrating the platform cooling holes of the present invention
communicating with a corresponding plurality of generally parallel
cooling veins formed in the airfoil and platform.
[0013] FIG. 5 is a cross-sectional view of a turbine blade with two
separate serpentine cooling passages in the airfoil, according to
various embodiments of the present invention.
[0014] FIGS. 6A and 6B are plots of the distribution of heat
transfer coefficient (HTC or film coefficient) along the leading
and trailing serpentine cooling circuits (shown in FIG. 5),
respectively, as a function of the distance from the air inlets at
the base of the blades, to each corresponding (leading or trailing)
cooling circuit, according to various embodiments of the present
invention.
[0015] FIGS. 7A through 7D are plots of the film coefficient and
cooling air temperature along the length of the platform cooling
holes as shown in FIG. 3, according to various embodiments of the
present invention. The film coefficients and temperatures are shown
as a function of distance from the point where the platform cooling
holes join the serpentine cooling circuits for each of the four
platform cooling holes.
[0016] FIGS. 8A and 8B are perspective views of the turbine blade
with (FIG. 8A) and without (FIG. 8B) the platform cooling holes
according to various embodiments of the present invention,
respectively, illustrating the metal temperature distribution on
the surface of the entire blade.
[0017] FIGS. 9A and 9B are perspective views of the turbine blade
with (FIG. 9A) and without (FIG. 9B) the platform cooling holes
according to various embodiments of the present invention,
respectively, illustrating the temperature distributions in the
region of these blades proximate the platform cooling holes. The
left hand illustration shows the blade platform in perspective view
from above and the right hand illustration shows the platform
temperatures looking from below in each of the two figures.
[0018] FIGS. 10A and 10B are perspective views of the turbine blade
with (FIG. 10A) and without (FIG. 10B) the platform cooling holes
according to various embodiments of the present invention,
respectively, illustrating the temperature distribution in the
region of the blade proximate the juncture of the platform and
trailing edge lowermost cooling hole.
[0019] FIGS. 11A and 11B are perspective views of the turbine blade
with (FIG. 11A) and without (FIG. 11B) the platform cooling holes
according to various embodiments of the present invention,
respectively, illustrating the equivalent stress distributions in
the platform region. The left hand illustration shows the sectioned
blade and platform looking down from above while the right hand
illustration shows the sectioned blade shank and platform looking
up from below, in each figure respectively.
[0020] FIGS. 12A and 12B are perspective views of the turbine blade
with (FIG. 12A) and without (FIG. 12B) the platform cooling holes
according to various embodiments of the present invention,
respectively, illustrating the axial stress distributions in the
platform region. The left hand illustration shows the sectioned
blade and platform looking down from above while the right hand
illustration shows the sectioned blade shank and platform looking
up from below, in each figure respectively.
[0021] FIGS. 13A and 13B are perspective views of the turbine blade
with (FIG. 13A) and without (FIG. 13B) the platform cooling holes
according to various embodiments of the present invention,
respectively, illustrating the stress distributions proximate the
juncture of the platform and lowermost, trailing edge cooling
hole.
DETAILED DESCRIPTION
[0022] The present invention will now be described with reference
to the following exemplary embodiments. Referring now to FIG. 2, a
turbine blade in accordance with the present invention is shown
generally by reference number 200. The turbine blade 200 has three
primary sections, a shank 202, which is designed to slide into a
disc on the shaft of the rotor (not shown); a platform 204
connected to the shank 202; and an airfoil 206 connected to the
platform 204. Platform 204 connects to shank 202 at a lower surface
of the platform 204, and to airfoil 206 at an upper surface of the
platform. Platform 204 has a thickness defined by the distance
between the lower surface and the upper surface. Moreover, platform
204 has four outside edges, which are generally orthogonal to the
lower surface and the upper surface. Generally, during the blade
200's initial manufacture, the shank 202, platform 204 and airfoil
206 are all cast as a single part.
[0023] The airfoil 206 is defined by a concave side wall 210, a
convex side wall 208, a leading edge 212 and an opposite trailing
edge 214; the leading and trailing edges being the two areas where
the concave side wall and convex side wall meet. The airfoil 206
has a root 216 which is proximate the platform 204 and a tip (or
shroud) 218 which is distal from the platform. As with prior art
turbine blades, air is supplied to the inside cavity (not shown) of
the airfoil 206 from the compressor to cool the airfoil. The
cooling air may exit through a plurality of cooling holes 220, at
least some of which may be formed in the trailing edge 214.
[0024] In accordance with the present invention, the platform 204
has a plurality of cooling holes 230 formed therein on the pressure
(concave) side of the airfoil 206, which is the region of the
platform that is susceptible to high stresses and often cracks,
including delaminating of coating, when present, and separation or
breakage of blade base material in extreme cases. In one
embodiment, four such cooling holes 230 are formed in the platform
204. The platform cooling holes 230 may be formed by an EDM
process. Alternatively, the platform cooling holes 230 can be
formed via STEM process or electro-chemical (ECM) drilling process
or other similar machining process. The process utilized to form
the cooling holes 230 may be selected to avoid removal of the TBC
layer formed on the turbine blade. In one embodiment, the platform
cooling holes 230 are generally cylindrical in shape, with center
axes generally parallel to the lower surface and the upper surface.
The cross-section of a platform cooling hole 230 at an outside edge
of the platform 204 may span approximately 50% of the platform
thickness, or the platform cooling holes 230 may have a diameter of
approximately 50% of the thickness of the platform 204. The
platform cooling holes 230 may also be formed at the approximate
mid-point of the thickness of the platform 204, i.e., the centers
of the cross-section of the platform cooling holes 230 at the
outside edge of the platform 204 are aligned at the mid-point of
the platform thickness so that an equal amount of platform material
is left above and below the platform cooling holes 230.
[0025] In one embodiment, the platform cooling holes 230 are formed
at an angle to the outside edge of the platform 204 into which they
are formed, which is best seen in FIG. 3. The platform cooling
holes 230 intersect the cooling cavity or passage 240, which
platform 204 shares with the airfoil 206 and which is fed by
cooling air from the compressor section of the turbine (not shown).
In the embodiment shown in FIG. 3, the common cooling passage 240
is defined by a pair of serpentine cooling circuits, namely a
leading serpentine cooling circuit 242 and a trailing serpentine
cooling circuit 244. In turn, each of the serpentine cooling
circuits is defined by a plurality of generally parallel channels
or pathways 246. The orientation and location of the serpentine
cooling circuits are shown in cross section in FIG. 5. As
illustrated in FIG. 3, the platform cooling holes 230 form an angle
a with the edge of the platform 204 which is approximately
45.degree.. Each of the platform cooling holes 230 is illustrated
as extending to, and communicating with, a distinct cooling pathway
246. The cooling air thus flows from the compressor to the turbine
blade 200 first through a cavity in the shank 202 (not shown), then
through the cooling pathways 246 of the serpentine cooling circuits
242, 244 and then through the platform cooling holes 230 before
exiting the turbine blade. As the cooling air flows through the
platform cooling holes 230 it cools the platform 204, thereby
preventing delamination of the TBC layer, formation of cracks, and,
worse, breakage and separation of the platform in that region
altogether.
[0026] In another embodiment (shown in FIG. 4), the platform
cooling holes 230 are formed in the platform 204 at an angle and
orientation within the thickness of the platform, but instead
extend to, and communicate with, a corresponding plurality of
generally parallel cooling veins 250 formed in the platform. In
other words, the common cooling passage 240 in FIG. 4 is a
plurality of discrete generally parallel cooling veins 250. The
cooling veins 250 may be formed by a number of processes, but
usually are formed by a STEM drilling process. The cooling veins
250 intersect a cavity (not shown) in the shank 202 of the turbine
blade 200, which is fed by cooling air from the compressor (also
not shown).
[0027] Without limiting the invention to a particular theory or
mechanism of action, it is nevertheless currently believed that the
overall cooling flow may increase and the internal cooling flow may
be re-distributed as a consequence of adding the platform cooling
holes 230. Table I lists the cooling mass flow which may occur as a
result of adding the platform cooling holes 230 to an example first
stage turbine blade with serpentine cooling passages.
TABLE-US-00001 TABLE I COMPARISON OF COOLING FLOW RATE Prior Art
Blade with Platform Blade Cooling Holes Difference Leading
Serpentine (lb.sub.m/hr) 453 456 +0.7% Trailing Serpentine
(lb.sub.m/hr) 512 518 +1.2% Total (lb.sub.m/hr) 965 974 +0.9%
[0028] As shown in the table, the cooling flow in the leading
serpentine cooling circuit may be .about.0.7% more than the prior
art blade configuration, and the cooling flow in the trailing
serpentine cooling circuit may increase by .about.1.2%. The total
cooling flow may increase only marginally (by .about.0.9%) with the
drilling of four platform cooling holes 230. The cooling flow of
the leading three platform cooling holes 230 may be 6.1, 5.8, and
6.5 pound mass per hour (lb.sub.m/hr), respectively. For the 4th
platform-cooling hole 230, which branches from the trailing
serpentine passage, the flow rate may be 6.1 lb.sub.m/hr. The total
platform cooling flow may be 24.4 lb.sub.m/hr, or about 2.5% of
total cooling flow available to the bucket.
[0029] FIG. 5 shows separate serpentine cooling passages in the
airfoil. The leading edge serpentine cooling circuit 252 cools the
leading, front half of the blade and receives its cooling air from
inlets 1 and 2, which are located at the base of the blade and lead
into cavity 256. The trailing edge serpentine cooling system 254
cools the trailing, back half of the blade and receives its cooling
air from inlets 3 and 4, which are located at the base of the blade
and lead into cavity 258.
[0030] FIGS. 6A and 6B show the distribution of heat transfer
coefficient (film coefficient) along the leading and trailing
serpentine circuits, respectively, according to one embodiment of
the invention. It can be seen that drilling four platform-cooling
holes 230 may have a minimal impact on the original cooling of the
main internal flow. The computed cooling flow parameters for each
platform cooling hole are shown in FIGS. 7A-7D, respectively.
[0031] The resulting surface temperature distributions of the
modified blade with platform cooling holes 230, according to one
embodiment of the invention, and a prior art blade are shown in
FIGS. 8A and 8B, respectively. As indicated by the results of the
cooling flow analysis, the thermal response in the airfoil above
the platform is basically unchanged when compared to the
temperature distribution of the original design configuration. As a
consequence of the insertion of four parallel platform cooling
holes 230, a substantial reduction of temperature was predicted in
the region encompassing the platform cooling holes 230. The peak
temperature predicted on the pressure side of the platform was
significantly reduced from approximately 1800.degree. F. for the
original design to 1600.degree. F. for the modified platform, e.g.
a drop of about 200.degree. F. This is illustrated in FIGS. 9A and
9B. As indicated by this drop, the platform cooling holes 230 are
effective as they extract fresh coolant air from the serpentine
cooling circuit and provide maximum coverage possible over the
pressure side region of the platform.
[0032] Further examining these results indicates that the benefit
of the proposed platform cooling strategy is likely to be at least
twofold. Through the additional convective cooling and conduction,
the gross reduction of the temperature in the platform region
should favorably lower the temperature gradients near the juncture
of platform and trailing edge lowermost cooling hole, which is
particularly susceptible to cracking, as indicated in FIGS. 10A and
10B. Temperatures near the trailing edge lowermost cooling hole may
be lowered by approximately 10.degree. F.
[0033] Equivalent and axial stress distributions of the blade
modified with platform cooling, according to one embodiment of the
invention, are plotted in FIGS. 11A and 11B and FIGS. 12A and 12B,
respectively. In the prior art turbine blades, there are large
compressive stresses induced by platform curling due to the
temperature gradients across the platform 204 and airfoil/shank
region (under steady load). The excessive compressive stress at
base load indicates a potential for substantial damage resulting
from the out-of-phase thermal-mechanical fatigue (TMF) that would
occur from each start-stop cycle. As shown in FIGS. 11A and 11B and
FIGS. 12A and 12B, overall stress levels on the pressure side of
the platform 204 may be reduced considerably by about 10-30%. At
the free edge, near the exit of the platform cooling holes 230, the
critical minimum principal stress may be reduced from 93 kilo-poind
per square inch (ksi) to 62 ksi as a result of platform cooling
modification. In the mid-span, the critical minimum principal
stress may decrease from 111 ksi to 100 ksi. This relatively mild
stress is localized and attributed to the thermal gradient across
the platform-cooling hole 230. Nevertheless, lowering the metal
temperature by 150.degree. F..about.200.degree. F. may
significantly enhance the associated fatigue properties and, hence,
increase the corresponding TMF life. TMF life may improve by as
much as 200% by taking into consideration the fatigue property
benefits resulting from the calculated temperature improvement
(Table II). In addition, with a much lower stress predicted at the
free edge near trailing edge of the blade, the fatigue crack
propagation life may improve substantially comparing to the
original design.
TABLE-US-00002 TABLE II COMPARISON OF STRESS RESULTS IN THE
PLATFORM Critical Min. Principal % Change Estimated % Change of
Stress (ksi) of Stress TMF Life Prior Art Blade 111 0% 0% Blade
with Platform 100 -10% +200%* Cooling Holes *taking into account
the temperature effect on TMF property
[0034] FIGS. 13A and 13B show the stress distribution in the
lowermost cooling hole region after the platform cooling
modification, according to one embodiment of the invention. As
illustrated in the plot, the lower thermal gradient near the
junction of airfoil trailing edge and platform favorably reduces
the stress at the critical location from 83 ksi to 76 ksi, or a
drop of about 8% (Table III). The corresponding TMF life may
increase by .about.100% as a consequence of the platform cooling
modification.
TABLE-US-00003 TABLE III COMPARISON OF STRESS RESULTS IN THE
LOWERMOST COOLING HOLE Critical Max. Principal % Change Estimated %
Change of Stress (ksi) of Stress TMF Life Prior Art Blade 83 0% 0%
Blade with Platform 76 -8% +100% Cooling Holes
[0035] In summary, the platform cooling hole modifications of the
present invention may be effective in both reducing the
temperatures and stresses in the cooled platform region. Moreover,
they may provide additional benefits in lowering the thermal
gradient near the juncture of platform and trailing edge, and
consequentially reduce the stress at the trailing edge lowermost
cooling hole. Based on a comparison to the results of the baseline
analysis, it is therefore considered as a viable design
modification--to be utilized in the course of forming a new turbine
blade--and/or recommended to implement during repair and
refurbishment of blades.
[0036] The terms "holes," "passages," "veins," "channels," and the
like are each used to describe conduits for the flow of air or
other cooling fluid. The use of different words for the various
conduits is not intended to be limiting in any way, but instead is
to assist the reader in fully understanding the interrelation
between the various conduits.
[0037] Therefore, the present invention is well adapted to attain
the ends and advantages mentioned as well as those that are
inherent therein. The particular embodiments disclosed above are
illustrative only, as the present invention may be modified and
practiced in different but equivalent manners apparent to those
skilled in the art, having the benefit of the teachings herein. For
example, as those of ordinary skill in the art will appreciate a
different number of platform cooling holes 230 may be implemented,
such platform cooling holes 230 may be formed at a different angle
than that disclosed herein, and such platform cooling holes 230 may
be oriented at a different location within the thickness of the
platform. Furthermore, no limitations are intended to the details
of construction or design herein shown, other than as described in
the claims below. It is therefore evident that the particular
illustrative embodiments disclosed above may be altered or modified
and all such variations are considered within the scope and spirit
of the present invention. Also, the terms in the claims have their
plain, ordinary meaning unless otherwise explicitly and clearly
defined by the patentee.
* * * * *