U.S. patent application number 14/143508 was filed with the patent office on 2015-07-02 for interior cooling circuits in turbine blades.
This patent application is currently assigned to General Electric Company. The applicant listed for this patent is General Electric Company. Invention is credited to Aaron Ezekiel Smith.
Application Number | 20150184538 14/143508 |
Document ID | / |
Family ID | 53372225 |
Filed Date | 2015-07-02 |
United States Patent
Application |
20150184538 |
Kind Code |
A1 |
Smith; Aaron Ezekiel |
July 2, 2015 |
INTERIOR COOLING CIRCUITS IN TURBINE BLADES
Abstract
An airfoil of a turbine rotor blade that includes a cooling
configuration having a plurality of elongated flow passages for
receiving and directing a coolant along a path through the airfoil.
The cooling configuration may include: a central flow passage
flanked to each side by near-wall flow passages that includes a
pressure side near-wall flow passage and a suction side near-wall
flow passage; a first port that fluidly connects the central flow
passage to the pressure side near-wall flow passage; a second port
that fluidly connects the central flow passage to the suction side
near-wall flow passage; and impingement connectors that fluidly
connect the central flow passage to a leading edge flow
passage.
Inventors: |
Smith; Aaron Ezekiel;
(Simpsonville, SC) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Assignee: |
General Electric Company
Schenectady
NY
|
Family ID: |
53372225 |
Appl. No.: |
14/143508 |
Filed: |
December 30, 2013 |
Current U.S.
Class: |
416/97R |
Current CPC
Class: |
Y02T 50/676 20130101;
F05D 2260/201 20130101; F01D 5/186 20130101; Y02T 50/60 20130101;
F01D 25/12 20130101; F01D 5/187 20130101 |
International
Class: |
F01D 25/12 20060101
F01D025/12 |
Claims
1. A turbine blade, comprising: an airfoil having a leading edge, a
trailing edge, an outboard tip, and an inboard end where the
airfoil attaches to a root configured to couple the turbine blade
to a disc, wherein the airfoil further includes a cooling
configuration comprising a plurality of elongated flow passages for
receiving and directing a coolant along a path through the airfoil,
the cooling configuration comprising: a central flow passage
flanked to each side by near-wall flow passages that includes a
pressure side near-wall flow passage and a suction side near-wall
flow passage; a first port that fluidly connects the central flow
passage to an upstream portion of the pressure side near-wall flow
passage; a second port that fluidly connects the central flow
passage to an upstream portion of the suction side near-wall flow
passage; a leading edge flow passage; and impingement connectors
that fluidly connect the central flow passage to the leading edge
flow passage.
2. The turbine blade according to claim 1, wherein the leading edge
flow passage is positioned in close proximity to the leading edge
of the airfoil, the leading edge flow passage extending radially
outward in spaced relation to the leading edge of the airfoil from
a first end positioned near the inboard end of the airfoil to a
second end positioned near the outboard tip of the airfoil.
3. The turbine blade according to claim 2, wherein the impingement
connectors to allow coolant to pass from the central flow passage
to the leading edge flow passage and impinge on an inner surface of
the wall forming the leading edge.
4. The turbine blade according to claim 3, wherein the leading edge
flow passage comprises surface outlets through which the coolant is
exhausted from the turbine blade.
5. The turbine blade according to claim 3, wherein the plurality of
impingement connectors are radially spaced between the first and
second end of the leading edge flow passage.
6. The turbine blade according to claim 2, wherein the pressure
side near-wall flow passage includes axially-stacked and parallel
first and second flow passages, each of which have an inner wall
defined by the pressure side outer wall of the airfoil.
7. The turbine blade according to claim 6, wherein the suction side
near-wall flow passage includes axially-stacked and parallel first
and second flow passages, each of which have an inner wall defined
by the suction side outer wall of the airfoil.
8. The turbine blade according to claim 1, wherein the pressure
side near-wall flow passage comprises a switchback circuit that
includes: a first segment that extends radially outward from a
first end positioned near the inboard end of the airfoil to a
second end positioned near the outboard end of the airfoil; a
second segment that extends radially inward from a first end
positioned near the outboard end of the airfoil to a second end
positioned near the inboard end of the airfoil; and a crossover
passage that, near the outboard end of the airfoil, fluidly
connects the second end of the first segment to the first end of
the second segment.
9. The turbine blade according to claim 8, wherein the suction side
near-wall flow passage comprises a switchback circuit that
includes: a first segment that extends radially outward from a
first end positioned near the inboard end of the airfoil to a
second end positioned near the outboard end of the airfoil; a
second segment that extends radially inward from a first end
positioned near the outboard end of the airfoil to a second end
positioned near the inboard end of the airfoil; and a crossover
passage that, near the outboard end of the airfoil, fluidly
connects the second end of the first segment to the first end of
the second segment.
10. The turbine blade according to claim 9, wherein the first and
second segments of the pressure side near-wall flow passage share a
common, partitioning wall that is configured so to maintain a fixed
spaced relation therebetween; and wherein the first and second
segments of the suction side near-wall flow passage share a common,
partitioning wall that is configured so to maintain a fixed spaced
relation therebetween; further comprising a sinusoidal tip flow
passage that connects to at least one of the crossover
passages.
11. The turbine blade according to claim 9, wherein the first and
second segments of the pressure side near-wall flow passage are
partitioned by a traverse rib that connects the pressure side outer
wall to a camber line rib; and wherein the first and second
segments of the suction side near-wall flow passage are partitioned
by a traverse rib that connects the suction side outer wall to a
camber line rib.
12. The turbine blade according to claim 9, wherein the first port
is disposed near the second end of the second segment of the
pressure side near-wall flow passage, and the second port is
disposed near the second end of the second segment of the pressure
side near-wall flow passage.
13. The turbine blade according to claim 9, wherein the first end
of the first segment of the pressure side near-wall flow passage
comprises a connection to a coolant feed passage formed through the
root of the turbine blade; and wherein the first end of the first
segment of the suction side near-wall flow passage comprises a
connection to a coolant feed passage formed through the root of the
turbine blade.
14. The turbine blade according to claim 1, wherein the turbine
blade comprises a turbine rotor blade, and wherein the cooling
configuration comprises a position near the leading edge of the
airfoil.
15. The turbine blade according to claim 1, wherein the central
flow passage, the pressure side near-wall flow passage, the suction
side near-wall flow passage, and the leading edge flow passage are
disposed between the leading edge of the airfoil and a midpoint of
a camber line of the airfoil.
16. A turbine blade comprising an airfoil defined by a concave
shaped pressure side outer wall and a convex shaped suction side
outer wall that connect along leading and trailing edges and,
therebetween, form a radially extending chamber for receiving the
flow of a coolant, wherein the chamber includes a cooling
configuration having: three laterally stacked flow passages
positioned between the pressure side outer wall and the suction
side outer wall: a pressure side near-wall flow passage disposed
adjacent to the pressure side outer wall; a suction side near-wall
flow passage disposed adjacent to the suction side outer wall; and
a central plenum disposed between the pressure side near-wall flow
passage and the suction side near-wall flow passage; and a leading
edge flow passage that is positioned in close proximity and
parallel to the leading edge of the airfoil; wherein ports fluidly
connect the central flow passage to a downstream portion of the
pressure side near-wall flow passage and a downstream portion of
the suction side near-wall flow passage; and wherein impingement
connectors fluidly connect the central flow passage to the leading
edge flow passage.
17. The turbine blade according to claim 16, wherein the pressure
side near-wall flow passage comprises an axially-stacked two-pass
serpentine circuit, wherein each pass includes an inner wall that
makes contact with the pressure side outer wall; and wherein the
suction side near-wall flow passage comprises an axially-stacked
two-pass serpentine circuit, wherein each pass includes an inner
wall that makes contact with the suction side outer wall.
18. The turbine blade according to claim 17, wherein the two-pass
serpentine circuit of the pressure side near-wall flow passage
comprises a 180 degree turn near an outboard tip of the airfoil;
and wherein the two-pass serpentine circuit of the suction side
near-wall flow passage comprises a 180 degree turn near the
outboard tip of the airfoil.
19. The turbine blade according to claim 18, wherein each of an
upstream end and the downstream portion of the two-pass serpentine
circuit of the pressure side near-wall flow passage comprises a
position near an inboard end of the airfoil; and wherein each of an
upstream end and the downstream portion of the two-pass serpentine
circuit of the suction side near-wall flow passage comprises a
position near the inboard end of the airfoil.
20. The turbine blade according to claim 16, further comprising a
plurality of exhaust orifices in the leading edge flow passage for
exhausting the coolant onto an exterior surface of the airfoil;
wherein the impingement connectors are adapted to allow coolant to
pass from the central flow passage to the leading edge flow passage
and impinge on an inner surface of the wall forming the leading
edge.
Description
BACKGROUND OF THE INVENTION
[0001] This invention relates to turbine airfoils, and more
particularly to hollow turbine airfoils, such as rotor or stator
blades, having internal channels for passing fluids such as air to
cool the airfoils.
[0002] Combustion or gas turbine engines (hereinafter "gas
turbines") include a compressor, a combustor, and a turbine. As is
well known in the art, air compressed in the compressor is mixed
with fuel and ignited in the combustor and then expanded through
the turbine to produce power. The components within the turbine,
particularly the circumferentially arrayed rotor and stator blades,
are subjected to a hostile environment characterized by the
extremely high temperatures and pressures of the combustion
products that are expended therethrough. In order to withstand the
repetitive thermal cycling as well as the extreme temperatures and
mechanical stresses of this environment, the airfoils must have a
robust structure and be actively cooled.
[0003] As will be appreciated, turbine rotor and stator blades
often contain internal passageways or circuits that form a cooling
system through which a coolant, typically air bled from the
compressor, is circulated. Such cooling circuits are typically
formed by internal ribs that provide the required structural
support for the airfoil, and include multiple flow paths designed
to maintain the airfoil within an acceptable temperature profile.
The air passing through these cooling circuits often is vented
through film cooling apertures formed on the leading edge, trailing
edge, suction side, and pressure side of the airfoil.
[0004] It will be appreciated that the efficiency of gas turbines
increases as firing temperatures rise. Because of this, there is a
constant demand for technological advances that enable turbine
blades to withstand ever higher temperatures. These advances
sometimes include new materials that are capable of withstanding
the higher temperatures, but just as often they involve improving
the internal configuration of the airfoil so to enhance the blades
structure and cooling capabilities. However, because the use of
coolant decreases the efficiency of the engine, new arrangements
that rely too heavily on increased levels of coolant usage merely
trade one inefficiency for another. As a result, there continues to
be demand for new airfoil designs that offer internal airfoil
configurations and coolant circulation that improves coolant
efficiency.
[0005] A consideration that further complicates design of
internally cooled airfoils is the temperature differential that
develops during operation between the airfoils internal and
external structure. That is, because they are exposed to the hot
gas path, the external walls of the airfoil typically reside at
much higher temperatures during operation than many of the internal
ribs, which, for example, may have coolant flowing through
passageways defined to each side of them. In fact, a common airfoil
configuration includes a "four-wall" arrangement in which lengthy
inner ribs run parallel to the pressure and suction side outer
walls. It is known that high cooling efficiency can be achieved by
the near-wall flow passages that are formed in the four-wall
arrangement, however, the outer walls experience a significantly
greater level of thermal expansion than the inner walls. This
imbalanced growth causes stress to develop at the points at which
the inner ribs and outer walls connect, which may cause low cyclic
fatigue that can shorten the life of the blade. As such, the
development of airfoil structures that use coolant more efficiently
while also reducing stress caused by imbalanced thermal expansion
between internal and external regions remains a significant
technological industry objection.
BRIEF DESCRIPTION OF THE INVENTION
[0006] The present application thus describes an airfoil having a
leading edge, a trailing edge, an outboard tip, and an inboard end
where the airfoil attaches to a root configured to couple the
turbine blade to a disc. The airfoil may further include a cooling
configuration comprising a plurality of elongated flow passages for
receiving and directing a coolant along a path through the airfoil.
The cooling configuration may include: a central flow passage
flanked to each side by near-wall flow passages that includes a
pressure side near-wall flow passage and a suction side near-wall
flow passage; a first port that fluidly connects the central flow
passage to an upstream portion of the pressure side near-wall flow
passage; a second port that fluidly connects the central flow
passage to an upstream portion of the suction side near-wall flow
passage; a leading edge flow passage; and impingement connectors
that fluidly connect the central flow passage to the leading edge
flow passage.
[0007] These and other features of the present application will
become apparent upon review of the following detailed description
of the preferred embodiments when taken in conjunction with the
drawings and the appended claims.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] These and other features of this invention will be more
completely understood and appreciated by careful study of the
following more detailed description of exemplary embodiments of the
invention taken in conjunction with the accompanying drawings, in
which:
[0009] FIG. 1 is a schematic representation of an exemplary turbine
engine in which certain embodiments of the present application may
be used;
[0010] FIG. 2 is a sectional view of the compressor section of the
combustion turbine engine of FIG. 1;
[0011] FIG. 3 is a sectional view of the turbine section of the
combustion turbine engine of FIG. 1;
[0012] FIG. 4 is a perspective view of a turbine rotor blade of the
type in which embodiments of the present invention may be
employed;
[0013] FIG. 5 is a side sectional view of a turbine rotor blade
having an inner wall configuration according to conventional
design; and
[0014] FIG. 6 is a cross-sectional view of the turbine rotor blade
of FIG. 5.
DETAILED DESCRIPTION OF THE INVENTION
[0015] As an initial matter, in order to clearly describe the
current invention it will become necessary to select certain
terminology when referring to and describing relevant machine
components within a gas turbine. When doing this, if possible,
common industry terminology will be used and employed in a manner
consistent with its accepted meaning. Unless otherwise stated, such
terminology should be given a broad interpretation consistent with
the context of the present application and the scope of the
appended claims. Those of ordinary skill in the art will appreciate
that often a particular component may be referred to using several
different or overlapping terms. What may be described herein as
being a single part may include and be referenced in another
context as consisting of multiple components. Alternatively, what
may be described herein as including multiple components may be
referred to elsewhere as a single part. Accordingly, in
understanding the scope of the present invention, attention should
not only be paid to the terminology and description provided
herein, but also to the structure, configuration, function, and/or
usage of the component.
[0016] In addition, several descriptive terms may be used regularly
herein, and it should prove helpful to define these terms at the
onset of this section. These terms and their definitions, unless
stated otherwise, are as follows. As used herein, "downstream" and
"upstream" are terms that indicate a direction relative to the flow
of a fluid, such as the working fluid through the turbine engine
or, for example, the flow of air through the combustor or coolant
through one of the turbine's component systems. The term
"downstream" corresponds to the direction of flow of the fluid, and
the term "upstream" refers to the direction opposite to the flow.
The terms "forward" and "aft", without any further specificity,
refer to directions, with "forward" referring to the front or
compressor end of the engine, and "aft" referring to the rearward
or turbine end of the engine. It is often required to describe
parts that are at differing radial positions with regard to a
center axis. The term "radial" refers to movement or position
perpendicular to an axis. In cases such as this, if a first
component resides closer to the axis than a second component, it
will be stated herein that the first component is "radially inward"
or "inboard" of the second component. If, on the other hand, the
first component resides further from the axis than the second
component, it may be stated herein that the first component is
"radially outward" or "outboard" of the second component. The term
"axial" refers to movement or position parallel to an axis.
Finally, the term "circumferential" refers to movement or position
around an axis. It will be appreciated that such terms may be
applied in relation to the center axis of the turbine.
[0017] By way of background, referring now to the figures, FIGS. 1
through 4 illustrate an exemplary combustion turbine engine in
which embodiments of the present application may be used. It will
be understood by those skilled in the art that the present
invention is not limited to this type of usage. As stated, the
present invention may be used in combustion turbine engines, such
as the engines used in power generation and airplanes, steam
turbine engines, and other types of rotary engines. The examples
provided are not meant to be limiting to the type of the turbine
engine.
[0018] FIG. 1 is a schematic representation of a combustion turbine
engine 10. In general, combustion turbine engines operate by
extracting energy from a pressurized flow of hot gas produced by
the combustion of a fuel in a stream of compressed air. As
illustrated in FIG. 1, combustion turbine engine 10 may be
configured with an axial compressor 11 that is mechanically coupled
by a common shaft or rotor to a downstream turbine section or
turbine 13, and a combustor 12 positioned between the compressor 11
and the turbine 13.
[0019] FIG. 2 illustrates a view of an exemplary multi-staged axial
compressor 11 that may be used in the combustion turbine engine of
FIG. 1. As shown, the compressor 11 may include a plurality of
stages. Each stage may include a row of compressor rotor blades 14
followed by a row of compressor stator blades 15. Thus, a first
stage may include a row of compressor rotor blades 14, which rotate
about a central shaft, followed by a row of compressor stator
blades 15, which remain stationary during operation.
[0020] FIG. 3 illustrates a partial view of an exemplary turbine
section or turbine 13 that may be used in the combustion turbine
engine of FIG. 1. The turbine 13 may include a plurality of stages.
Three exemplary stages are illustrated, but more or less stages may
be present in the turbine 13. A first stage includes a plurality of
turbine buckets or turbine rotor blades 16, which rotate about the
shaft during operation, and a plurality of nozzles or turbine
stator blades 17, which remain stationary during operation. The
turbine stator blades 17 generally are circumferentially spaced one
from the other and fixed about the axis of rotation. The turbine
rotor blades 16 may be mounted on a turbine wheel (not shown) for
rotation about the shaft (not shown). A second stage of the turbine
13 also is illustrated. The second stage similarly includes a
plurality of circumferentially spaced turbine stator blades 17
followed by a plurality of circumferentially spaced turbine rotor
blades 16, which are also mounted on a turbine wheel for rotation.
A third stage also is illustrated, and similarly includes a
plurality of turbine stator blades 17 and rotor blades 16. It will
be appreciated that the turbine stator blades 17 and turbine rotor
blades 16 lie in the hot gas path of the turbine 13. The direction
of flow of the hot gases through the hot gas path is indicated by
the arrow. As one of ordinary skill in the art will appreciate, the
turbine 13 may have more, or in some cases less, stages than those
that are illustrated in FIG. 3. Each additional stage may include a
row of turbine stator blades 17 followed by a row of turbine rotor
blades 16.
[0021] In one example of operation, the rotation of compressor
rotor blades 14 within the axial compressor 11 may compress a flow
of air. In the combustor 12, energy may be released when the
compressed air is mixed with a fuel and ignited. The resulting flow
of hot gases from the combustor 12, which may be referred to as the
working fluid, is then directed over the turbine rotor blades 16,
the flow of working fluid inducing the rotation of the turbine
rotor blades 16 about the shaft. Thereby, the energy of the flow of
working fluid is transformed into the mechanical energy of the
rotating blades and, because of the connection between the rotor
blades and the shaft, the rotating shaft. The mechanical energy of
the shaft may then be used to drive the rotation of the compressor
rotor blades 14, such that the necessary supply of compressed air
is produced, and also, for example, a generator to produce
electricity.
[0022] FIG. 4 is a perspective view of a turbine rotor blade 16 of
the type in which embodiments of the present invention may be
employed. The turbine rotor blade 16 includes a root 21 by which
the rotor blade 16 attaches to a rotor disc. The root 21 may
include a dovetail configured for mounting in a corresponding
dovetail slot formed in the perimeter of the rotor disc. The root
21 may further include a shank that extends between the dovetail
and a platform 24, which is disposed at the junction of the airfoil
25 and the root 21 and defines a portion of the inboard boundary of
the flow path through the turbine 13. It will be appreciated that
the airfoil 25 is the active component of the rotor blade 16 that
intercepts the flow of working fluid and induces the rotor disc to
rotate. While the blade of this example is a turbine rotor blade
16, it will be appreciated that the present invention also may be
applied to other types of blades within the turbine engine 10,
including turbine stator blades 17. It will be seen that the
airfoil 25 of the rotor blade 16 includes a concave pressure side
outer wall 26 and a circumferentially or laterally opposite convex
suction side outer wall 27 extending axially between opposite
leading and trailing edges 28, 29 respectively. The sidewalls 26
and 27 also extend in the radial direction from the platform 24 to
an outboard tip 31.
[0023] FIGS. 5 and 6 provide exemplary embodiments of internal wall
structures that define a cooling configuration according to the
present invention. As indicated, the cooling configuration may
include a plurality of elongated flow passages for receiving and
directing a coolant through the airfoil 25. The cooling
configuration may be positioned near the leading edge 28 of the
airfoil 25. In a preferred embodiment, the several flow passages
that are included in the cooling configuration of the present
invention are positioned in the forward half of the airfoil 25.
[0024] In general, as illustrated in FIGS. 5 and 6, the cooling
configuration of the present invention includes a central flow
passage 40 that is flanked to each side by near-wall flow passages
43, 44. The near-wall flow passages include a pressure side
near-wall flow passage 43 and a suction side near-wall flow passage
44. Positioned forward of the central flow passage 40, a leading
edge flow passage 42 may be positioned in close proximity and
parallel to the leading edge 28 of the airfoil 25. A port 46 may
fluidly connect the central flow passage 40 to a downstream portion
of the pressure side near-wall flow passage 43. Another port 46 may
fluidly connects the central flow passage 40 to a downstream
portion of the suction side near-wall flow passage 44. Finally,
impingement connectors 48 may fluidly connect the central flow
passage 40 to the leading edge flow passage 42.
[0025] It will be appreciated that, between the pressure side outer
wall 26 and the suction side outer wall 27, the cooling
configuration of the present invention provides for the laterally
stacking of three flow passages: the pressure side near-wall flow
passage 43, which is disposed adjacent to the pressure side outer
wall 26; the suction side near-wall flow passage 44, which is
disposed adjacent to the suction side outer wall 27; and the
central flow passage 40, which is disposed between the pressure
side and the suction side near-wall flow passages 43, 44. Given the
ports 46 that connect the downstream portions of the pressure side
near-wall flow passage 43 and a downstream portion of the suction
side near-wall flow passage 44, it will be appreciated that the
flow through the central flow passage 40 represents the combined
flow of coolant from the two near-wall flow passages 43, 44.
[0026] A number of impingement connectors 48 fluidly connect the
central flow passage 40 to the leading edge flow passage 42. As
illustrated, the leading edge flow passage 42 is positioned in
close proximity to the leading edge 28 of the airfoil 25. In
preferred embodiments, the leading edge flow passage 42 extends
radially outward in spaced relation to the leading edge 28 of the
airfoil 25. At one end, the leading edge flow passage 42 is
positioned near the inboard end of the airfoil 25. At the opposite
end, the leading edge flow passage 42 is positioned near the
outboard tip of the airfoil 25. The impingement connectors 48 are
configured to allow coolant to pass from the central flow passage
40 to the leading edge flow passage 42, while also impinging the
flow of coolant against an inner surface of the wall forming the
leading edge 28 of the airfoil 25. It will be appreciated that the
leading edge 28 of the airfoil 25 is a region that requires a
significant level of coolant, and that impinging the flow of
coolant in this manner enhances its effectiveness. In preferred
embodiments, the central flow passage 40 may extend radially
alongside the leading edge flow passage 42. The many impingement
connectors 48 may be radially spaced between inboard and outboard
ends of the leading edge flow passage 42 so that the flow of
coolant is evenly applied.
[0027] The leading edge flow passage 42 may include a number of
surface outlets 52. These may be configured to provide an outlet
through which exhausted coolant is expelled from the airfoil 25. It
will be appreciated that the surface outlets 52 also provide
outlets through which film cooling may be applied to targeted
surface areas of the airfoil 25.
[0028] In a preferred embodiment, the pressure side near-wall flow
passage 43 includes axially-stacked and parallel first and second
flow passages. As illustrated, each of these flow passages 43 may
be defined on one side by the pressure side outer wall 26 of the
airfoil 25. The pressure side near-wall flow passage 43 may include
a switchback circuit that includes: a first segment that extends
radially outward from a first end positioned near the inboard end
of the airfoil 25 to a second end positioned near the outboard end
of the airfoil 25; a second segment that extends radially inward
from a first end positioned near the outboard end of the airfoil 25
to a second end positioned near the inboard end of the airfoil 25;
and a crossover passage 47 that, near the outboard end of the
airfoil 25, fluidly connects the second end of the first segment to
the first end of the second segment. Given this configuration, the
first and second segments of the pressure side near-wall flow
passage 43 share a common, partitioning wall that may be configured
so to maintain a fixed spaced relation between the two
passages.
[0029] The suction side near-wall flow passage 44 may be similarly
formed. That is, the suction side near-wall flow passage 44 may
include axially-stacked and parallel first and second flow
passages. As indicated, each of these flow passages 44 are defined
on one side by the suction side outer wall 27 of the airfoil 25.
The suction side near-wall flow passage 44 may include a switchback
circuit that includes: a first segment that extends radially
outward from a first end positioned near the inboard end of the
airfoil 25 to a second end positioned near the outboard end of the
airfoil 25; a second segment that extends radially inward from a
first end positioned near the outboard end of the airfoil 25 to a
second end positioned near the inboard end of the airfoil 25; and a
crossover passage 47 that, near the outboard end of the airfoil 25,
fluidly connects the second end of the first segment to the first
end of the second segment. In this arrangement, the first and
second segments of the suction side near-wall flow passage 44 share
a common, partitioning wall that may be configured so to maintain a
fixed spaced relation between the two passages.
[0030] It will be appreciated that the cooling passages of this
type of configuration typically are formed with interconnecting
rib-like structural members (hereinafter "ribs"). Such ribs 60 may
be divided into two groups depending on their orientation and
length. A first type, a camber line rib 62, is typically a lengthy
rib that extends in parallel or approximately parallel to the
camber line of the airfoil 25. (The camber line of the airfoil 25
is a reference line stretching from the leading edge 28 to the
trailing edge 29 that connects the midpoints between the pressure
side outer wall 26 and the suction side outer wall 27.) The second
type of rib is referred to herein as a traverse rib 66. Traverse
ribs 66 are the shorter ribs that are shown connecting the outer
walls 26, 27 and the camber line ribs 62. The partitioning wall
between the first and second segments of the pressure side
near-wall flow passage 43 may be a traverse rib 66 that connects
the pressure side outer wall 26 to a camber line rib 62. Similarly,
the partitioning wall between the first and second segments of the
suction side near-wall flow passage 44 may be a traverse rib 66
that connects the suction side outer wall 27 to a camber line rib
62. As mentioned, the central flow passage 40, the pressure side
near-wall flow passage 43, the suction side near-wall flow passage
44, and the leading edge flow passage 42 have a position within the
forward half of the airfoil 25. The forward half of the airfoil may
be defined with reference to airfoil camber line. That is,
according to embodiments of the present invention, the forward half
of the airfoil may be defined as the region between the leading
edge 28 and the midpoint of the airfoil camber line.
[0031] The upstream end of the pressure side near-wall flow passage
43 (i.e., the inboard end of the first segment) may be connected to
a coolant supply feed 45 formed through the root 21 of the turbine
blade 16. The upstream end of the suction side near-wall flow
passage 44 (i.e., the inboard end of the first segment) also may be
connected to the coolant supply feed 45. As stated, the first
segment of both the pressure side near-wall flow passage 43 and the
suction side near-wall flow passage 44 may extend from the
connection made with the supply feed 45 in an outward radial
direction and connect, respectively, to one of the crossover
passage 47 formed near the outer tip 31 of the airfoil 25. It will
be appreciated that the crossover passages 47 fluidly connect the
two segments of each of the two near-wall flow passages 43, 44.
Then, from the crossover passage 47, the second segment of both the
pressure side near-wall flow passage 43 and the suction side
near-wall flow passage 44 may extend in an inward radial direction
toward the ports 46 that are positioned at the inboard end of the
airfoil 25. It will be appreciated that, formed in this manner, the
pressure side near-wall flow passage 43 and the suction side
near-wall flow passage 44 provide an axially-stacked two-pass
serpentine circuit, with the two-pass serpentine circuit of each
making a 180 degree turn near the outboard tip 31 of the airfoil
25.
[0032] In an alternative embodiment, a tip flow passage (not shown)
may be included near the tip 31 of the airfoil 25. The tip flow
passage may connect to and be supplied by the crossover passages 47
of either of the pressure side near-wall flow passage 43 and the
suction side near-wall flow passage 44. In a preferred embodiment,
the tip flow passage extends parallel to the tip 31 of the airfoil
25 toward the aft portions of the blade, and may have a winding or
sinusoidal configuration.
[0033] In operation, according to a preferred embodiment of the
present invention, the cooling circuit described herein may operate
to introduce a fresh supply of coolant via a supply feed through
the root to the upstream ends of leading edge near-wall flow
passages 43, 44, which may be positioned on the pressure and
suction sides 26, 27 of the airfoil 25. Each of these near-wall
flow passages 43, 44 may have a serpentine form that first directs
the fresh coolant through an outboard segment that extends the
length of the airfoil 25. The coolant may then make an approximate
180.degree. turn and then be directed by an inboard segment that
carries the coolant back to the inboard end of the airfoil 25. At
this downstream end of the near-wall flow passages, ports 46 may be
positioned that combine the two flows in a central flow passage 40.
The central flow passage 40 acts as a plenum by which a number of
radially spaced impingement connectors 48 are supplied coolant. The
impingement connectors 48 may fluidly deliver coolant from the
central flow passage 40 to the leading edge flow passage 42. The
impingement connectors 40 are configured to deliver an impinged
flow of coolant against the walls that form the leading edge of the
airfoil 25. From the leading edge flow passage 42, the coolant may
be expelled through surface outlets 52, which may arranged to
provide film cooling to targeted surface areas of the airfoil
25.
[0034] It will be appreciated that, configured in this manner, the
cooling configuration of the present invention introduces a
relatively fresh supply of coolant to serpentine near-wall flow
passages 43, 44 formed near the leading edge 28 of the airfoil 25.
By circulating a coolant first along the outer walls 26, 27 and
then combining the flows in the central flow passage 40 at a
downstream location and after each has absorbed heat, the present
invention reduces the temperature differential that typically
occurs between the internal structure and the external walls of the
airfoil 25. This will advantageously reduce the stresses that
typically arise due to the unbalanced thermal expansion that high
temperature differentials cause. Additionally, the present
invention allows for coolant "pre-heating" such that the total
cooling flow requirement is less than direct feeding the leading
edge with fresh coolant. These advantages are achieved even though
the cooling circuit of the present invention remains relatively
simple. It will be appreciated that the simplified circuit of the
present invention minimizes the pressure losses and backflow issues
that come with circuits having more serpentine or torturous paths.
Further, it will be appreciated that the present invention provides
a configuration that is conveniently tuned. Specifically, given the
separated pressure side and suction side near wall flow passages
43, 44, the configuration allows convenient adjustment of the level
of coolant directed toward each side of the airfoil 25. This is
advantageous given that the forward portions of an airfoil 25 have
very different heat load requirements from pressure side outer wall
26 to suction side outer wall 27. As such, significant flow savings
may be achieved by locally tuning convective heat transfer on each
side.
[0035] As one of ordinary skill in the art will appreciate, the
many varying features and configurations described above in
relation to the several exemplary embodiments may be further
selectively applied to form the other possible embodiments of the
present invention. For the sake of brevity and taking into account
the abilities of one of ordinary skill in the art, all of the
possible iterations is not provided or discussed in detail, though
all combinations and possible embodiments embraced by the several
claims below or otherwise are intended to be part of the instant
application. In addition, from the above description of several
exemplary embodiments of the invention, those skilled in the art
will perceive improvements, changes and modifications. Such
improvements, changes and modifications within the skill of the art
are also intended to be covered by the appended claims. Further, it
should be apparent that the foregoing relates only to the described
embodiments of the present application and that numerous changes
and modifications may be made herein without departing from the
spirit and scope of the application as defined by the following
claims and the equivalents thereof.
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