U.S. patent application number 13/205475 was filed with the patent office on 2013-02-14 for system and method for controlling flow in turbomachinery.
This patent application is currently assigned to General Electric Company. The applicant listed for this patent is Sushil Babu Mane, Vishal Rajpurohit, Karthik Srinivasan. Invention is credited to Sushil Babu Mane, Vishal Rajpurohit, Karthik Srinivasan.
Application Number | 20130039772 13/205475 |
Document ID | / |
Family ID | 46639371 |
Filed Date | 2013-02-14 |
United States Patent
Application |
20130039772 |
Kind Code |
A1 |
Mane; Sushil Babu ; et
al. |
February 14, 2013 |
SYSTEM AND METHOD FOR CONTROLLING FLOW IN TURBOMACHINERY
Abstract
A system includes a turbine. The turbine includes a first
turbine blade comprising a leading edge, a blade platform coupled
to the first turbine blade, and a protrusion disposed on the blade
platform adjacent the leading edge of the first turbine blade. The
protrusion is configured to increase a first static pressure of a
cooling flow near the leading edge above a second static pressure
of a hot gas flow near the leading edge.
Inventors: |
Mane; Sushil Babu;
(Bangalore, IN) ; Srinivasan; Karthik; (Bangalore,
IN) ; Rajpurohit; Vishal; (Bangalore, IN) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Mane; Sushil Babu
Srinivasan; Karthik
Rajpurohit; Vishal |
Bangalore
Bangalore
Bangalore |
|
IN
IN
IN |
|
|
Assignee: |
General Electric Company
Schenectady
US
|
Family ID: |
46639371 |
Appl. No.: |
13/205475 |
Filed: |
August 8, 2011 |
Current U.S.
Class: |
416/223R ;
29/889 |
Current CPC
Class: |
F01D 11/02 20130101;
F01D 11/001 20130101; Y02T 50/673 20130101; Y10T 29/49316 20150115;
F01D 5/081 20130101; Y02T 50/60 20130101; Y02T 50/676 20130101 |
Class at
Publication: |
416/223.R ;
29/889 |
International
Class: |
F01D 5/14 20060101
F01D005/14; B21D 53/78 20060101 B21D053/78 |
Claims
1. A system, comprising: a turbine, comprising: a first turbine
blade comprising a leading edge; a blade platform coupled to the
first turbine blade; and a protrusion disposed on the blade
platform adjacent the leading edge of the first turbine blade,
wherein the protrusion is configured to increase a first static
pressure of a cooling flow near the leading edge above a second
static pressure of a hot gas flow near the leading edge.
2. The system of claim 1, wherein the protrusion is configured to
block entry of the hot gas flow into a wheel space cavity near the
leading edge of the first turbine blade.
3. The system of claim 1, comprising a second turbine blade,
wherein the protrusion does not extend across a space between the
first turbine blade and the second turbine blade.
4. The system of claim 1, wherein the protrusion is configured to
increase the first static pressure of the cooling flow only near
the leading edge of the first turbine blade.
5. The system of claim 1, wherein the protrusion comprises at least
one of a circular cross-sectional shape, an oval cross-sectional
shape, a triangular cross-sectional shape, a square cross-sectional
shape, a rectangular cross-sectional shape, a polygonal
cross-sectional shape, an aerodynamic cross-sectional shape, an
arcuate cross-sectional shape, or a combination thereof.
6. The system of claim 1, wherein an aspect ratio of the protrusion
is less than approximately 2:1, the aspect ratio comprises a ratio
of a first dimension of the protrusion to a second dimension of the
protrusion.
7. The system of claim 1, wherein the protrusion is disposed
between a front edge of the blade platform and the leading edge of
the first turbine blade.
8. The system of claim 1, comprising a one-piece structure having
the protrusion integrally formed on the blade platform, or a
multi-piece structure having the protrusion coupled to the blade
platform.
9. A system, comprising: an ingestion restrictor configured to
mount on a blade platform adjacent a leading edge of a turbine
blade of a turbine, wherein the ingestion restrictor is configured
to block entry of a hot gas flow into a wheel space cavity near the
leading edge of the turbine blade.
10. The system of claim 9, wherein the ingestion restrictor is
configured to increase a first static pressure of a cooling flow
near the leading edge above a second static pressure of the hot gas
flow near the leading edge.
11. The system of claim 9, wherein a width of the ingestion
restrictor is less than a radial height of the blade platform.
12. The system of claim 11, wherein a ratio of the radial height of
the blade platform to the width of the ingestion restrictor is less
than approximately 2:1.
13. The system of claim 9, wherein the ingestion restrictor is
disposed between a front edge of the blade platform and the leading
edge of the turbine blade.
14. The system of claim 9, wherein the ingestion restrictor
comprises at least one of a circular cross-sectional shape, an oval
cross-sectional shape, a triangular cross-sectional shape, a square
cross-sectional shape, a rectangular cross-sectional shape, a
polygonal cross-sectional shape, an aerodynamic cross-sectional
shape, an arcuate cross-sectional shape, or a combination
thereof.
15. The system of claim 9, wherein the ingestion restrictor is
oriented at an angle from an axial axis of the turbine.
16. The system of claim 9, comprising the turbine having the
ingestion restrictor.
17. A method, comprising: mounting an ingestion restrictor on a
blade platform upstream of a leading edge of a first turbine blade
of a turbine, and increasing a first static pressure of a cooling
flow near the leading edge of the first turbine blade above a
second static pressure of a hot gas flow using the ingestion
restrictor.
18. The method of claim 17, comprising blocking entry of the hot
gas flow into a wheel space cavity near the leading edge of the
first turbine blade using the ingestion restrictor.
19. The method of claim 17, comprising disposing the ingestion
restrictor between a front edge of the blade platform and the
leading edge of the first turbine blade.
20. The method of claim 17, comprising locating the ingestion
restrictor to not extend across a space between the first turbine
blade and a second turbine blade.
Description
BACKGROUND OF THE INVENTION
[0001] The subject matter disclosed herein relates to
turbomachinery, and more specifically, to controlling flow within
turbines.
[0002] In general, gas turbine engines combust a mixture of
compressed air and fuel to produce hot combustion gases. The
combustion gases may flow through one or more turbine stages to
generate power for a load and/or compressor. Surfaces of the gas
turbine engine expected to come in contact with the hot combustion
gases may be made from materials able to withstand hot
temperatures. Such high-temperature materials may be difficult to
obtain and/or expensive. Other surfaces of the gas turbine engine
that are not expected to come in contact with the hot combustion
gases may be made from more readily available and/or less expensive
materials. Such low-temperature materials may not be designed to be
exposed to the hot combustion gases. Unfortunately, in certain
situations, the hot combustion gases may come in contact with the
surfaces not designed to withstand high temperatures, thereby
potentially degrading these surfaces and the gas turbine engine.
Similarly, other turbomachinery, such as steam turbines, for
example, may be susceptible to such degradation caused by high
temperature gases contacting surfaces designed for exposure to
lower temperatures.
BRIEF DESCRIPTION OF THE INVENTION
[0003] Certain embodiments commensurate in scope with the
originally claimed invention are summarized below. These
embodiments are not intended to limit the scope of the claimed
invention, but rather these embodiments are intended only to
provide a brief summary of possible forms of the invention. Indeed,
the invention may encompass a variety of forms that may be similar
to or different from the embodiments set forth below.
[0004] In a first embodiment, a system includes a turbine. The
turbine includes a first turbine blade comprising a leading edge, a
blade platform coupled to the first turbine blade, and a protrusion
disposed on the blade platform adjacent the leading edge of the
first turbine blade. The protrusion is configured to increase a
first static pressure of a cooling flow near the leading edge above
a second static pressure of a hot gas flow near the leading
edge.
[0005] In a second embodiment, a system includes an ingestion
restrictor configured to mount on a blade platform adjacent a
leading edge of a turbine blade of a turbine. The ingestion
restrictor is configured to block entry of a hot gas flow into a
wheel space cavity near the leading edge of the turbine blade.
[0006] In a third embodiment, a method includes mounting an
ingestion restrictor on a blade platform upstream of a leading edge
of a first turbine blade of a turbine and increasing a first static
pressure of a cooling flow near the leading edge of the first
turbine blade above a second static pressure of a hot gas flow
using the ingestion restrictor.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] These and other features, aspects, and advantages of the
present invention will become better understood when the following
detailed description is read with reference to the accompanying
drawings in which like characters represent like parts throughout
the drawings, wherein:
[0008] FIG. 1 is a schematic flow diagram of an embodiment of a gas
turbine engine that may employ ingestion restrictors;
[0009] FIG. 2 is a cross-sectional side view of an embodiment of
the gas turbine engine of FIG. 1 taken along the longitudinal
axis;
[0010] FIG. 3 is a cross-sectional side view of an embodiment of a
stage of a gas turbine engine that may employ injection
restrictors;
[0011] FIG. 4 is a perspective view of an embodiment of a stage of
a gas turbine engine that may employ ingestion restrictors;
[0012] FIG. 5 is a perspective view of an embodiment of an
ingestion restrictor;
[0013] FIG. 6 is a cross-sectional top view of an embodiment of an
ingestion restrictor and a gas turbine blade;
[0014] FIG. 7 is a cross-sectional top view of an embodiment of an
ingestion restrictor;
[0015] FIG. 8 is a cross-sectional top view of an embodiment of an
ingestion restrictor;
[0016] FIG. 9 is a cross-sectional top view of an embodiment of an
ingestion restrictor;
[0017] FIG. 10 is a cross-sectional top view of an embodiment of an
ingestion restrictor;
[0018] FIG. 11 is a cross-sectional top view of an embodiment of an
ingestion restrictor;
[0019] FIG. 12 is a cross-sectional top view of an embodiment of an
ingestion restrictor;
[0020] FIG. 13 is a cross-sectional top view of an embodiment of an
ingestion restrictor; and
[0021] FIG. 14 is a cross-sectional side view of an embodiment of
an ingestion restrictor.
DETAILED DESCRIPTION OF THE INVENTION
[0022] One or more specific embodiments of the present invention
will be described below. In an effort to provide a concise
description of these embodiments, all features of an actual
implementation may not be described in the specification. It should
be appreciated that in the development of any such actual
implementation, as in any engineering or design project, numerous
implementation-specific decisions must be made to achieve the
developers' specific goals, such as compliance with system-related
and business-related constraints, which may vary from one
implementation to another. Moreover, it should be appreciated that
such a development effort might be complex and time consuming, but
would nevertheless be a routine undertaking of design, fabrication,
and manufacture for those of ordinary skill having the benefit of
this disclosure.
[0023] When introducing elements of various embodiments of the
present invention, the articles "a," "an," "the," and "said" are
intended to mean that there are one or more of the elements. The
terms "comprising," "including," and "having" are intended to be
inclusive and mean that there may be additional elements other than
the listed elements.
[0024] The present disclosure is directed to turbomachinery, such
as multi-stage turbines, e.g., gas turbine engines or steam turbine
engines. The multi-stage turbine may include a first turbine blade
that includes a leading edge and a blade platform coupled to the
first turbine blade. A protrusion may be disposed on the blade
platform adjacent the leading edge of the first turbine blade. In
certain embodiments, the protrusion may increase a first static
pressure of a cooling flow near the leading edge above a second
static pressure of a hot gas flow near the leading edge. For
example, the cooling flow may flow radially outward from an axial
axis of the turbine to help cool components of the turbine near the
axial axis. In a gas turbine engine, the hot gas flow may be
generated by combustion of a fuel with oxygen and may flow axially
along a hot gas flow path through the turbine. In a steam turbine
engine, the hot gas flow may be a steam flow. By increasing the
first static pressure of the cooling flow, the protrusion may help
block the hot gas flow from reaching surfaces and/or areas of the
turbine not in the hot gas flow path, such as the surfaces cooled
by the cooling flow. Entry of the hot gas flow into such areas of
the turbine may be referred to as ingestion, and the protrusion may
be referred to as an injection restrictor.
[0025] Use of such protrusions, or ingestion restrictors, may offer
several advantages. For example, the first turbine blade and other
components of the turbine in the hot gas flow path may be made from
materials suitable for high temperatures. However, the components
of the turbine not in the hot gas flow path may be made from
different materials not suitable for high temperatures. Such
materials may be degraded upon exposure to the hot gas flow. Thus,
by helping to block the hot gas flow from entering areas of the
turbine outside the hot gas flow path, the protrusion may help
reduce the possibility of degradation of components of the turbine
disposed in those areas not designed for exposure to high
temperatures. In addition, the disclosed protrusions may offer
advantages compared to other methods of ingestion restriction. For
example, the protrusions may be located adjacent to only the
leading edge of the first turbine blade, where ingestion may be
more likely. Thus, the protrusions are not disposed upstream of the
spaces between adjacent turbine blades of the turbine. Thus, any
reduction in the aerodynamic performance of the turbine caused by
the protrusions may be reduced. In addition, such small protrusions
may be easy to manufacture, inexpensive, and easy to install in the
turbine. Embodiments of the protrusions may be used to avoid
ingestion in any turbomachinery, such as gas turbine engines and
steam turbine engines, for example. In the following discussion, a
gas turbine engine is used as one non-limiting example of
turbomachinery in which such embodiments of the protrusions may be
used.
[0026] Turning to the drawings, FIG. 1 is a block diagram of an
exemplary system 10 including a gas turbine engine 12 that may
employ protrusions, or ingestion restrictors, as described in
detail below. In certain embodiments, the system 10 may include an
aircraft, a watercraft, a locomotive, a power generation system, or
combinations thereof. The illustrated gas turbine engine 12
includes an air intake section 16, a compressor 18, a combustor
section 20, a turbine 22, and an exhaust section 24. The turbine 22
is coupled to the compressor 18 via a shaft 26.
[0027] As indicated by the arrows, air may enter the gas turbine
engine 12 through the intake section 16 and flow into the
compressor 18, which compresses the air prior to entry into the
combustor section 20. The illustrated combustor section 20 includes
a combustor housing 28 disposed concentrically or annularly about
the shaft 26 between the compressor 18 and the turbine 22. The
compressed air from the compressor 18 enters combustors 30, where
the compressed air may mix and combust with fuel within the
combustors 30 to drive the turbine 22.
[0028] From the combustor section 20, the hot combustion gases flow
through the turbine 22, driving the compressor 18 and a load via
the shaft 26. For example, the combustion gases may apply motive
forces to turbine rotor blades within the turbine 22 to rotate the
shaft 26, which may be used to drive a load, such as an electrical
generator. After flowing through the turbine 22, the hot combustion
gases may exit the gas turbine engine 12 through the exhaust
section 24. As discussed below, the turbine 22 may include a
plurality of protrusions (e.g., ingestion restrictors) disposed
adjacent leading edges of the turbine rotor blades.
[0029] FIG. 2 is a cross-sectional side view of an embodiment of
the gas turbine engine 12 of FIG. 1 taken along a longitudinal axis
32. As depicted in FIG. 2, the gas turbine 22 includes three
separate stages 34. Each stage 34 includes a set of blades 36
coupled to a rotor wheel 38 that may be rotatably attached to the
shaft 26 (FIG. 1). The blades 36 extend radially outward from the
rotor wheels 38 and are partially disposed within the path of the
hot combustion gases. Seals 40 extend between and are supported by
adjacent rotor wheels 38. As discussed below, protrusions may be
disposed adjacent the leading edges of the blades 36. Although the
gas turbine 22 is illustrated as a three-stage turbine, the
protrusions described herein may be employed in any suitable type
of turbine with any number of stages and shafts. For example, the
protrusions may be included in a single stage gas turbine, in a
dual turbine system that includes a low-pressure turbine and a
high-pressure turbine, or in a steam turbine. Further, the
protrusions described herein may also be employed in a rotary
compressor, such as the compressor 18 illustrated in FIG. 1. The
protrusions may be made from various high-temperature alloys, such
as, but not limited to, nickel based alloys.
[0030] As described above with respect to FIG. 1, air enters
through the air intake section 16 and is compressed by the
compressor 18. The compressed air from the compressor 18 is then
directed into the combustor section 20 where the compressed air is
mixed with fuel. The mixture of compressed air and fuel is
generally burned within the combustor section 20 to generate
high-temperature, high-pressure combustion gases, which are used to
generate torque within the turbine 22. Specifically, the combustion
gases apply motive forces to the blades 36 to turn the wheels 38.
In certain embodiments, a pressure drop may occur at each stage 34
of the turbine 22, which may allow gas leakage flow through
unintended paths. For example, the hot combustion gases may leak
into the interstage volume between turbine wheels 38, which may
place thermal stresses on the turbine components. In certain
embodiments, the interstage volume may be cooled by a cooling flow,
such as discharge air bled from the compressor or provided by
another source. However, flow of hot combustion gases into the
interstage volume may abate the cooling effects of the cooling
flow. Accordingly, the protrusions may be disposed adjacent leading
edges of the blades 36 to block the hot combustion gases from
leaking into the interstage volume. In addition, the protrusions
may increase a static pressure of the cooling flow near the leading
edges of the blades 36 above a static pressure of the hot
combustion gases near the leading edges. The hot combustion gases
may be blocked from entering the interstage volume because the
cooling flow has a higher static pressure than the hot combustion
gases.
[0031] FIG. 3 is a cross-sectional side view of an embodiment of
one of the rotor stages 34 shown in FIG. 2. In the following
discussion, reference may be made to an axial direction or axis 50,
a radial direction or axis 52, and a circumferential direction or
axis 54, relative to the longitudinal axis 32 of the gas turbine
engine 12. Hot fluids, such as hot combustion gases or steam, with
a flow path 56 (illustrated generally by arrows) enters at an
upstream side 58 and exits at a downstream side 60 relative to the
stage 34. For illustrative purposes, only a portion of the stage 34
is illustrated in FIG. 3. In addition, a portion of a stator stage
42 upstream of the rotor stage 34 is shown in FIG. 3. For example,
a portion of a stator blade 44 is shown adjacent to the rotor blade
36. The rotor blade 36 and the stator blade 44 may be made from
materials selected for high temperature service because of their
exposure to the hot fluids during operation of the gas turbine
engine 12. In addition, the rotor blade 36 may have a leading edge
46 facing the upstream side 58 and a trailing edge 48 facing the
downstream side 60.
[0032] As shown in FIG. 3, a cooling flow enters an interstage
volume 61, or wheel space cavity, between the stator stage 42 and
the rotor stage 34, flowing in the radial direction 52 along a
cooling flow path 62 (illustrated generally by arrows). A
temperature of the cooling flow may be less than a temperature of
the hot fluids. Thus, components of the interstage volume 61 (e.g.,
the rotor wheel 38 and the seal 40) may be made from materials that
are unsuitable for high temperature service. If such components are
exposed to the hot fluids, the components may experience thermal
stresses that may degrade the components. The coolant (e.g., air)
flows in the radial direction 52 past the rotor wheel 38 and toward
a first wing seal 64, which may help to seal the interstage volume
61 from the hot fluids. The cooling flow moves past the first wing
seal 64 and then moves in the axial direction 50 after encountering
a first projection 66 of the stator stage 42. The cooling flow then
turns in the radial direction 52 again toward a second wing seal
68, which also may help seal the interstage volume 61 from the hot
fluids. The first and second wing seals 64 and 68 may extend
circumferentially 54 about the axial axis 50 without any
circumferential gaps or spaces. The cooling flow moves past the
second wing seal 68 and then moves in the axial direction 50 after
encountering a second projection 70 of the stator stage 42. The
cooling flow 62 then turns in the radial direction 52 again toward
a blade platform 72, which couples the blade 36 to the seal 40. The
blade platform 72 may be defined by a radial height 73.
[0033] A protrusion, or an ingestion restrictor, 74 may be disposed
on the upstream side 58 of the blade platform 72. As shown in FIG.
3, a portion 76 of the hot combustion gases is blocked from
entering the interstage volume 61 by the protrusion 74. As
described in detail below, the protrusion 74 may help increase a
static pressure of the cooling flow above a static pressure of the
hot fluids adjacent to the blade 36, thereby helping to block the
hot fluids from entering the interstage volume 61. The protrusion
74 may be defined by a height 78 and a width 80. In certain
embodiments, the height 78 and/or the width 80 may be between
approximately 0.5 cm to 15 cm, 1.5 cm to 10 cm, or 2.5 cm to 5 cm.
In other embodiments, the height 78 may be less than the width 80
to help reduce any effect of the protrusion 74 on the aerodynamic
efficiency of the gas turbine engine 12. For example, an aspect
ratio (e.g., ratio of the width 80 to the height 78) may be between
approximately 1.1:1 to 5:1, 1.3:1 to 3:1, or 1.5:1 to 2:1. In
general, the aspect ratio of the protrusion 74 may refer to any
ratio of a first dimension of the protrusion 74 to a second
dimension of the protrusion 74. In addition, a ratio of the radial
height 73 of the blade platform 72 to the width 80 of the
protrusion 72 may be between approximately 1.1:1 to 4:1, 1.3:1 to
3:1, or 1.5:1 to 2:1. By blocking the portion 76 of the hot
combustion gases, the protrusion 74 helps to protect the components
of the interstage volume 61 from degradation caused by the hot
fluids.
[0034] FIG. 4 is a perspective view of an embodiment of one of the
rotor stages 34 shown in FIG. 2. Elements in FIG. 4 in common with
those shown in FIG. 3 are labeled with the same reference numerals.
As shown in FIG. 4, the protrusions 74 are disposed adjacent the
leading edges 46 of the turbine blades 36. Thus, the protrusions 74
are separated by a distance 90. In other words, the protrusions 74
are disposed only adjacent to the leading edges 46 of the turbine
blades 36 and are not disposed in, and/or do not extend into, gaps
or spaces between the turbine blades 36. Thus, each protrusion 74
may be aligned or centered in front of only one blade 36, such that
the protrusion 74 generally does not impede the gas flow between
adjacent blades 36. Accordingly, the protrusions 74 are different
from other components of the gas turbine engine 12 that extend
completely circumferentially 54 about the axial axis 50, such as,
for example, the first and second wing seals 64 and 68. Thus, the
protrusions 74 may have a minimal impact on the overall aerodynamic
efficiency of the gas turbine engine 12. Further, it may be
relatively simple and inexpensive to add the protrusions 74 to the
blade platform 72. In certain embodiments, the protrusions 74 may
be coupled to the blade platform 72 in a multi-piece structure
using a variety of techniques, such as, but not limited to,
welding, brazing, adhesives, fasteners, and so forth. In other
embodiments, the protrusions 74 may be integrally formed on the
blade platforms 74 as a one-piece structure. In certain
embodiments, a ratio of the width 80 of the protrusion 72 to the
distance 90 may be between approximately 0.1:1 to 4:1, 0.2:1 to
2:1, or 0.3:1 to 0.5:1.
[0035] FIG. 5 is a perspective view of the protrusions 74 and the
turbine blades 36. As shown in FIG. 5, the leading edges 46 of the
turbine blades 36 are separated by a distance 100. The distance 100
may correspond to approximately the separation distance 90 between
the protrusions 74. Thus, the protrusions 74 are disposed adjacent
to the leading edges 46 of the turbine blades 36. In other words,
the protrusions 74 do not extend into the distance 100 between the
blades 36 and are generally limited to the area in front of the
leading edges 46. In addition, the protrusions 74 are separated
from the leading edges 46 by a distance 102. In certain
embodiments, the distance 102 may be between approximately 0.5 cm
to 15 cm, 1.5 cm to 10 cm, or 2.5 cm to 5 cm. In addition, the
protrusions 74 are separated from a front edge 103 of the blade
platform 72 by a distance 104. In certain embodiments, the distance
104 may be between approximately 0.5 cm to 15 cm, 1.5 cm to 10 cm,
or 2.5 cm to 5 cm. Thus, the protrusions 74 may be generally
limited to the area between the front edge 103 of the blade
platform 72 and the leading edge 46 of the turbine blades 36.
[0036] In addition, a graph of static pressures associated with the
protrusions 74 is superimposed over the perspective view shown in
FIG. 5. Specifically, an x-axis 105 corresponds to the
circumferential direction 54 and a y-axis 106 corresponds to a
static pressure of a fluid. A first curve 108 shows the static
pressure associated with the cooling flow flowing along the cooling
flow path 62. A second curve 110 shows the static pressure
associated with the hot fluids flowing along the flow path 56. As
shown in FIG. 5, the cooling flow static pressure 108 is greater
than the hot fluid static pressure 110 in the entire
circumferential direction 54. In other words, the protrusions 74
help the cooling flow static pressure 108 not to fall below the hot
fluid static pressure 110. For example, the hot fluid static
pressure 110 increases at peaks 112 near the leading edges 46
because the blades 36 at least partially resist the flow of the hot
fluids. Similarly, the protrusions 74 adjacent to the leading edges
46 help to increase the cooling flow static pressure 108 at the
peaks 112 because the protrusions 74 at least partially resist the
flow of the cooling flow. When the cooling flow static pressure 108
exceeds the hot fluid static pressure 110, the hot fluids may be
blocked from entering the interstage volume 61. Without the
protrusions 74, the hot fluid static pressure 110 may exceed the
cooling flow static pressure 108 at the peaks 112, thereby enabling
the hot fluids to enter the interstage volume 61. At troughs 114
between the blades 36, the cooling flow static pressure 108 exceeds
the hot fluid static pressure 110 because the turbine blades 36 are
not present to resist the flow of the hot fluids. In other words,
the protrusions 74 are not used between the turbine blades 36
because the cooling flow static pressure 108 already exceeds the
hot fluid static pressure 110, thereby blocking the hot fluids from
entering the interstage volume 61. Thus, the protrusions 74
increase the cooling flow static pressure 108 only where needed to
help block ingestion of the hot fluids into the interstage volume
61, namely adjacent to the leading edges 46 of the blades 36.
[0037] FIG. 6 is cross-sectional top view of an embodiment of the
protrusion 74 and the blade 36. As shown in FIG. 6, the leading
edge 46 of the blade 36 faces the upstream side 58 and the trailing
edge 48 faces the downstream side 60. The leading edge 46 of the
blade 36 may have width 134. In certain embodiments, the width 134
may be between approximately 0.5 cm to 15 cm, 1.5 cm to 10 cm, or
2.5 cm to 5 cm. Thus, a ratio of the width 134 of the blade 36 to
the width 80 of the protrusion 74 may be between approximately
0.5:1 to 4:1, 0.75:1 to 3:1, or 1:1 to 2:1. The leading edge 46 of
the blade 36 may be generally bisected by an axis 136, which is
offset by an angle 138 from the axial axis 50. In certain
embodiments, the angle 138 may be between approximately 10 degrees
to 90 degrees, 25 degrees to 75 degrees, or 40 degrees to 60
degrees. As shown in FIG. 6, the blade 36 curves away from the axis
136 toward the trailing edge 48. The protrusion 74 may also have a
leading edge 140 and a trailing edge 142. In the illustrated
embodiment, the protrusion 74 has an aerodynamic shape, such as an
airfoil shape or a teardrop shape. Such a shape for the protrusion
74 may help reduce the resistance to gases flowing past the
protrusion 74. Thus, an aerodynamically-shaped protrusion 74 may
reduce any negative effect on the aerodynamic efficiency of the gas
turbine engine 12 caused by the protrusion 74. As shown in FIG. 6,
the protrusion 74 may be generally aligned with the axis 136. In
other words, the protrusion 74 is oriented at the angle 138 from
the axial axis 50. In addition, the protrusion 74 may have a length
144, which may be between approximately 1 cm to 30 cm, 3 cm to 20
cm, or 5 cm to 10 cm. Thus, the length 144 of the protrusion 74 may
be greater than the width 80. In other embodiments, the length 144
may be less than the width 80 or the length 144 and width 80 may be
approximately the same. The separation distance 102 between the
protrusion 74 and the blade 36 may be adjusted to determine the
effect of the protrusion 74 on the cooling flow static pressure
108. For example, as the separation distance 102 is increased, the
cooling flow static pressure 108 may decrease. Similarly, as the
separation distance 102 is decreased, the cooling flow static
pressure 108 may increase. Thus, the separation distance 102 may be
selected to help the cooling flow static pressure 108 to generally
exceed the hot fluid static pressure 110 under a wide variety of
operating conditions of the gas turbine engine 12. In certain
embodiments, the separation distance 102 may be between
approximately 1 cm to 30 cm, 3 cm to 20 cm, or 5 cm to 10 cm.
[0038] FIGS. 7 through 13 show cross-sectional top views of various
embodiments of the protrusions 74 with different cross-sectional
shapes. For example, in FIG. 7, the protrusion 74 has a circular
cross-sectional shape. Such a shape may be easier to manufacture
than other shapes. In addition, the orientation of the protrusion
74 with respect to the axis 136 may be unimportant. In FIG. 8, the
protrusion 74 has a generally oval shape, which may also be easy to
manufacture because of the curved shape of the protrusion 74. As
shown in FIG. 8, the leading and trailing edges 140 and 142 are
generally curved. In FIG. 9, the protrusion 74 has a bullet shape
or football shape with sharp edges for the leading and trailing
edges 140 and 142. Such a configuration of the protrusion 74 may
offer aerodynamic characteristics different from the protrusions
shown in FIGS. 7 and 8, for example. In FIG. 10, the protrusion 74
has a generally triangular shape. In other words, the leading edge
140 has a sharp edge and the trailing edge 142 is generally flat.
In FIG. 11, the protrusion 74 has a generally square or rectangular
shape. Again, the leading and trailing edges 140 and 142 have
generally sharp edges. In FIG. 12, the protrusion 74 has a
generally polygonal cross-sectional shape (e.g., hexagonal shape)
with sharp edges for the leading and trailing edges 140 and 142. In
FIG. 13, the protrusion 74 has a crescent or arcuate
cross-sectional shape. The leading and trailing edges 140 and 142
may have generally curved surfaces facing toward and away from the
upstream and downstream sides 58 and 60. FIGS. 7 through 13 are
meant to only provide several examples of cross-sectional shapes
for the protrusions 74. In other embodiments, other shapes may be
selected based on the requirements of a particular application. For
example, the protrusion 74 may have a cross-sectional shape that is
a combination of one or more of the shapes shown in FIGS. 7 through
13.
[0039] FIG. 14 is a cross-sectional side view of an embodiment of
the protrusion 74 taken along line 14-14 of FIG. 7. As shown in
FIG. 14, a base 160 of the protrusion 74 may be wider than a tip
162. In other words, a width 164 of the tip 162 may be less than
the width 80 of the base 160. Thus, the protrusion has a generally
triangular, or tapered, side cross-sectional shape, which may be
used with any of the previous embodiments of the protrusion 74
described in FIGS. 6 through 13. In other embodiments, the width 80
of the base 160 and the width 164 of the tip 162 may be
approximately the same. In further embodiments, the width 164 of
the tip 162 may be greater than the width 80 of the base 160. Thus,
the protrusion 74 may have a variety of side cross-sectional shapes
including at least one of, but not limited to, circles, ovals,
squares, rectangles, triangles, polygons, hourglass, trapezoids,
and so forth. The particular side cross-sectional shape selected
may depend on the requirements of a particular application. For
example, additional blocking of the hot fluids may be needed near
the base 160 of the protrusion 74. Thus, the width 80 of the base
160 may be configured to be greater than the width 164 of the tip
162, thereby reducing any effect on the aerodynamic efficiency of
the gas turbine engine 12.
[0040] This written description uses examples to disclose the
invention, including the best mode, and also to enable any person
skilled in the art to practice the invention, including making and
using any devices or systems and performing any incorporated
methods. The patentable scope of the invention is defined by the
claims, and may include other examples that occur to those skilled
in the art. Such other examples are intended to be within the scope
of the claims if they have structural elements that do not differ
from the literal language of the claims, or if they include
equivalent structural elements with insubstantial differences from
the literal language of the claims.
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