U.S. patent application number 11/543523 was filed with the patent office on 2008-04-10 for turbine airfoil cooling system with enhanced tip corner cooling channel.
This patent application is currently assigned to Siemens Power Generation, Inc.. Invention is credited to George Liang.
Application Number | 20080085193 11/543523 |
Document ID | / |
Family ID | 39301669 |
Filed Date | 2008-04-10 |
United States Patent
Application |
20080085193 |
Kind Code |
A1 |
Liang; George |
April 10, 2008 |
Turbine airfoil cooling system with enhanced tip corner cooling
channel
Abstract
A cooling system for a turbine airfoil of a turbine engine
having a serpentine cooling channel with a portion positioned
proximate to an intersection between a trailing edge and the tip
section of the airfoil. The serpentine cooling channel may extend
generally spanwise between a root and the tip section of the
airfoil and may include a portion extending generally chordwise.
The chordwise portion of the serpentine cooling channel may extend
between the spanwise portion of the serpentine cooling channel and
a trailing edge of the airfoil. The chordwise portion may include
an exhaust orifice in the trailing edge and at a trailing edge turn
in the serpentine cooling channel. The trailing edge turn provides
a more efficient cooling channel for the cooling system which
reduces the local temperature of the airfoil at the tip flag.
Inventors: |
Liang; George; (Palm City,
FL) |
Correspondence
Address: |
SIEMENS CORPORATION;INTELLECTUAL PROPERTY DEPARTMENT
170 WOOD AVENUE SOUTH
ISELIN
NJ
08830
US
|
Assignee: |
Siemens Power Generation,
Inc.
|
Family ID: |
39301669 |
Appl. No.: |
11/543523 |
Filed: |
October 5, 2006 |
Current U.S.
Class: |
416/97R |
Current CPC
Class: |
F05D 2250/185 20130101;
F05D 2260/22141 20130101; F01D 5/187 20130101 |
Class at
Publication: |
416/97.R |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Claims
1. A turbine airfoil, comprising: a generally elongated, hollow
airfoil having a leading edge, a trailing edge, a tip section at a
first end, a root coupled to the airfoil at an end generally
opposite to the first end for supporting the airfoil and for
coupling the airfoil to a disc, and a cooling system formed from at
least one cavity in the elongated, hollow airfoil; the cooling
system comprising a serpentine cooling channel positioned in a
mid-chord region of the generally elongated airfoil and being
formed from a first leg, a second leg, a third leg, a fourth leg,
and a fifth leg; wherein the first, fourth and fifth legs are
generally aligned with each other and extend in a generally
spanwise direction; wherein the second and third legs are generally
aligned with each other and extend in a generally chordwise
direction proximate to an intersection between the trailing edge
and the tip section of the generally elongated, hollow airfoil; and
a trailing edge turn between the second and third legs including a
trailing edge exhaust orifice in the trailing edge.
2. The turbine airfoil of claim 1, further comprising a rib
positioned in the trailing edge turn between the second and third
legs.
3. The turbine airfoil of claim 2, wherein the rib extends
generally chordwise and contacts the trailing edge.
4. The turbine airfoil of claim 1, further comprising a trailing
edge cooling channel in contact with the trailing edge and
extending from the root to the second leg of the serpentine cooling
channel.
5. The turbine airfoil of claim 4, further comprising at least one
trailing edge exhaust orifice in fluid communication with the
trailing edge cooling channel.
6. The turbine airfoil of claim 5, further comprising at least one
impingement rib in the trailing edge cooling channel that extends
in a spanwise direction, wherein the impingement rib includes a
plurality of impingement orifices.
7. The turbine airfoil of claim 1, further comprising at least one
leading edge cooling channel positioned proximate to the leading
edge.
8. The turbine airfoil of claim 7, further comprising at least one
leading edge supply channel in fluid communication with the at
least one leading edge cooling channel.
9. The turbine airfoil of claim 8, further comprising a plurality
of impingement orifices in a rib separating the at least one
leading edge supply channel from the at least one leading edge
cooling channel.
10. The turbine airfoil of claim 1, further comprising at least one
trip strip in at least one of the legs of the serpentine cooling
channel.
11. The turbine airfoil of claim 10, wherein the at least one trip
strip protrudes from an outer wall forming a pressure side of the
generally elongated airfoil.
12. The turbine airfoil of claim 11, wherein the at least one trip
strip comprises a plurality of trip strips protruding from the
outer walls forming the pressure side and a suction side of the
generally elongated airfoil.
13. A turbine airfoil, comprising: a generally elongated, hollow
airfoil having a leading edge, a trailing edge, a tip section at a
first end, a root coupled to the airfoil at an end generally
opposite to the first end for supporting the airfoil and for
coupling the airfoil to a disc, and a cooling system formed from at
least one cavity in the elongated, hollow airfoil; the cooling
system comprising a serpentine cooling channel positioned in a
mid-chord region of the generally elongated airfoil and being
formed from a first leg, a second leg, a third leg, a fourth leg,
and a fifth leg; wherein the first, fourth and fifth legs are
generally aligned with each other and extend in a generally
spanwise direction; wherein the second and third legs are generally
aligned with each other and extend in a generally chordwise
direction between a trailing edge cooling channel and an
intersection between the trailing edge and the tip section of the
generally elongated, hollow airfoil; a trailing edge trailing edge
turn between the second and third legs including a trailing edge
exhaust orifice in the trailing edge; the trailing edge cooling
channel in contact with the trailing edge and extending from the
root to the second leg of the serpentine cooling channel and the
trailing edge trailing edge turn; and at least one leading edge
cooling channel positioned proximate to the leading edge.
14. The turbine airfoil of claim 13, further comprising at least
one trailing edge exhaust orifice in fluid communication with the
trailing edge cooling channel.
15. The turbine airfoil of claim 14, further comprising at least
one impingement rib in the trailing edge cooling channel that
extends in a spanwise direction, wherein the impingement rib
includes a plurality of impingement orifices.
16. The turbine airfoil of claim 15, further comprising at least
one leading edge supply channel in fluid communication with the at
least one leading edge cooling channel.
17. The turbine airfoil of claim 16, further comprising a plurality
of impingement orifices in a rib separating the at least one
leading edge supply channel from the at least one leading edge
cooling channel.
18. The turbine airfoil of claim 13, further comprising at least
one trip strip in at least one of the legs of the serpentine
cooling channel, wherein the at least one trip strip comprises a
plurality of trip strips protruding from the outer walls forming
the pressure side and a suction side of the generally elongated
airfoil.
19. The turbine airfoil of claim 13, further comprising a rib
positioned in the trailing edge turn between the second and third
legs.
20. The turbine airfoil of claim 19, wherein the rib extends
generally chordwise and contacts the trailing edge.
Description
FIELD OF THE INVENTION
[0001] This invention is directed generally to turbine airfoils,
and more particularly to cooling systems in hollow turbine
airfoils.
BACKGROUND
[0002] Typically, gas turbine engines include a compressor for
compressing air, a combustor for mixing the compressed air with
fuel and igniting the mixture, and a turbine blade assembly for
producing power. Combustors often operate at high temperatures that
may exceed 2,500 degrees Fahrenheit. Typical turbine combustor
configurations expose turbine blade assemblies to these high
temperatures. As a result, turbine blades must be made of materials
capable of withstanding such high temperatures. In addition,
turbine blades often contain cooling systems for prolonging the
life of the blades and reducing the likelihood of failure as a
result of excessive temperatures.
[0003] Typically, turbine blades are formed from a root portion
having a platform at one end and an elongated portion forming a
blade that extends outwardly from the platform coupled to the root
portion. The blade is ordinarily composed of a tip opposite the
root section, a leading edge, and a trailing edge. The inner
aspects of most turbine blades typically contain an intricate maze
of cooling channels forming a cooling system. The cooling channels
in a blade receive air from the compressor of the turbine engine
and pass the air through the blade. The cooling channels often
include multiple flow paths that are designed to maintain all
aspects of the turbine blade at a relatively uniform temperature.
However, centrifugal forces and air flow at boundary layers often
prevent some areas of the turbine blade from being adequately
cooled, which results in the formation of localized hot spots.
Localized hot spots, depending on their location, can reduce the
useful life of a turbine blade and can damage a turbine blade to an
extent necessitating replacement of the blade. Thus, a need exists
for a cooling system capable of providing sufficient cooling to
turbine airfoils.
SUMMARY OF THE INVENTION
[0004] This invention relates to a turbine airfoil cooling system
for a turbine airfoil used in turbine engines. In particular, the
turbine airfoil cooling system includes a plurality of internal
cavities positioned between outer walls of the turbine airfoil. The
cavity may be formed from a mid-chord serpentine cooling channel
positioned between a leading edge cooling channel and a trailing
edge cooling channel. The mid-chord serpentine cooling channel may
be formed from a spanwise portion extending in a generally spanwise
direction between a root and a tip section and from a chordwise
section extending in a generally chordwise direction between the
spanwise section and a trailing edge of the airfoil. The mid-chord
serpentine cooling channel may facilitate the efficient removal of
heat from the airfoil, especially at the intersection between the
tip section and the trailing edge.
[0005] The turbine airfoil cooling system may be positioned in a
generally elongated, hollow airfoil having a leading edge, a
trailing edge, a tip section at a first end, a root coupled to the
airfoil at an end generally opposite to the first end for
supporting the airfoil and for coupling the airfoil to a disc, and
a cooling system formed from at least one cavity in the elongated,
hollow airfoil. The cooling system may be formed from a serpentine
cooling channel positioned in a mid-chord region of the generally
elongated airfoil. The serpentine cooling channel may be formed
from a first leg, a second leg, a third leg, a fourth leg, and a
fifth leg. The first, fourth and fifth legs may be generally
aligned with each other and may extend in a generally spanwise
direction. The second and third legs may be generally aligned with
each other and may extend in a generally chordwise direction
proximate to an intersection between the trailing edge and the tip
section of the generally elongated, hollow airfoil. The trailing
edge turn may be positioned between the second and third legs and
may include a trailing edge exhaust orifice in the trailing edge. A
rib may be positioned in the trailing edge turn between the second
and third legs. The rib may extend generally chordwise and may
contact the trailing edge.
[0006] A trailing edge cooling channel may be in contact with the
trailing edge and may extend from the root to the second leg of the
serpentine cooling channel. One or more trailing edge exhaust
orifices may be in fluid communication with the trailing edge
cooling channel. One or more impingement ribs may be positioned in
the trailing edge cooling channel that extends in a spanwise
direction, wherein the impingement rib may include a plurality of
impingement orifices. The cooling system may also include one or
more leading edge cooling channels positioned proximate to the
leading edge. One or more leading edge supply channels may be in
fluid communication with the at least one leading edge cooling
channel. A plurality of impingement orifices may be positioned in a
rib separating the at least one leading edge supply channel from
the at least one leading edge cooling channel.
[0007] The cooling system may include one or more trip strips in at
least one of the legs of the serpentine cooling channel, the
leading edge supply channel, the leading edge cooling channel, and
the trailing edge cooling channel. The trip strips may protrude
from an outer wall forming a pressure side of the generally
elongated airfoil. In one embodiment, the trip strip may be formed
by a plurality of trip strips protruding from the outer walls
forming the pressure side and a suction side of the generally
elongated airfoil.
[0008] An advantage of this invention is that the second and third
legs of the serpentine cooling channel utilizes combined cooling
fluids for the airfoil mid-body, blade tip section, and the tip
flag, thereby increasing the total amount of cooling mass flow
through the tip corner region and yields higher internal cooling
capacity for the blade tip flag region at the intersection between
the tip section and the trailing edge.
[0009] Another advantage of this invention is that a large portion
of cooling fluids flows under the tip section, which provides
additional convective cooling enhancement for the blade trailing
edge tip section.
[0010] Still another advantage of this invention is that the
pressure side tip section film cooling holes are angled in a
direction opposite to the serpentine flow, thereby minimizing
internal hole plugging problems.
[0011] These and other embodiments are described in more detail
below.
BRIEF DESCRIPTION OF THE DRAWINGS
[0012] The accompanying drawings, which are incorporated in and
form a part of the specification, illustrate embodiments of the
presently disclosed invention and, together with the description,
disclose the principles of the invention.
[0013] FIG. 1 is a perspective view of a turbine airfoil having
features according to the instant invention.
[0014] FIG. 2 is a cross-sectional view, commonly referred to as a
filleted view, of the turbine airfoil shown in FIG. 1 taken along
line 2-2.
[0015] FIG. 3 is a schematic diagram of the flow pattern of the
cooling channels shown in FIG. 2.
DETAILED DESCRIPTION OF THE INVENTION
[0016] As shown in FIGS. 1-3, this invention is directed to a
turbine airfoil cooling system 10 for a turbine airfoil 12 used in
turbine engines. In particular, the turbine airfoil cooling system
10 includes a plurality of internal cavities 14, as shown in FIG.
2, positioned between outer walls 16 of the turbine airfoil 12. The
cavity 14 may be formed from a mid-chord serpentine cooling channel
18 positioned between a leading edge cooling channel 20 and a
trailing edge cooling channel 22. The mid-chord serpentine cooling
channel 18 may be formed from a spanwise portion 24 extending in a
generally spanwise direction between a root 26 and a tip section 28
and from a chordwise section 30 extending in a generally chordwise
direction between the spanwise section 24 and a trailing edge 32 of
the airfoil 12. The mid-chord serpentine cooling channel 18 may
facilitate the efficient removal of heat from the airfoil 12,
especially at the intersection 34 between the tip section 28 and
the trailing edge 32.
[0017] As shown in FIGS. 1 and 2, the turbine airfoil 12 may be
formed from a generally elongated, hollow airfoil 34 coupled to the
root 26 at a platform 36. The turbine airfoil 12 may be formed from
conventional metals or other acceptable materials. The generally
elongated airfoil 34 may extend from the root 26 to the tip section
28 and may include a leading edge 38 and trailing edge 32. Airfoil
34 may have an outer wall 16 adapted for use, for example, in a
first stage of an axial flow turbine engine. Outer wall 16 may form
a generally concave shaped portion forming pressure side 40 and may
form a generally convex shaped portion forming suction side 42. The
cavity 14, as shown in FIG. 2, may be positioned in inner aspects
of the airfoil 34 for directing one or more gases, which may
include air received from a compressor (not shown), through the
airfoil 34 to reduce the temperature of the airfoil 34.
[0018] The cooling system 10, as shown in FIGS. 2 and 3, may
include a mid-chord serpentine cooling channel 18 forming the
internal cavity 14. The mid-chord serpentine cooling channel 18 may
be formed from a first leg 44, a second leg 46, a third leg 48, a
fourth leg 50, and a fifth leg 52. The legs 44, 46, 48, 50 and 52
may be formed with internal ribs 54, 55, 57. The legs 44, 46, 48,
50 and 52 may extend from the outer wall 16 forming the pressure
side 40 to the outer wall 16 forming a suction side 42. In other
embodiments, the serpentine cooling channel 18 may be formed from
more or less number of legs. The serpentine cooling channel 18 may
extend from close proximity of the tip section 28 of the airfoil 34
to the root 26. In other embodiments, the serpentine cooling
chamber 18 may have a shorter length.
[0019] The serpentine cooling chamber 18 may be in fluid
communication with one or more cooling fluid supply channels 56 in
the root 26. In particular, the first leg 44 may be in fluid
communication with the supply channels 56 such that cooling fluids
flow radially from the root 26 toward the tip section 28. The
serpentine cooling chamber 18 may be positioned such that the first
leg 44 extends in a generally spanwise direction and is positioned
proximate to the trailing edge cooling channel 22. In at least one
embodiment, the first, fourth and fifth legs 44, 50, 52 may be
aligned with each other and may be positioned in a generally
spanwise direction. The first, fourth and fifth legs 44, 50, 52 may
form a spanwise portion 24 of the serpentine cooling chamber 18.
The fourth leg 50 may be positioned between the first leg 44 and
the fifth leg 52. The fourth and fifth legs 50, 52 may be separated
by internal rib 57. The first and fourth legs 44, 50 may be
separated by internal rib 55. The first and second legs 44, 46 may
be in fluid communication through turn 72. The third and fourth
legs 48, 50 may be in fluid communication through turn 74 and the
fourth and fifth legs 50, 52 may be in fluid communication through
turn 76. The turn 74 at the third and fourth legs 48, 50 may
include a gap 86 that is used in the manufacturing process. The gap
86 is sealed with a plate 88.
[0020] As shown in FIG. 2, the second and third legs 46, 48 may be
positioned proximate to the tip section 28 and may extend in a
generally chordwise direction forming the chordwise portion 30. The
second and third legs 46, 48 may be positioned between the trailing
edge cooling channel 22 and the tip section 28. The second and
third legs 46, 48 may be in fluid communication with each other
through a trailing edge turn 58. The trailing edge turn 58 may be
in contact with the trailing edge 32 and may include a trailing
edge exhaust orifice 60 opening the serpentine cooling chamber 18
at the trailing edge 32. The second and third legs 46, 48 may be
separated by the internal rib 54. The internal rib 54 may terminate
in the chordwise direction toward the leading edge 38 from the
trailing edge 32 to form the trailing edge turn 58. The size of the
trailing edge turn 58 may be sized depending on the application of
use of the turbine airfoil 12.
[0021] A rib 62 may be positioned in the trailing edge turn 58. The
rib 62 may extend generally in line with the internal rib 54
separating the first and fourth legs 46, 50 and the second and
third legs 46, 48. The rib 62 may extend generally in a chordwise
direction and may extend from the outer wall 16 forming the
pressure side 40 to the outer wall 16 forming the suction side
42.
[0022] As shown in FIG. 2, the turbine airfoil cooling system 10
may include a trailing edge cooling channel 64 in contact with the
trailing edge 32 and extending from the root 26 to the second leg
46 of the serpentine cooling channel 18. The trailing edge cooling
channel 64 may include one or more impingement ribs 66 extending in
a generally spanwise direction. The impingement ribs 66 may include
one or more impingement orifices 68. The impingement orifices 68 in
adjacent impingement ribs 66 may be aligned or offset from each
other. The trailing edge cooling channel 64 may also include a
plurality of bleed off ribs 70 at the trailing edge 32. The bleed
off ribs 70 may be staggered apart from each other to enable
cooling fluids to be exhausted from the cooling system 10.
[0023] The turbine airfoil cooling system 10 may include one or
more leading edge cooling channels 20 extending generally in a
spanwise direction along the leading edge 38. A leading edge supply
channel 78 may extend proximate to the leading edge cooling channel
20 and may be in fluid communication with the fluid supply channels
56. The leading edge supply channel 20 may be in fluid
communication with the leading edge cooling channels 20 through one
or more orifices 80.
[0024] The cooling system 10 may include one or more trip strips 82
for increasing the cooling capacity of the system 10. In at least
one embodiment, the trip strips 82 may be positioned on inner
surfaces 84 of the mid-chord serpentine cooling channel 18, the
leading edge cooling channel 20, or the trailing edge cooling
channel 22, or any combination thereof. The trips strips 82 may be
positioned on the pressure side 40 or the suction side 42 of the
cooling channels or both. The trip strips 82 may be positioned
orthogonal to the flow of the cooling fluids through the cooling
channels or may be positioned at an angle, as shown in FIG. 2.
[0025] The tip section 28 may include film cooling orifices 90 in
fluid communication with the serpentine cooling channel 18 and the
leading edge supply channel 78. The film cooling orifices 90 may be
angled in a downstream direction to prevent the orifices from
becoming clogged. The film cooling orifices 90 may also extend
through the outer wall 16 forming the pressure and suction sides
40, 42.
[0026] During use, cooling fluids may flow into the cooling system
10 from a cooling fluid supply source (not shown) through the
cooling fluid supply channel 56. In particular, the cooling fluids
may enter into the first leg 44 of the serpentine cooling channel
18 and flow spanwise toward the tip section 28. The fluids may flow
through the turn 72, through the second leg 46 and into the
trailing edge turn 58 where the outer walls 16 proximate to the tip
section 28 and the trailing edge 32 are cooled. Some of the cooling
fluids may be exhausted through the trailing edge exhaust orifice
60 and the remainder of the cooling fluids may pass into the third
leg 48. The cooling fluids may flow through the turn 74, into the
fourth leg 50, through turn 76, into the fifth leg 52 flowing
radially outwardly and be exhausted through film cooling
orifices.
[0027] Cooling fluids may also flow from a cooling fluid supply
channel 56 into the trailing edge cooling channel 64. The cooling
fluids may flow through one or more impingement orifices 68 in the
impingement rib 66 and past the bleed off ribs 70 where the cooling
fluids are exhausted from the turbine airfoil 12. A portion of the
cooling fluids may also be exhausted through film cooling
orifices.
[0028] Cooling fluids may also flow from a cooling fluid supply
channel 56 into the leading edge supply channel 78. Cooling fluids
may flow from the leading edge supply channel 78 into the leading
edge cooling channel 20 through one or more leading edge
impingement orifices 80. The cooling fluids may be exhausted from
the leading edge cooling channel 20 through film cooling orifices
forming a showerhead.
[0029] The foregoing is provided for purposes of illustrating,
explaining, and describing embodiments of this invention.
Modifications and adaptations to these embodiments will be apparent
to those skilled in the art and may be made without departing from
the scope or spirit of this invention.
* * * * *