U.S. patent application number 09/756902 was filed with the patent office on 2002-07-11 for method and apparatus for reducing turbine blade tip temperatures.
Invention is credited to Lee, Ching-Pang, Prakash, Chander, Rinck, Gerard Anthony, Shelton, Monty Lee, Singh, Hardev, Starkweather, John Howard.
Application Number | 20020090301 09/756902 |
Document ID | / |
Family ID | 25045543 |
Filed Date | 2002-07-11 |
United States Patent
Application |
20020090301 |
Kind Code |
A1 |
Lee, Ching-Pang ; et
al. |
July 11, 2002 |
METHOD AND APPARATUS FOR REDUCING TURBINE BLADE TIP
TEMPERATURES
Abstract
A rotor blade for a gas turbine engine including a tip region
that facilitates reducing operating temperatures of the rotor blade
is described. The tip region includes a first tip wall and a second
tip wall extending radially outward from a tip plate of an airfoil.
The tip walls extend from adjacent a leading edge of the airfoil to
connect at a trailing edge of the airfoil. A notch is defined
between the first and second tip walls at the airfoil leading edge.
At least a portion of the second tip wall is recessed to define a
tip shelf.
Inventors: |
Lee, Ching-Pang;
(Cincinnati, OH) ; Prakash, Chander; (Cincinnati,
OH) ; Shelton, Monty Lee; (Loveland, OH) ;
Starkweather, John Howard; (Cincinnati, OH) ; Singh,
Hardev; (Mason, OH) ; Rinck, Gerard Anthony;
(Cincinnati, OH) |
Correspondence
Address: |
JOHN S. BEULICK
C/O ARMSTRONG TEASDALE LLP
ONE METROPOLITAN SQUARE
SUITE 2600
ST. LOUIS
MO
63102-2740
US
|
Family ID: |
25045543 |
Appl. No.: |
09/756902 |
Filed: |
January 9, 2001 |
Current U.S.
Class: |
416/224 ;
416/228; 416/236R; 416/92; 416/97R |
Current CPC
Class: |
F05D 2240/121 20130101;
F01D 5/20 20130101; F05D 2260/202 20130101; F05D 2250/70 20130101;
F01D 5/186 20130101; F05D 2240/303 20130101 |
Class at
Publication: |
416/224 ;
416/228; 416/236.00R; 416/92; 416/97.00R |
International
Class: |
F01D 005/18 |
Claims
What is claimed is:
1. A method for fabricating a rotor blade for a gas turbine engine
to facilitate reducing operating temperatures of a tip portion of
the rotor blade, the rotor blade including a leading edge, a
trailing edge, a first sidewall, and a second sidewall, the first
and second sidewalls connected axially at the leading and trailing
edges, and extending radially between a rotor blade root to a rotor
blade tip plate, said method comprising the steps of: forming a
first tip wall extending from the rotor blade tip plate along the
first sidewall; and forming a second tip wall extending from the
rotor blade tip plate along the second sidewall such that the
second tip wall connects with the first tip wall at the rotor blade
trailing edge, and such that a notch is defined between the first
and second tip walls along the rotor blade leading edge.
2. A method in accordance with claim 1 further comprising the step
of forming a guide wall extending from the rotor blade notch
aftward towards the rotor blade trailing edge such that flow
entering the notch is directed with the guide wall towards the
first sidewall.
3. A method in accordance with claim 1 wherein said step of forming
a first tip wall further comprises the step of recessing at least a
portion of the first tip wall with respect to the rotor blade first
sidewall such that a first tip shelf is defined.
4. A method in accordance with claim 3 wherein said step of forming
a second tip wall further comprises the step of recessing at least
a portion of the second tip wall with respect to the rotor blade
second sidewall such that a second tip shelf is defined.
5. A method in accordance with claim 1 wherein said step of forming
a second tip wall further comprises the step of forming the second
tip wall such that a notch extends from the tip plate and is
defined between the first and second tip walls.
6. An airfoil for a gas turbine engine, said airfoil comprising: a
leading edge; a trailing edge; a tip plate; a first sidewall
extending in radial span between an airfoil root and said tip
plate; a second sidewall connected to said first sidewall at said
leading edge and said trailing edge, said second sidewall extending
in radial span between the airfoil root and said tip plate; a first
tip wall extending radially outward from said tip plate along said
first sidewall; a second tip wall extending radially outward from
said tip plate along said second sidewall, said first tip wall
connected to said second tip wall at said trailing edge; and a
notch extending between said first tip wall and said second tip
wall along said airfoil leading edge.
7. An airfoil in accordance with claim 6 wherein said notch
comprises a guide wall extending from said notch towards said
airfoil trailing edge.
8. An airfoil in accordance with claim 7 wherein said guide wall
configured to channel flow entering said notch towards said first
tip wall.
9. An airfoil in accordance with claim 6 wherein said first tip
wall recessed at least partially from said first sidewall to define
a first tip shelf.
10. An airfoil in accordance with claim 9 wherein said second tip
wall recessed at least partially from said second sidewall to
define a second tip shelf.
11. An airfoil in accordance with claim 6 wherein said first tip
wall and said second tip wall are substantially equal in
height.
12. An airfoil in accordance with claim 6 wherein said first tip
wall extends a first distance from said tip plate, said second tip
wall extends a second distance from said tip plate.
13. An airfoil in accordance with claim 12 wherein said notch
extends from said tip plate at least one of said first distance or
said second distance.
14. A gas turbine engine comprising a plurality of rotor blades,
each said rotor blade comprising an airfoil comprising a leading
edge, a trailing edge, a first sidewall, a second sidewall, a first
tip wall, a second tip wall, and a notch, said airfoil first and
second sidewalls connected axially at said leading and trailing
edges, said first and second sidewalls extending radially from a
blade root to said tip plate, said first tip wall extending
radially outward from said tip plate along said first sidewall,
said second tip wall extending radially outward from said tip plate
along said second sidewall, said notch along said airfoil leading
edge between said first tip wall and said second tip wall, said
notch extending from said tip plate.
15. A gas turbine engine in accordance with claim 14 wherein said
rotor blade airfoil first sidewall is concave, said rotor blade
airfoil second sidewall is convex.
16. A gas turbine engine in accordance with claim 15 wherein said
rotor blade airfoil notch comprises a guide wall extending from
said notch towards said rotor blade trailing edge, said guide wall
configured to channel flow entering said notch towards said first
tip wall.
17. A gas turbine engine in accordance with claim 15 wherein said
rotor blade first tip wall at least partially recessed with respect
to said rotor blade first sidewall to define a first tip shelf
18. A gas turbine engine in accordance with claim 17 wherein said
rotor blade second tip wall at least partially recessed with
respect to said rotor blade second sidewall to define a second tip
shelf
19. A gas turbine engine wherein said rotor blade notch extends
radially outward from said rotor blade tip plate.
20. A gas turbine engine wherein said rotor blade first tip wall
and said rotor blade second tip wall have approximately equal
heights.
Description
BACKGROUND OF THE INVENTION
[0001] This application relates generally to gas turbine engine
rotor blades and, more particularly, to methods and apparatus for
reducing rotor blade tip temperatures.
[0002] Gas turbine engine rotor blades typically include airfoils
having leading and trailing edges, a pressure side, and a suction
side. The pressure and suction sides connect at the airfoil leading
and trailing edges, and span radially between the airfoil root and
the tip. To facilitate reducing combustion gas leakage between the
airfoil tips and stationary stator components, the airfoils include
a tip region that extends radially outward from the airfoil
tip.
[0003] The airfoil tip regions include a first tip wall extending
from the airfoil leading edge to the trailing edge, and a second
tip wall also extending from the airfoil leading edge to connect
with the first tip wall at the airfoil trailing edge. The tip
region prevents damage to the airfoil if the rotor blade rubs
against the stator components.
[0004] During operation, combustion gases impacting the rotating
rotor blades transfer heat into the blade airfoils and tip regions.
Over time, continued operation in higher temperatures may cause the
airfoil tip regions to thermally fatigue. To facilitate reducing
operating temperatures of the airfoil tip regions, at least some
known rotor blades include slots within the tip walls to permit
combustion gases at a lower temperature to flow through the tip
regions.
[0005] To facilitate minimizing thermal fatigue to the rotor blade
tips, at least some known rotor blades include a shelf adjacent the
tip region to facilitate reducing operating temperatures of the tip
regions. The shelf is defined within the pressure side of the
airfoil and disrupt combustion gas flow as the rotor blades rotate,
thus enabling a film layer of cooling air to form against the
pressure side of the airfoil. The film layer insulates the blade
from the higher temperature combustion gases.
BRIEF SUMMARY OF THE INVENTION
[0006] In an exemplary embodiment, a rotor blade for a gas turbine
engine includes a tip region that facilitates reducing operating
temperatures of the rotor blade, without sacrificing aerodynamic
efficiency of the turbine engine. The tip region includes a first
tip wall and a second tip wall that extend radially outward from an
airfoil tip plate. The first tip wall extends from adjacent a
leading edge of the airfoil to a trailing edge of the airfoil. The
second tip wall also extends from adjacent the airfoil leading edge
and connects with the first tip wall at the airfoil trailing edge
to define an open-top tip cavity. At least a portion of the second
tip wall is recessed to define a tip shelf. A notch extends from
the tip plate and is defined between the first and second tip walls
at the airfoil leading edge. The notch is in flow communication
with the tip cavity.
[0007] During operation, as the rotor blades rotate, combustion
gases at a higher temperature near each rotor blade leading edge
migrate to the airfoil tip region. Because the tip walls extend
from the airfoil, a tight clearance is defined between the rotor
blade and stationary structural components that facilitates
reducing combustion gas leakage therethrough. If rubbing occurs
between the stationary structural components and the rotor blades,
the tip walls contact the components and the airfoil remains
intact. As the rotor blade rotates, combustion gases at lower
temperatures near the leading edge flow through the notch and
induce cooler gas temperatures into the tip cavity. The combustion
gases on a pressure side of the rotor blade also flow over the tip
region shelf and mix with film cooling air. As a result, the notch
and shelf facilitate reducing operating temperatures of the rotor
blade within the tip region, but without consuming additional
cooling air, thus improving turbine efficiency.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] FIG. 1 is schematic illustration of a gas turbine
engine;
[0009] FIG. 2 is a partial perspective view of a rotor blade that
may be used with the gas turbine engine shown in FIG. 1;
[0010] FIG. 3 is a cross-sectional view of an alternative
embodiment of the rotor blade shown in FIG. 2; and
[0011] FIG. 4 is a partial perspective view of another alternative
embodiment of a rotor blade that may be used with the gas turbine
engine shown in FIG. 1.
DETAILED DESCRIPTION OF THE INVENTION
[0012] FIG. 1 is a schematic illustration of a gas turbine engine
10 including a fan assembly 12, a high pressure compressor 14, and
a combustor 16. Engine 10 also includes a high pressure turbine 18,
a low pressure turbine 20, and a booster 22. Fan assembly 12
includes an array of fan blades 24 extending radially outward from
a rotor disc 26. Engine 10 has an intake side 28 and an exhaust
side 30.
[0013] In operation, air flows through fan assembly 12 and
compressed air is supplied to high pressure compressor 14. The
highly compressed air is delivered to combustor 16. Airflow (not
shown in FIG. 1) from combustor 16 drives turbines 18 and 20, and
turbine 20 drives fan assembly 12.
[0014] FIG. 2 is a partial perspective view of a rotor blade 40
that may be used with a gas turbine engine, such as gas turbine
engine 10 (shown in FIG. 1). In one embodiment, a plurality of
rotor blades 40 form a high pressure turbine rotor blade stage (not
shown) of gas turbine engine 10. Each rotor blade 40 includes a
hollow airfoil 42 and an integral dovetail (not shown) used for
mounting airfoil 42 to a rotor disk (not shown) in a known
manner.
[0015] Airfoil 42 includes a first sidewall 44 and a second
sidewall 46. First sidewall 44 is convex and defines a suction side
of airfoil 42, and second sidewall 46 is concave and defines a
pressure side of airfoil 42. Sidewalls 44 and 46 are joined at a
leading edge 48 and at an axially-spaced trailing edge 50 of
airfoil 42 that is downstream from leading edge 48.
[0016] First and second sidewalls 44 and 46, respectively, extend
longitudinally or radially outward to span from a blade root (not
shown) positioned adjacent the dovetail to a tip plate 54 which
defines a radially outer boundary of an internal cooling chamber
(not shown). The cooling chamber is defined within airfoil 42
between sidewalls 44 and 46. Internal cooling of airfoils 42 is
known in the art. In one embodiment, the cooling chamber includes a
serpentine passage cooled with compressor bleed air. In another
embodiment, sidewalls 44 and 46 include a plurality of film cooling
openings (not shown), extending therethrough to facilitate
additional cooling of the cooling chamber. In yet another
embodiment, airfoil 42 includes a plurality of trailing edge
openings (not shown) used to discharge cooling air from the cooling
chamber.
[0017] A tip region 60 of airfoil 42 is sometimes known as a
squealer tip, and includes a first tip wall 62 and a second tip
wall 64 formed integrally with airfoil 42. First tip wall 62
extends from adjacent airfoil leading edge 48 along airfoil first
sidewall 44 to airfoil trailing edge 50. More specifically, first
tip wall 62 extends from tip plate 54 to an outer edge 65 for a
height 66. First tip wall height 66 is substantially constant along
first tip wall 62.
[0018] Second tip wall 64 extends from adjacent airfoil leading
edge 48 along second sidewall 46 to connect with first tip wall 62
at airfoil trailing edge 50. More specifically, second tip wall 64
is laterally spaced from first tip wall 62 such that an open-top
tip cavity 70 is defined with tip walls 62 and 64, and tip plate
54. Second tip wall 64 also extends radially outward from tip plate
54 to an outer edge 72 for a height 74. In the exemplary
embodiment, second tip wall height 74 is equal first tip wall
height 66. Alternatively, second tip wall height 74 is not equal
first tip wall height 66.
[0019] A notch 80 is defined between first tip wall 62 and second
tip wall 64 along airfoil leading edge 48. More specifically, notch
80 has a width 82 extending between first and second tip walls 62
and 64, and a height 84 measured between a bottom 86 of notch 80
defined by tip plate 54, and first and second tip wall outer edges
65 and 72, respectively.
[0020] In an alternative embodiment, notch 80 does not extend from
tip plate 54, but instead extends from first and second tip wall
outer edges 65 and 72, respectively, towards tip plate 54 for a
distance (not shown) that is less than notch height 84, and
accordingly, notch bottom 86 is a distance (not shown) from tip
plate 54. In a further alternative embodiment, second tip wall 64
is not connected to first tip wall 62 at airfoil trailing edge 50,
and an opening (not shown) is defined between first tip wall 62 and
second tip wall 64 at airfoil trailing edge 50.
[0021] Notch 80 is in flow communication with open-top tip cavity
70 and permits combustion gas at a lower temperature to enter
cavity 70 for lower heating purposes. In one embodiment, notch 80
also includes a guidewall (not shown in FIG. 2) used to channel
flow entering open-top tip cavity 70 towards second tip wall 64.
More specifically, the guidewall extends from notch 80 towards
airfoil trailing edge 50.
[0022] Second tip wall 64 is recessed at least in part from airfoil
second sidewall 46. More specifically, second tip wall 64 is
recessed from airfoil second sidewall 46 toward first tip wall 62
to define a radially outwardly facing first tip shelf 90 which
extends generally between airfoil leading and trailing edges 48 and
50. More specifically, shelf 90 includes a front edge 94 and an aft
edge 96. Front edge 94 and aft edge 96 each taper to be flush with
second sidewall 46. Shelf front edge 94 is a distance 98 downstream
of airfoil leading edge 48, and shelf aft edge 96 is a distance 100
upstream from airfoil trailing edge 50.
[0023] Recessed second tip wall 64 and shelf 90 define a generally
L-shaped trough 102 therebetween. In the exemplary embodiment, tip
plate 54 is generally imperforate and only includes a plurality of
openings 106 extending through tip plate 54 at tip shelf 90.
Openings 106 are spaced axially along shelf 90 and are in flow
communication between trough 102 and the internal airfoil cooling
chamber. In one embodiment, tip region 60 and airfoil 42 are coated
with a thermal barrier coating.
[0024] During operation, squealer tip walls 62 and 64 are
positioned in close proximity with a conventional stationary stator
shroud (not shown), and define a tight clearance (not shown)
therebetween that facilitates reducing combustion gas leakage
therethrough. Tip walls 62 and 64 extend radially outward from
airfoil 42. Accordingly, if rubbing occurs between rotor blades 40
and the stator shroud, only tip walls 62 and 64 contact the shroud
and airfoil 42 remains intact.
[0025] Because combustion gases assume a parabolic profile flowing
through a turbine flowpath, combustion gases near turbine blade tip
region leading edge 48 are at a lower temperature than gases near
turbine blade tip region trailing edge 50. As cooler combustion
gases flow into notch 80, a heat load of tip region 60 is reduced.
More specifically, combustion gases flowing into notch 80 are at a
higher pressure and reduced temperature than gases leaking from
rotor blade pressure side 46 through the tip clearance to rotor
blade suction side 44. As a result, notch 80 facilitates reducing
an operating temperatures within tip region 60.
[0026] Furthermore, as combustion gases flow past airfoil first tip
shelf 90, trough 102 provides a discontinuity in airfoil pressure
side 46 which causes the combustion gases to separate from airfoil
second sidewall 46, thus facilitating a decrease in heat transfer
thereof. Additionally, trough 102 provides a region for cooling air
to accumulate and form a film against sidewall 46. First tip shelf
openings 106 discharge cooling air from the airfoil internal
cooling chamber to form a film cooling layer on tip region 60.
Because of blade rotation, combustion gases outside rotor blade 40
at leading edge 48 near a blade pitch line (not shown) will migrate
in a radial flow toward airfoil tip region 60 near trailing edge 50
along second sidewall 46 such that leading edge tip operating
temperatures are lower than trailing edge tip operating
temperatures. First tip shelf 90 functions as a backward facing
step in the migrated radial flow and provides a shield for the film
of cooling air accumulated against sidewall 46. As a result, shelf
90 facilitates improving cooling effectiveness of the film to lower
operating temperatures of sidewall 46.
[0027] FIG. 3 is a cross-sectional view of an alternative
embodiment of a rotor blade 120 that may be used with a gas turbine
engine, such as gas turbine engine 10 (shown in FIG. 1). Rotor
blade 120 is substantially similar to rotor blade 40 shown in FIG.
2, and components in rotor blade 120 that are identical to
components of rotor blade 40 are identified in FIG. 3 using the
same reference numerals used in FIG. 2. Accordingly, rotor blade
120 includes airfoil 42 (shown in FIG. 2), sidewalls 44 and 46
(shown in FIG. 2) extending between leading and trailing edges 48
and 50, respectively, and notch 80. Furthermore, rotor blade 120
includes second tip wall 64 and first tip shelf 90. Additionally,
rotor blade 120 includes a first tip wall 122. Notch 80 is defined
between first and second tip walls 122 and 64, respectively.
[0028] First tip wall 122 extends from adjacent airfoil leading
edge 48 along first sidewall 44 to connect with second tip wall 64
at airfoil trailing edge 50. More specifically, first tip wall 122
is laterally spaced from second tip wall 64 to define opentop tip
cavity 70. First tip wall 122 also extends a height (not shown)
radially outward from tip plate 54 to an outer edge 126. In the
exemplary embodiment, the first tip wall height is equal second tip
wall height 74. Alternatively, the first tip wall height is not
equal second tip wall height 74.
[0029] First tip wall 122 is recessed at least in part from airfoil
first sidewall 44. More specifically, first tip wall 122 is
recessed from airfoil first sidewall 44 toward second tip wall 64
to define a radially ouwardly facing second tip shelf 130 which
extends generally between airfoil leading and trailing edges 48 and
50. More specifically, shelf 130 includes a front edge 134 and an
aft edge 136. Front edge 134 and aft edge 136 each taper to be
flush with first sidewall 44. Shelf front edge 134 is a distance
138 downstream of airfoil leading edge 48, and shelf aft edge 136
is a distance 140 upstream from airfoil trailing edge 50.
[0030] Recessed first tip wall 122 and second tip shelf 130 define
therebetween a generally L-shaped trough 144. In the exemplary
embodiment, tip plate 54 is generally imperforate and includes
plurality of openings 106 extending through tip plate 54 at first
tip shelf 90, and a plurality of openings 146 extending through tip
plate 54 at second tip shelf 130. Openings 146 are spaced axially
along second tip shelf 130 and are in flow communication between
trough 144 and the internal airfoil cooling chamber. In one
embodiment, tip region 62 and airfoil 42 are coated with a thermal
barrier coating.
[0031] During operation, squealer tip walls 122 and 64 are
positioned in close proximity with a conventional stationary stator
shroud (not shown), and define a tight clearance (not shown)
therebetween to facilitate reducing combustion gas leakage
therethrough. Tip wall 122 functions in an identical manner as tip
wall 62 described above, and extends radially outward from airfoil
42. Accordingly, if rubbing occurs between rotor blades 40 and the
stator shroud, only tip walls 122 and 64 contact the shroud and
airfoil 42 remains intact.
[0032] Furthermore, as rotor blades 40 rotate and combustion gases
flow past airfoil tip shelves 90 and 130, troughs 102 and 144,
respectively provide a discontinuity in airfoil pressure side 46
and airfoil suction side 44, respectively, which causes the
combustion gases to separate from airfoil sidewalls 46 and 44,
respectively, thus facilitating a decrease in heat transfer thereof
Trough 144 functions similarly with trough 102 to facilitate film
cooling circulation..
[0033] FIG. 4 is a partial perspective view of an alternative
embodiment of a rotor blade 200 that may be used with a gas turbine
engine, such as gas turbine engine 10 (shown in FIG. 1). Rotor
blade 200 is substantially similar to rotor blade 40 shown in FIG.
2, and components in rotor blade 200 that are identical to
components of rotor blade 40 are identified in FIG. 4 using the
same reference numerals used in FIG. 2. Accordingly, rotor blade
200 includes airfoil 42, sidewalls 44 and 46 extending between
leading and trailing edges 48 and 50, respectively, and notch 80.
Furthermore, rotor blade 200 includes first tip wall 62, notch 80,
and a second tip wall 202. Notch 80 is defined between first and
second tip walls 62 and 202, respectively.
[0034] Second tip wall 202 extends from adjacent airfoil leading
edge 48 along airfoil first sidewall 44 to airfoil trailing edge
50. More specifically, second tip wall 202 extends from tip plate
54 to an outer edge 204 for a height (not shown). The second tip
wall height is substantially constant along second tip wall 202.
Second tip wall 202 is laterally spaced from first tip wall 62 to
define open-top tip cavity 70 in the exemplary embodiment, the
second tip wall height is equal first tip wall height 66.
Alternatively, the second tip wall height is not equal first tip
wall height 66.
[0035] Notch 80 includes a guidewall 210 extending from first tip
wall 62 towards airfoil trailing edge. More specifically, guidewall
210 curves to extend from first tip wall 62 to define a curved
entrance 212 for notch 80. Guidewall 210 has a length 214 that is
selected to channel airflow entering open-top tip cavity 70 towards
second tip wall 202.
[0036] The above-described rotor blade is cost-effective and highly
reliable. The rotor blade includes a leading edge notch defined
between leading edges of first and second tip walls. The tip walls
connect at a trailing edge of the rotor blade and define a tip
cavity. In the exemplary embodiment, one of the tip walls is
recessed to define a tip shelf. During operation, as the rotor
blade rotates, the tip walls prevent the rotor blade from rubbing
against stationary structural members. As combustion gases flow
past the rotor blade, the rotor blade notch facilitates lowering
heating of the tip cavity without increasing cooling air
requirements and sacrificing aerodynamic efficiency of the rotor
blade. Furthermore, the tip shelf disrupts combustion gases flowing
past the airfoil to facilitate a cooling layer being formed against
the shelf. As a result, cooler operating temperatures within the
rotor blade facilitate extending a useful life of the rotor blades
in a cost-effective and reliable manner.
[0037] While the invention has been described in terms of various
specific embodiments, those skilled in the art will recognize that
the invention can be practiced with modification within the spirit
and scope of the claims.
* * * * *