Information
-
Patent Grant
-
6422821
-
Patent Number
6,422,821
-
Date Filed
Tuesday, January 9, 200123 years ago
-
Date Issued
Tuesday, July 23, 200221 years ago
-
Inventors
-
Original Assignees
-
Examiners
- Look; Edward K.
- McCoy; Kimya N.
Agents
- Andes; William Scott
- Armstrong Teasdale LLP
-
CPC
-
US Classifications
Field of Search
US
- 416 228
- 416 236 R
- 416 92
- 416 97 R
- 416 224
- 416 1734
- 416 96 R
- 415 115
- 415 116
-
International Classifications
-
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.
Description
BACKGROUND OF THE INVENTION
This application relates generally to gas turbine engine rotor blades and, more particularly, to methods and apparatus for reducing rotor blade tip temperatures.
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.
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.
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.
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
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.
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
FIG. 1
is schematic illustration of a gas turbine engine;
FIG. 2
is a partial perspective view of a rotor blade that may be used with the gas turbine engine shown in
FIG. 1
;
FIG. 3
is a cross-sectional view of an alternative embodiment of the rotor blade shown in
FIG. 2
; and
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
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
.
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
.
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.
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
.
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.
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
.
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
.
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.
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
.
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
.
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
.
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.
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.
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
.
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
.
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.
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 open-top 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
.
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 outwardly 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
.
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.
Second tip wall
202
extends from adjacent airfoil leading edge
48
along airfoil first sidewall
46
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
.
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..
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.
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
.
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
.
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.
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.
Claims
- 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 afterward 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 is 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 is 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, and connected to said first tip wall at said trailing edge, 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 in accordance with claim 15 wherein said rotor blade notch extends radially outward from said rotor blade tip plate.
- 20. A gas turbine engine in accordance with claim 15 wherein said rotor blade first tip wall and said rotor blade second tip wall have approximately equal heights.
US Referenced Citations (3)
Number |
Name |
Date |
Kind |
5261789 |
Butts et al. |
Nov 1993 |
A |
5503527 |
Lee et al. |
Apr 1996 |
A |
6059530 |
Lee |
May 2000 |
A |