Information
-
Patent Grant
-
6758045
-
Patent Number
6,758,045
-
Date Filed
Friday, August 30, 200221 years ago
-
Date Issued
Tuesday, July 6, 200420 years ago
-
Inventors
-
Original Assignees
-
Examiners
Agents
- Andes; William Scott
- Armstrong Teasdale LLP
- Reeser, III; Robert B.
-
CPC
-
US Classifications
Field of Search
US
- 060 804
- 060 737
- 060 747
- 060 748
- 060 752
- 060 772
-
International Classifications
-
Abstract
A method enables a gas turbine engine multi-domed combustor including an outer liner and an inner liner that define a combustion chamber therebetween to be assembled. The method comprises coupling a first dome including a heat shield that includes an annular endbody that extends a first distance axially from the heat shield to the combustor outer liner, and coupling a second dome including a heat shield that includes an annular endbody that extends a second distance axially from the heat shield to the first dome, such that the second dome is radially aligned with respect to the first dome, and wherein the second dome second distance is less than the first dome first distance.
Description
BACKGROUND OF THE INVENTION
This application relates generally to gas turbine engines and, more particularly, to combustors for gas turbine engines.
At least some known gas turbine engines include annular combustors which facilitate reducing nitrogen oxide emissions during gas turbine engine operation. Because of the heat generated within such combustors during operation, at least some known multiple annular combustors include a plurality of multiple dome assemblies that are radially aligned between the combustor dome plate and the combustion chamber. Each dome assembly includes a heat shield to protect the dome plate from excessive heat generated during engine operation.
At least some known dome assembly heat shields include annular endbodies that extend an axial distance downstream from the heat shield to separate the domes or stages of the combustor to enable primary dilution air to be directed into a pilot stage reaction zone, thus facilitating combustion stability of the pilot stage of combustion at various operating points. However, because the endbodies extend axially towards the combustion chamber, the endbodies are exposed to a high temperature and high acoustic energy environment. Over time, the combination of the high temperatures and high acoustic energy may induce thermal stresses, low cycle fatigue (LCF), and/or high cycle fatigue (HCF) into the heat shield assembly. Continued operation with such stresses may lead to cracking within the heat shield which may shorten the useful life of the combustor.
To facilitate reducing the effects of exposure to the high temperature and high acoustic energy environment, at least some known heat shield assemblies have employed various design changes to facilitate improving heat shield durability by addressing thermal and LCF failures. Such improvements have included for example, increased impingement cooling flow, surface film cooling, material changes, and/or heat shield contour changes to attempt to stiffen the component. However, such improvements did not completely address HCF failures caused by combustor acoustics. More specifically, due to engine-to-engine operating variation, and manufacturing/assembly tolerances, despite the improvements, at least some known heat shield natural frequencies remain within the combustor acoustic operating range, and over time, may still experience failures due to HCF fatigue.
BRIEF SUMMARY OF THE INVENTION
In one aspect, a method for assembling a gas turbine engine multi-domed combustor including an outer liner and an inner liner that define a combustion chamber therebetween is provided. The method comprises coupling a first dome including a heat shield that includes an annular endbody that extends a first distance axially from the heat shield to the combustor outer liner, and coupling a second dome including a heat shield that includes an annular endbody that extends a second distance axially from the heat shield to the first dome, such that the second dome is radially aligned with respect to the first dome, and wherein the second dome second distance is less than the first dome first distance.
In another aspect of the invention, an annular combustor for a gas turbine engine is provided. The combustor includes an outer liner, an inner liner, a first dome, and a second dome. The inner liner is spaced radially inwardly from the outer liner to define a combustion chamber therebetween. The first dome includes an outer end coupled to the outer liner and a heat shield including an annular endbody that extends outwardly a first distance axially from the heat shield towards the combustion chamber. The second dome is spaced radially inwardly from, and radially aligned with respect to the first dome. The second dome includes an outer end coupled to an inner end of the first dome, and a heat shield including at least one annular endbody that extends outwardly a second distance from the second dome heat shield. The second distance is less than the first dome first distance.
In a further aspect, a gas turbine engine including a combustor having a natural combustor acoustic operating range is provided. The combustor includes an outer liner, an inner liner, and a plurality of radially-aligned domes. The outer liner is coupled to the inner liner to define a combustion chamber therebetween. The plurality of domes include at least a first dome and a second dome. The first dome includes a heat shield including an annular endbody that extends a first axial distance from the first dome heat shield. The second dome is radially inward from the first dome and includes a heat shield including an annular endbody extending a second axial distance from the first dome heat shield. The second axial distance is less than the first dome first distance.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1
is a schematic illustration of a gas turbine engine;
FIG. 2
is a cross-sectional view of a combustor that may be used with the gas turbine engine shown in
FIG. 1
; and
FIG. 3
is an enlarged cross-sectional view of a portion of the combustor shown in FIG.
2
.
DETAILED DESCRIPTION OF THE INVENTION
FIG. 1
is a schematic illustration of a gas turbine engine
10
including a low pressure compressor
12
, a high pressure compressor
14
, and a combustor
16
. Engine
10
also includes a high pressure turbine
18
and a low pressure turbine
20
. Combustor
16
is a lean premix combustor. Compressor
12
and turbine
20
are coupled by a first shaft
21
, and compressor
14
and turbine
18
are coupled by a second shaft
22
. A load (not shown) may also be coupled to gas turbine engine
10
with first shaft
21
. In one embodiment, gas turbine engine
10
is an LM6000 available from General Electric Aircraft Engines, Cincinnati, Ohio.
In operation, air flows through low pressure compressor
12
and compressed air is supplied from low pressure compressor
12
to high pressure compressor
14
. The highly compressed air is delivered to combustor
30
. Airflow from combustor
16
drives turbines
18
and
20
and exits gas turbine engine
10
through a nozzle
24
.
FIG. 2
is a cross-sectional view of a combustor
30
that may be used with gas turbine engine
10
.
FIG. 3
is an enlarged cross-sectional view of a portion of combustor
30
. Because a fuel/air mixture supplied to combustor
30
contains more air than is required to fully combust the fuel, and because the air is mixed with the fuel prior to combustion, combustor
30
is a lean premix combustor. Accordingly, a fuel/air mixture equivalence ratio for combustor
30
is less than one. Furthermore, because a gas and a liquid fuel are supplied to combustor
30
, and because combustor
30
does not include water injection, combustor
30
is a dual fuel dry low emissions combustor.
Combustor
30
includes an annular outer liner
40
, an annular inner liner
42
, and a domed end or dome plate
44
extending between outer and inner liners
40
and
42
, respectively. Outer liner
40
and inner liner
42
are spaced radially inward from a combustor casing
45
and define a combustion chamber
46
. Combustor casing
45
is generally annular and extends downstream from a diffuser
48
. Combustion chamber
46
is generally annular in shape and is disposed radially inward from liners
40
and
42
. Outer liner
40
and combustor casing
45
define an outer passageway
52
and inner liner
42
and combustor casing
45
define an inner passageway
54
. Outer and inner liners
40
and
42
extend to a turbine nozzle
55
disposed downstream from diffuser
48
.
Combustor domed end
44
includes a plurality of domes
56
. In the exemplary embodiment, domes
56
are arranged in a triple annular configuration. Alternatively, combustor domed end
44
includes a double annular configuration. An outer dome
58
includes an outer end
60
fixedly attached to combustor outer liner
40
and an inner end
62
fixedly attached to a middle dome
64
. Middle dome
64
includes an outer end
66
attached to outer dome inner end
62
and an inner end
68
attached to an inner dome
70
. Accordingly, middle dome
64
is between outer and inner domes
58
and
70
, respectively. Inner dome
70
includes an outer end
72
attached to middle dome inner end
68
and an inner end
74
fixedly attached to combustor inner liner
42
.
Each dome
56
includes a plurality of premixer cups
80
to permit uniform mixing of fuel and air therein and to channel the fuel/air mixture into combustion chamber
46
. In one embodiment, premixer cups
80
are available from Parker Hannifin, 6035 Parkland Blvd., Cleveland, Ohio. Combustor domed end
44
also includes an outer dome heat shield
100
, a middle dome heat shield
102
, and an inner dome heat shield
104
to insulate each respective dome
58
,
64
, and
70
from heat generated within combustion chamber
46
. Heat shields
100
,
102
, and
104
are radially aligned within engine
10
.
Outer dome heat shield
100
includes an annular endbody
106
to insulate combustor outer liner
40
from flames burning in an outer primary combustion zone
108
. Endbody
106
extends outwardly an axial distance
110
from a downstream side
112
of heat shield
100
towards combustion chamber
46
. Distance
110
is commonly known as a heat shield wing length. In one embodiment, distance
110
is approximately equal 1.95 inches. In the exemplary embodiment, endbody
106
extends substantially perpendicularly from heat shield
100
.
Middle dome heat shield
102
includes annular heat shield centerbodies
120
and
122
to segregate middle dome
64
from outer and inner domes
58
and
70
, respectively. Middle dome heat shield centerbodies
120
and
122
are positioned radially outwardly from a middle primary combustion zone
114
, and each extends outwardly an axial distance
126
and
128
, respectively, from a downstream side
130
of heat shield
102
towards combustion chamber
46
. In the exemplary embodiment, endbodies
120
and
122
each extend substantially perpendicularly from heat shield
102
, and as such are substantially parallel outer dome heat shield endbody
106
.
Middle dome heat shield distance
126
is approximately equal distance
128
. Endbody distances
126
and
128
are shorter than outer dome heat shield endbody length
110
. More specifically, middle dome endbody distances
126
and
128
are at least 0.5 inches shorter than outer dome heat shield endbody length
110
. In the exemplary embodiment, middle dome endbody distances are each equal approximately 1.25 inches.
Inner dome heat shield
104
includes an annular endbody
140
to insulate combustor inner liner
42
from flames burning in an inner primary combustion zone
142
. Endbody
140
extends outwardly an axial distance
144
from a downstream side
146
of heat shield
100
towards combustion chamber
46
. Endbody distance
144
is approximately equal outer dome heat shield distance
110
. In one embodiment, endbody distance
144
is approximately equal 1.95 inches. In the exemplary embodiment, endbody
106
extends substantially perpendicularly from heat shield
100
.
During operation of gas turbine engine
10
, as combustor
30
uses radial fuel flow staging to facilitate reducing NOx and CO emissions over the engine operating range, combustor
30
has a natural acoustic operating range. Middle dome heat shield endbodies
120
and
122
facilitate providing additional structural support to middle dome
56
. Specifically, because heat shield endbodies
120
and
122
have a shorted winglength
126
and
128
than outer dome and inner dome endbodies
106
and
140
, respectively, middle dome endbodies
120
and
122
facilitate increasing a stiffness of middle dome heat shield
102
such that the natural frequency of middle dome heat shield
102
is increased above that of the combustor natural acoustic operating range, without adversely impacting engine operability. More specifically, the shortened winglength
126
and
128
does not adversely impact NOx and/or CO emissions, but does facilitate reducing vibrational stresses that may be induced to middle dome
56
. As such, middle dome endbodies
120
and facilitate extending a useful life of combustor
30
.
The above-described combustor system for a gas turbine engine is cost-effective and reliable. The combustor system includes a combustor including a heat shield that includes at least one endbody that has a shortend winglength in comparison to the other heatshield endbodies. The shortened winglength facilitates reducing vibrational stresses that may be induced to the dome assembly by increasing the natural frequency of the endbody above that of the combustor acoustic operating range, but without adversely affecting engine operability. As a result, the endbody facilitates extending a useful life of the combustor in cost effective and reliable manner.
Exemplary embodiments of combustor assemblies are described above in detail. The systems are not limited to the specific embodiments described herein, but rather, components of each assembly may be utilized independently and separately from other components described herein. Each combustor assembly component can also be used in combination with other combustor assembly components
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 assembling a gas turbine engine multi-domed combustor including an outer liner and an inner liner that define a combustion chamber therebetween, said method comprising:coupling a first dome including a heat shield that includes an annular endbody that extends a first distance axially from the heat shield to the combustor outer liner; and coupling a second dome including a heat shield that includes an annular endbody that extends a second distance axially from the heat shield to the first dome, such that the second dome is radially aligned with respect to the first dome, and wherein the second dome second distance is less than the first dome first distance.
- 2. A method in accordance with claim 1 wherein coupling a second dome including a heat shield further comprises coupling a second dome including an annular endbody that extends outwardly a second distance from the heat shield, wherein the second distance is at least approximately 0.5 inches shorter than the first dome first distance.
- 3. A method in accordance with claim 1 wherein coupling a second dome including a heat shield further comprises coupling a second dome including an annular endbody that extends outwardly a second distance from the heat shield that is approximately equal 1.25 inches.
- 4. A method in accordance with claim 1 wherein coupling a second dome including a heat shield further comprises coupling a second dome including an annular endbody that extends outwardly a second distance from the heat shield, wherein the second distance is less than approximately 1.25 inches.
- 5. A method in accordance with claim 1 wherein coupling a second dome including a heat shield further comprises coupling a second dome including an annular endbody that extends outwardly a second distance from the heat shield to facilitate increasing a natural frequency of the annular endbody above a combustor natural acoustic operating range.
- 6. An annular combustor for a gas turbine engine, said combustor comprising:an outer liner; an inner liner spaced radially from said outer liner to define a combustion chamber therebetween; a first dome comprising an outer end coupled to said outer liner and a heat shield comprising an annular endbody extending outwardly a first distance axially from said heat shield towards said combustion chamber; and a second dome spaced radially inwardly from, and radially aligned with respect to said first dome, said second dome comprising an outer end coupled to an inner end of said first dome, and a heat shield comprising at least one an annular endbody extending outwardly a second distance from said second dome heat shield, said second distance less than said first dome first distance.
- 7. A combustor in accordance with claim 6 wherein said second dome second distance is at least approximately 0.5 inches shorter than said first dome first distance.
- 8. A combustor in accordance with claim 6 wherein said second dome second distance is less than approximately 1.50 inches.
- 9. A combustor in accordance with claim 6 wherein said second dome second distance is approximately equal 1.25 inches.
- 10. A combustor in accordance with claim 6 further comprising a third dome spaced radially inwardly from said second dome, said third dome radially aligned with respect to said first and second domes, said third dome comprising an outer end coupled to an inner end of said second dome, and a heat shield comprising an endbody extending outwardly a third distance from said third heat shield, said second dome second distance less than said third dome third distance.
- 11. A combustor in accordance with claim 6 wherein said second dome further comprises a plurality of annular endbodies extending outwardly a second distance from said second dome heat shield.
- 12. A combustor in accordance with claim 6 wherein the combustor has a natural acoustic operating range, said second dome at least one annular endbody facilitates increasing a natural frequency of said at least one annular endbody above the combustor natural acoustic operating range.
- 13. A gas turbine engine comprising a combustor having a natural combustor acoustic operating range, said combustor comprising an outer liner, an inner liner, and a plurality of radially-aligned domes, said outer liner coupled to said inner liner to define a combustion chamber therebetween, said plurality of domes comprising at least a first dome and a second dome, said first dome comprising a heat shield comprising an annular endbody extending a first axial distance from said first dome heat shield, said second dome radially inward from said first dome and comprising a heat shield comprising an annular endbody extending a second axial distance from said first dome heat shield, said second axial distance less than said first dome first distance.
- 14. A gas turbine engine in accordance with claim 13 wherein said combustor second dome second distance configured to facilitate increasing a natural frequency of said second dome endbody above the combustor natural acoustic operating range.
- 15. A gas turbine engine in accordance with claim 14 wherein said combustor second dome second distance is at least approximately 0.5 inches shorter than said first dome first distance.
- 16. A gas turbine engine in accordance with claim 14 wherein said combustor second dome second distance is less than approximately 1.50 inches.
- 17. A gas turbine engine in accordance with claim 14 wherein said combustor second dome second distance is approximately equal 1.25 inches.
- 18. A gas turbine engine in accordance with claim 14 wherein said combustor second dome further comprises a plurality of annular endbodies extending outwardly a second distance from said second dome heat shield.
- 19. A gas turbine engine in accordance with claim 14 wherein said combustor further comprises a third dome spaced radially inwardly from said second dome, said third dome comprising an endbody extending outwardly a third distance from said third heat shield, said third distance approximately equal said first dome first distance.
- 20. A gas turbine engine in accordance with claim 14 wherein said combustor second dome endbody configured to facilitate extending a useful life of said combustor.
US Referenced Citations (13)