IMPINGEMENT COOLING OF COMBUSTOR LINERS
A gas turbine engine may include an impingement cooled double-walled liner, having an inner liner and an outer liner, disposed around a combustion space of the turbine engine. The double-walled liner may extend from an upstream end to a downstream end. The gas turbine engine may also include a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner. Each nozzle of the plurality of nozzles may extend radially inwards from a first distal end to a second proximal end. The plurality of nozzles may be arranged such that a radial gap between the second end of a nozzle and the outer liner decreases from the upstream end to the downstream end. The at least one nozzle of the plurality of nozzles may include multiple air holes at the second end.
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The present disclosure relates generally to systems and methods of impingement cooling a combustor liner of a gas turbine engine.
BACKGROUNDCombustor liners of gas turbine engines are exposed to high temperatures of combustion and therefore require cooling. A type of combustor liner, called a double-walled liner, includes an inner liner that encloses a volume where combustion occurs and an outer liner that surrounds the inner liner. An annular space between the inner liner and the outer liner assists in the cooling of the liner. There are various methods that are employed to cool combustion liners during operation of the engine. These method include film cooling and jet impingement cooling. In film cooling, air in the annular space is directed into the combustor through holes in the inner liner to mix with the hot combustion gases within. The air absorbs the heat from the inner liner as it flows therethrough. In jet impingement cooling, air jets impinge upon and cool the back surface of the inner liner. These air jets may be directed to the back surface of the inner liner through an array of holes on the outer liner. After impinging on the back surface of the inner liner, the spent cooling air flows downstream through the annular space. This spent air flow, called cross-flow, is known to degrade the cooling ability of downstream air jets.
U.S. Patent Application No. 2008/0271458 to Ekkad et al. (the '458 publication) describes an impingement cooled liner with ports extending from the outer liner to the inner liner to reduce the effects of cross-flow. While the extended ports of the '458 publication may reduce the effects of cross-flow, they may have limitations. For instance, dimensional changes during operation of the turbine engine may force portions of the inner liner against the extended ports preventing air flow therethrough. The systems and methods of the current disclosure are directed to overcoming one or more of the problems set forth above.
SUMMARYIn one aspect, a gas turbine engine is disclosed. The gas turbine engine may include an impingement cooled double-walled liner, having an inner liner and an outer liner, disposed around a combustion space of the turbine engine. The double-walled liner may extend from an upstream end to a downstream end. The gas turbine engine may also include a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner. Each nozzle of the plurality of nozzles may extend radially inwards from a first distal end to a second proximal end. The plurality of nozzles may be arranged such that a radial gap between the second end of a nozzle and the outer liner decreases from the upstream end to the downstream end. The at least one nozzle of the plurality of nozzles may include multiple air holes at the second end.
In another aspect, a method of impingement cooling a double-walled combustor liner of a gas turbine engine is disclosed. The double-walled liner may extend from an upstream end to a downstream end and include an inner liner and an outer liner positioned radially outwards the inner liner. The method may include combusting a fuel in a combustor of the gas turbine engine, and directing cooling air through a plurality of nozzles that extend radially inwards through the outer liner to impinge upon and cool the inner liner. The cooling air may be directed such that the cooling air exits the plurality of nozzles closer to the inner liner at the downstream end than at the upstream end. The cooling air directed through at least one nozzle of the plurality of nozzles may exit the at least one nozzle through multiple air flow paths symmetrically arranged about a longitudinal axis of the at least one nozzle.
In yet another aspect, a gas turbine engine is disclosed. The gas turbine engine may include an impingement cooled double-walled liner. The double-walled liner may include an inner liner and an outer liner disposed around a combustion space of the turbine engine and extend from an upstream end to a downstream end. The gas turbine engine may also include a plurality of nozzles that extend radially inwards through the outer liner to direct cooling air towards the inner liner. Each nozzle of the plurality of nozzles may extend radially inwards from a first distal end to a second proximal end. Each nozzle of the plurality of nozzles may include multiple air holes arranged in a shower head pattern at the second end.
The combustion space 58 is fluidly coupled to turbine system 70 at the downstream end. The plurality of fuel injectors 30, positioned on the dome assembly, direct the fuel-air mixture to the combustion space 58 for combustion. This fuel-air mixture burns in a combustion zone (proximate the upstream end) of the combustion space 58 to produce high pressure combustion gases that flow downstream towards the turbine system 70. The combustion of fuel-air mixture within the combustion space 58 heats the outer and the inner combustor walls (22, 24). For increased reliability and performance of GTE 100, it is desirable to cool these walls. The outer combustor wall 22 includes an inner liner 22b and an outer liner 22a, and the inner combustor wall 24 includes an inner liner 24b and an outer liner 24a. The inner liners 22b, 24b are radially spaced apart from the outer liners 22a, 24a to define annular cooling spaces 26, 28 between them. These cooling spaces 26, 28 extend from an upstream end 44 to a downstream end 46 of the combustor 50. The combustion in the combustion space 58 may create oscillations of pressure (pressure waves) within the combustion space 58 that causes radial expansion and contraction (bulging) of the inner liners 22b, 24b with respect to the outer liners 22a, 24a. The outer liners 22a, 24a include a plurality perforations 32, 34 that direct high pressure air from the enclosure 72 to impinge on, and cool, the inner liners 22b, 24b. This technology of impingement cooling the combustor liners is referred to in the industry as Augmented Backside Cooled (ABC) technology. It is known that the use of ABC technology decreases the emission of pollutants into the atmosphere. It should be noted that the general configuration of combustor system 20 illustrated in
Although
In some embodiments, as illustrated in
In some applications, the pressure pulses generated in the combustion space 58 during combustion may cause portions of the inner liner 22b to bulge outwards toward the nozzles 48 in corresponding portions of the outer liner 22a. Contact between the inner liner 22b and a nozzle 48 may restrict, or even block, air flow (air jets 36) through the nozzle 48, and result in uneven cooling of the liner. Some embodiments of the nozzles 48 of the current disclosure may be configured to allow the air flow to continue even when they are in contact with the inner liner 22b.
In some embodiments, nozzle 48A may include one or more projections 74 that project outwards from the second end 68. In some embodiments, at least one of these projections 74 may be located between the outlets of the multiple air holes 62 at the second end 68. Other projections (if any) may be located anywhere on, or proximate, the second end 68. For instance, in some embodiments, the projections 74 may be substantially evenly distributed on the second end 68 of the nozzle 48A. These projections 74 may contact a bulging inner liner 22b, 24b and act as a standoff to allow air flow through the air holes 62 of the nozzle 48A. The projections 74 may have any shape and size. For instance, in some embodiments, arc-shaped projections may extend towards the inner liner 22b, 24b from the periphery of nozzle 48A. In some embodiments, as illustrated in
In some embodiments (as illustrated in nozzles 48A and 48B of
Any type of nozzle (such as, for example nozzles 48, 48A, 48B, 48C, etc.) may be used in an application. In some applications, a nozzle having one air hole 62 (such as nozzle 48 of
The disclosed systems and methods of impingement cooling a cylinder liner may be applicable to any turbine engine to reliably and effectively cool the cylinder liner. The disclosed system of impingement cooling is configured to prevent the impingement air flow from being blocked as a result of dimensional changes of the combustor liner during operation of the turbine engine. The operation of a gas turbine engine using a disclosed system of impingement cooling will now be explained.
With reference to
It will be apparent to those skilled in the art that various modifications and variations can be made to the disclosed impingement cooling system and method. Other embodiments will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed cooling system. It is intended that the specification and examples be considered as exemplary only, with a true scope being indicated by the following claims and their equivalents.
Claims
1. A gas turbine engine, comprising:
- an impingement cooled double-walled liner, including an inner liner and an outer liner, disposed around a combustion space of the turbine engine and extending from an upstream end to a downstream end; and
- a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner, each nozzle of the plurality of nozzles extending radially inwards from a first distal end to a second proximal end, the plurality of nozzles being arranged such that a radial gap between the second end of a nozzle and the outer liner decreases from the upstream end to the downstream end, wherein, at least one nozzle of the plurality of nozzles includes multiple air holes at the second end.
2. The gas turbine engine of claim 1, wherein the at least one nozzle further includes a longitudinal axis extending from the first end to the second end and each air hole of the multiple air holes includes a central axis, the multiple air holes being symmetrically arranged about the longitudinal axis.
3. The gas turbine engine of claim 2, wherein the central axis of each air hole is substantially parallel to the longitudinal axis.
4. The gas turbine engine of claim 2, wherein the central axis of each air hole is inclined with respect to the longitudinal axis such that the cooling air exiting the at least one nozzle diverges.
5. The gas turbine engine of claim 1, wherein the multiple air holes in the at least nozzle is arranged in a shower head pattern at the second end.
6. The gas turbine engine of claim 1, wherein the second end of the at least one nozzle includes a projection that extends towards the inner liner.
7. The gas turbine engine of claim 6, wherein the projection is centrally positioned on the second end and each air hole of the multiple air holes is symmetrically positioned about the projection.
8. The gas turbine engine of claim 1, wherein the second end of the at least one nozzle is curved such that a central portion of the second end forms a proximal-most portion of the nozzle.
9. The gas turbine engine of claim 1, wherein the radial gap decreases substantially linearly from the upstream end to the downstream end.
10. A method of impingement cooling a double-walled combustor liner of a gas turbine engine, the double-walled liner extending from an upstream end to a downstream end and including an inner liner and an outer liner positioned radially outwards the inner liner, comprising:
- combusting a fuel in a combustor of the gas turbine engine; and
- directing cooling air through a plurality of nozzles extending radially inwards through the outer liner to impinge upon and cool the inner liner, such that the cooling air exits the plurality of nozzles closer to the inner liner at the downstream end than at the upstream end, wherein the cooling air directed through at least one nozzle of the plurality of nozzles exit the at least nozzle through multiple air flow paths symmetrically arranged about a longitudinal axis of the at least one nozzle.
11. The method of claim 10, wherein directing the cooling air includes directing the cooling air through the multiple air flow paths of the at least one nozzle such that the cooling air diverges.
12. The method of claim 10, wherein directing the cooling air includes directing the cooling air though the multiple air flow paths of the at least nozzle such that the cooling air through each of the multiple air flow paths flow substantially parallel to one another.
13. A gas turbine engine, comprising:
- an impingement cooled double-walled liner, including an inner liner and an outer liner, disposed around a combustion space of the turbine engine and extending from an upstream end to a downstream end; and
- a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner, each nozzle of the plurality of nozzles extending radially inwards from a first distal end to a second proximal end, wherein each nozzle of the plurality of nozzles include multiple air holes arranged in a shower head pattern at the second end.
14. The gas turbine engine of claim 13, wherein the plurality of nozzles are arranged such that a radial gap of the second end of a nozzle to the inner liner decreases as a function of distance from the upstream end to the downstream end.
15. The gas turbine engine of claim 13, wherein the multiple air holes are symmetrically positioned about a longitudinal axis of each nozzle.
16. The gas turbine engine of claim 15, wherein each air hole of the multiple air holes are inclined with respect to the longitudinal axis such that the cooling air exiting each nozzle diverges.
17. The gas turbine engine of claim 16, wherein an inclination of each air hole of the multiple air holes with respect to the longitudinal axis is substantially the same.
18. The gas turbine engine of claim 13, wherein the second end of each nozzle includes a projection that extends towards the inner liner.
19. The gas turbine engine of claim 18, wherein the multiple air holes are symmetrically arranged about the projection.
20. The gas turbine engine of claim 13, wherein the second end of each nozzle is curved such that a central portion of the second end forms a proximal-most portion of the nozzle.
Type: Application
Filed: Sep 30, 2011
Publication Date: Apr 4, 2013
Applicant:
Inventor: Yong Weon Kim (San Diego, CA)
Application Number: 13/250,274
International Classification: F23R 3/42 (20060101);