U.S. patent application number 13/250274 was filed with the patent office on 2013-04-04 for impingement cooling of combustor liners.
This patent application is currently assigned to Solar Turbines Incorporated. The applicant listed for this patent is Yong Weon Kim. Invention is credited to Yong Weon Kim.
Application Number | 20130081401 13/250274 |
Document ID | / |
Family ID | 47991348 |
Filed Date | 2013-04-04 |
United States Patent
Application |
20130081401 |
Kind Code |
A1 |
Kim; Yong Weon |
April 4, 2013 |
IMPINGEMENT COOLING OF COMBUSTOR LINERS
Abstract
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.
Inventors: |
Kim; Yong Weon; (San Diego,
CA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Kim; Yong Weon |
San Diego |
CA |
US |
|
|
Assignee: |
Solar Turbines Incorporated
|
Family ID: |
47991348 |
Appl. No.: |
13/250274 |
Filed: |
September 30, 2011 |
Current U.S.
Class: |
60/772 ;
60/752 |
Current CPC
Class: |
F23R 3/50 20130101; F23R
3/06 20130101; F23R 2900/03044 20130101 |
Class at
Publication: |
60/772 ;
60/752 |
International
Class: |
F23R 3/42 20060101
F23R003/42 |
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.
Description
TECHNICAL FIELD
[0001] The present disclosure relates generally to systems and
methods of impingement cooling a combustor liner of a gas turbine
engine.
BACKGROUND
[0002] Combustor 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.
[0003] 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.
SUMMARY
[0004] In 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.
[0005] 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.
[0006] 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.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] FIG. 1 is an illustration of an exemplary disclosed gas
turbine engine;
[0008] FIG. 2 is a cut-away illustration of an exemplary combustor
system of the gas turbine engine of FIG. 1;
[0009] FIG. 3 is a cross-sectional view of an embodiment of the
outer combustor wall of the gas turbine engine of FIG. 1;
[0010] FIG. 4A is a cross-sectional view of another embodiment of
the outer combustor wall of the gas turbine engine of FIG. 1;
[0011] FIG. 4B is a cross-sectional view of another embodiment of
the outer combustor wall of the gas turbine engine of FIG. 1;
[0012] FIG. 5A is a cross-sectional view of an embodiment of a
nozzle of the gas turbine engine of FIG. 1;
[0013] FIG. 5B is another cross-sectional view of the exemplary
nozzle of FIG. 5A;
[0014] FIG. 6 is a cross-sectional view of another embodiment of a
nozzle of the gas turbine engine of FIG. 1; and
[0015] FIG. 7 is a cross-sectional view of another embodiment of a
nozzle of the gas turbine engine of FIG. 1.
DETAILED DESCRIPTION
[0016] FIG. 1 illustrates an exemplary gas turbine engine (GTE) 100
having a compressor system 10, a combustor system 20, a turbine
system 70, and an exhaust system 90 arranged lengthwise along an
engine axis 98. The compressor system 10 may compress air and
deliver the compressed air to an enclosure 72 of the combustor
system 20. The compressed air may be mixed with a fuel and directed
into a combustor 50 through one or more fuel injectors 30. The
fuel-air mixture may ignite and burn in the combustor 50 to produce
combustion gases that may be directed to the turbine system 70. The
turbine system 70 may extract energy from these combustion gases,
and direct the exhaust gases to the atmosphere through the exhaust
system 90.
[0017] FIG. 2 is a cut-away view of combustor system 20 showing the
combustor 50. Combustor 50 includes an outer combustor wall 22 and
an inner combustor wall 24 annularly disposed about the engine axis
98. The outer and the inner combustor walls (22, 24) are joined
together at an upstream end by a dome assembly to define a
combustion space 58 therebetween.
[0018] 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 FIG. 2 is
exemplary only, and that several variations are possible.
[0019] FIG. 3 is a cross-sectional schematic of the outer combustor
wall 22 illustrating the impingement cooling of the inner liner
22b. A high pressure stream of air ("air jets 36") enters the
cooling space 26 through perforations 32 on the outer liner 22a.
These air jets 36 impinge on, and cool, the inner liner 22b. After
impingement, the spent air stream flows towards the downstream end
46 to form the cross-flow air 38 that may be mixed with the
combustion gases or discarded. It is known that cross-flow air 38
from the upstream end 44, interacts with, and degrades the ability
of the air jets 36 at the downstream end 46 to impinge on, and
cool, the inner liner 22b. For instance, the cross-flow air 38 from
a first perforation 32a may degrade the ability of the air jet 36
from a second perforation 32b, downstream of the first perforation
32a, to impinge on the region of the inner liner 22b under the
second perforation 32b. Similarly, the cross-flow air 38 from the
first and second perforations 32a, 32b may collectively further
degrade the cooling ability of an air jet 36 from a third
perforation 32c, further downstream of the first perforation 32a,
to cool the inner liner 22b under the third perforation 32c. In
some embodiments, some (or all) of the perforations 32 may include
extended ports or nozzles 48 (see FIGS. 4A-7) to reduce the impact
of the cross-flow air 38 from an upstream perforation 32 on a
downstream air jet 36.
[0020] FIG. 4A illustrates a cross-sectional view of the outer
combustor wall 22 illustrating nozzles 48 attached to the
perforations 32. The nozzles 48 may include air holes 62 that
extend from a first end 66, positioned in the enclosure 72 outside
the outer liner 22a, to a second end 68 positioned in the cooling
space 26 inside the outer liner 22a. These air holes 62 may direct
the compressed air in the enclosure 72 (air jets 36 of FIG. 3) to
impinge on the inner liner 22b. The nozzles 48 may be a separate
part attached to the outer liner 22a (by any conventional
attachment process, such as brazing, etc.) or may be a region of
the outer liner 22a that is bent towards the inner liner 22b (such
as for example, the rim of a perforation that is folded towards the
inner liner 22b). The nozzles 48 may be arranged on the outer liner
22a such that a radial gap (t) between the second end 68 of a
nozzle 48 and the inner liner 22b decreases from the upstream end
44 to the downstream end 46 (that is,
t.sub.a>t.sub.b>t.sub.c>t.sub.d>t.sub.e). To achieve
this decreasing radial gap (t) from the upstream end 44 to the
downstream end 46, in some embodiments (as illustrated in FIG. 4A),
the length of the nozzles 48 may progressively increase from the
upstream end 44 to the downstream end 46. Because the air jets 36
enter the cooling space 26 closer to the inner liner 22b at the
downstream end 46, the effect of the cross-flow air 38 from the
upstream air jets 36 on the downstream air jets 36 will be lower.
In some embodiments, substantially all the rows of perforations on
the outer liner 22a will include nozzles 48, while in other
embodiments, only selected rows of perforations along the length of
the outer liner 22a will include nozzles 48. In some embodiments,
the perforations 34 on the inner combustor wall 24 (see FIG. 2)
will also include nozzles 48 so that the air jets 36 enter the
cooling space 28 closer to the inner liner 24b at the downstream
end 46 than at the upstream end 44.
[0021] Although FIG. 4A illustrates the radial gap (t) between the
nozzles 48 and the inner liner 22b as decreasing substantially
linearly from the upstream end 44 to the downstream end 46, this is
only exemplary. In general, the radial gap (t) may vary in any
manner (such as, for example, decrease exponentially from the
upstream end to the downstream end). In some embodiments, although
the radial gap (t) may generally decrease from the upstream end 44
to the downstream end 46, the radial gaps (t) of selected adjacent
nozzles 48 may be substantially the same (such as, for example,
t.sub.a.apprxeq.t.sub.b>t.sub.c>t.sub.d.apprxeq.t.sub.e).
[0022] In some embodiments, as illustrated in FIG. 4B, only
perforations 32 in selected regions of the outer liner 22a may
include nozzles 48 to direct the air jets 36 in these regions
closer to the inner liner 22b. For example, in some embodiments,
nozzles 48 may only be included in a few rows of perforations 32 at
the downstream end 46 in applications where only the air jets 36
from those few rows are detrimentally affected by the cross-flow
air 38 from the upstream end 44. In some embodiments (as
illustrated in FIG. 4B), the radial gap (t) between these nozzles
48 and the inner liner 22b may be substantially the same (that is,
t.sub.a.apprxeq.t.sub.b). Although FIGS. 4A and 4B illustrate
embodiments, where nozzles 48 are used to decrease the radial gap
(t) from the upstream end 44 to the downstream end 46, it is
contemplated that in some embodiments, the radial gap (t) may
instead be decreased by decreasing the distance between the inner
liner 22b and the outer liner 22a (that is, the thickness of the
cooling space 26) from the upstream end 44 to the downstream end
46.
[0023] 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.
[0024] FIGS. 5A and 5B are cross-sectional illustrations of an
exemplary embodiment of a nozzle 48A of the current disclosure.
FIG. 5A illustrates a cross-sectional view along a plane parallel
to a longitudinal axis 88 of nozzle 48A, and FIG. 5B illustrates a
cross-sectional view along a plane transverse to the longitudinal
axis 88. In the discussion that follows, reference will be made to
both FIGS. 5A and 5B. One or more air holes 62 may direct
compressed air out of nozzle 48A at second end 68. In some
embodiments, as illustrated in FIG. 5A, the one or more air holes
62 may form a shower head pattern of air holes 62 at the second end
68. In some embodiments, all the air holes 62 may extend from the
first end 66 to the second end 68, while in other embodiments (as
illustrated in FIG. 5A), a single air hole 62 that extends from the
first end 66 may be divided into multiple air holes 62 to form a
shower head pattern at the second end 68. The single air hole may
be divided into multiple air holes anywhere along the length of
nozzle 48A. In some embodiments, as illustrated in FIG. 5A, the
single air hole may be divided into multiple air holes proximate
the second end 68. In some embodiments, multiple (for example, 2,
3, 4, 5, 6 etc.) air holes 62 may be positioned symmetrically
around the longitudinal axis 88 at the second end 68. If the inner
liner 22b, 24b bulges during operation, the bulging liner may
contact a central portion (proximate longitudinal axis 88) of the
second end 68 of a nozzle 48. And, since the air holes 62 are
distributed around the central portion, some or all of the air
holes 62 may remain unblocked by the bulging inner liner 22b, 24b.
Even if flow through some of the air holes 62 is restricted (or
even blocked) by the contacting inner liner 22b, 24b, the flow
though the remaining air holes 62 may provide sufficient cooling
for the inner liner 22b, 24b. Thus, a shower head pattern of air
holes 62 in nozzle 48A may allow air flow to continue through at
least some of the air holes 62 when there is contact between the
nozzle 48A and the inner liner 22b, 24b.
[0025] 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 FIG. 5A, the projections 74
may have a rounded edge to reduce bearing stresses on the inner
liner 22b, 24b during contact. Although described as a projection,
it is contemplated that other features (such as grooves, cut-outs,
etc.) that allow air from the air holes 62 to exit out of the
nozzle 48A when there is contact between the nozzle 48A and the
inner liner 22b, 24b, may be provided. In place of, or in addition
to, discrete projections 74 on the second end 68, in some
embodiments, the shape of a nozzle 48A may include a projecting
region on the second end 68. For example, as illustrated in FIG. 6,
the second end 68 of an exemplary nozzle 48B may have a curved
shape with a projecting central region. In these embodiments, the
projecting central region may act as the projection 74 that
contacts a bulging inner liner 22b, 24b. In addition to this
projecting central region, in some embodiments, additional
projections 74 may also be provided on second end 68 of nozzle
48B.
[0026] In some embodiments (as illustrated in nozzles 48A and 48B
of FIGS. 5A-6), central axes (64a, 64b, 64c, 64d, etc.) of the
multiple air holes 62 may be substantially parallel to the
longitudinal axis 88. However, in other embodiments, the central
axis of an air hole 62 may be inclined with respect to the
longitudinal axis 88 of the nozzle. FIG. 7 illustrates an
embodiment of a nozzle 48C in with the central axes 64a, 64b of the
air holes 62a and 62b make angles .theta..sub.a and .theta..sub.b,
respectively, with respect to the longitudinal axis 88. The angles
.theta..sub.a and .theta..sub.b may have the same or different
magnitudes. In these embodiments, a bulging inner liner 22b, 24b
may contact the central portion of the second end 68 and allow air
flow through the inclined air holes 62 even in the absence of a
projection at the second end 68. In some embodiments, one or more
projections 74 may also be provided at the second end 68 of nozzle
48C to act as a stand-off. The inclined air holes 62 may also allow
the air flowing through them to diverge and impinge on a larger
area of the inner liner 22b, 24b.
[0027] 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
FIGS. 4A and 4B) may be used in areas where the possibility of
contact with the inner liner 22b, 24b is minimal, and a nozzle
having multiple inclined air holes 62 (such as nozzle 48C of FIG.
7) may be used where the possibility of contact with the inner
liner 22b, 24b exists. It should be noted that, although contact
between a nozzle 48 and the inner liner 22b, 24b is described as
being a result of a pressure wave in the combustor 50 that causes a
portion of the inner liner 22b, 24b to bulge and contact one or
more nozzles 48 on the outer liner 22a, 24b, this is only
exemplary. In some applications, vibration of the combustor 50 may
cause contact between the inner liner 22b, 24b and the nozzles 48.
Contact between a nozzle 48 and the inner liner 22b, 24b can occur
for various other reasons, and the disclosed system can be used to
provide continuous air flow through the nozzles 48 during contact
that occurs for any reason.
INDUSTRIAL APPLICABILITY
[0028] 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.
[0029] With reference to FIGS. 1 and 2, during operation of GTE
100, air may be drawn into compressor section 10 and compressed.
This compressed air may then be directed to enclosure 72 around the
combustor 50. The combustor may enclose a combustion space 58
bounded by a double-walled liner (including inner liners 22b, 24b
and outer liners 22a, 24a). A portion of the compressed air may be
mixed with fuel and combusted in the combustion space 58. The
combustion heats the inner liners 22b, 24b of the combustor 50. A
portion of the compressed air in the enclosure 72 is directed
though the perforations 32, 34 on the outer liner 22a, 24a to
impinge on, and cool, the hot inner liner 22b, 24b (FIGS. 3). To
reduce the impact of cross-flow air 38, from upstream perforations,
on the cooling effectiveness of downstream perforations, nozzles 48
are provided on some or all the perforations 32, 34. These nozzles
48 deliver the impingement air jets closer to the inner liner 22b,
24b at the downstream end 46 of the combustor 50 and thereby reduce
the effect of the cross-flow air on the cooling effectiveness of
the downstream air jets. To reduce the possibility of the air jets
being blocked by dimensional variations of the liner walls during
operation of the turbine engine (such as bulging of the inner liner
22b, 24b), the air jets may be provided in a shower head pattern at
the tip of the nozzles 48. A shower head pattern of air jets may
allow some of the air jets to continue to impinge on, and cool, the
inner liner 22b, 24b even when a bulging inner liner contacts and
blocks some of the air jets.
[0030] 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.
* * * * *