U.S. patent application number 12/131403 was filed with the patent office on 2008-09-25 for turbine shroud segment transpiration cooling with individual cast inlet and outlet cavities.
This patent application is currently assigned to PRATT & WHITNEY CANADA CORP.. Invention is credited to Eric DUROCHER, Assaf FARAH.
Application Number | 20080232963 12/131403 |
Document ID | / |
Family ID | 36917246 |
Filed Date | 2008-09-25 |
United States Patent
Application |
20080232963 |
Kind Code |
A1 |
DUROCHER; Eric ; et
al. |
September 25, 2008 |
TURBINE SHROUD SEGMENT TRANSPIRATION COOLING WITH INDIVIDUAL CAST
INLET AND OUTLET CAVITIES
Abstract
A shroud segment of a turbine shroud of a gas turbine engine
comprises a platform with front and rear legs. The platform defines
a plurality of axially extending holes with individual inlets on an
outer surface of the platform for transpiration cooling of the
platform of the turbine shroud segment.
Inventors: |
DUROCHER; Eric; (Vercheres,
CA) ; FARAH; Assaf; (Charlemagne, CA) |
Correspondence
Address: |
OGILVY RENAULT LLP (PWC)
1981 MCGILL COLLEGE AVENUE, SUITE 1600
MONTREAL
QC
H3A 2Y3
CA
|
Assignee: |
PRATT & WHITNEY CANADA
CORP.
Longueuil
CA
|
Family ID: |
36917246 |
Appl. No.: |
12/131403 |
Filed: |
June 2, 2008 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
11183741 |
Jul 19, 2005 |
|
|
|
12131403 |
|
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Current U.S.
Class: |
416/1 ;
416/97R |
Current CPC
Class: |
F05D 2250/51 20130101;
F05D 2230/21 20130101; F05D 2250/52 20130101; F05D 2260/20
20130101; F01D 25/12 20130101; F05D 2240/11 20130101; F01D 11/08
20130101; F01D 9/04 20130101 |
Class at
Publication: |
416/1 ;
416/97.R |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Claims
1-14. (canceled)
15. A gas turbine engine cooling arrangement for serially cooling a
turbine static shroud and a turbine vane outer shroud, the cooling
arrangement comprising a plurality of passages extending through a
turbine static shroud platform substantially in a direction of a
turbine rotating axis for directing cooling air through the turbine
static shroud platform to cool the turbine shroud, the passages
having respective outlets positioned and configured to direct air
exhausted from the passages to impinge substantially all of a
circumferential leading end of the turbine vane outer shroud
16. The cooling arrangement as claimed in claim 15 wherein each of
the passages extends through a major section of, an entire axial
length of the shroud platform.
17. The cooling arrangement as claimed in claim 16 wherein each of
the passages is in fluid communication with a cavity defined
between front and rear legs of the turbine shroud segment.
18. The cooling arrangement as claimed in claim 17 wherein each of
the passages discharges the cooling air at a trailing end of the
shroud platform for impingement, on the turbine vane outer
shroud.
19. The cooling arrangement as claimed in claim 18 wherein the
passages comprise a plurality of substantially axial straight
holes.
20. A gas turbine engine assembly defining a portion of an outer
wall of a gas path, the assembly comprising a turbine static shroud
having a platform, and a turbine vane outer shroud disposed
immediately downstream of the turbine static shroud, the turbine
vane outer shroud having a plurality of vanes extending radially
inwardly from the turbine vane outer shroud, the turbine static
shroud defining a plurality of passages extending substantially
axially in a direction of a turbine rotating axis through the
platform of the turbine static shroud, the passages having outlets
on a trailing end of the platform which substantially radially
align with a leading edge of the turbine vane outer shroud, the
outlets configured for discharging air exhausted from the passages
to impinge on the turbine vane outer shroud leading edge.
21. The gas turbine engine structure as claimed in claim 20 wherein
the passages comprise a plurality of holes having at least one
inlet thereof on an outer surface of the platform.
22. The gas turbine engine structure as claimed in claim 21 wherein
the at least one inlet of the holes is located in a position close
to and downstream a front leg of the shroud.
23. The gas turbine engine structure as claimed in claim 21 wherein
the holes extend in a substantially straight direction over a major
section of the entire axial length of the platform.
24. A method of reusing turbine shroud cooling air for impingement
cooling on a downstream turbine vane outer shroud, the method
comprising steps of: (a) directing cooling air within and through a
platform of a shroud segment of a turbine shroud for cooling the
turbine shroud; and (b) using the cooling air of step (a) to form a
plurality of substantially straight cooling air streams axially
towards a leading end of the turbine vane outer shroud for
impingement cooling on the turbine vane outer shroud.
25. The method as claimed in claim 24 wherein steps (a) and (b) are
conducted substantially simultaneously.
26. The method as claimed in claim 25 wherein steps (a) and (b) are
practiced by directing the cooling air through a plurality of
substantially axial and straight passages extending within the
platform of the shroud segment of the turbine shroud to deliver the
substantially straight cooling, air streams with a high velocity
thereof.
27. The method as claimed in claim 26 wherein the cooling air is
directed from a cavity defined between front and rear legs of the
shroud segment, into the substantially axial and straight passages,
and is discharged therefrom at a trailing end of the platform.
28. The method as claimed in claim 24 further comprising a step (c)
of discharging the cooling air into a gas path upon the impingement
cooling thereof on the turbine vane outer shroud.
29. The method as claimed in claim 24 comprising directing the
substantially straight cooling air streams in a manner for
impingement cooling on a substantially entire circumference of the
leading end of the turbine vane outer shroud.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This is a Continuation of Applicant's U.S. patent
application Ser. No. 11/183,741, filed on Jul. 19, 2005.
TECHNICAL FIELD
[0002] The invention relates generally to gas turbine engines and
more particularly to turbine shroud segments configured for
transpiration cooling of a turbine shroud assembly.
BACKGROUND OF THE ART
[0003] A gas turbine engine usually includes a hot section, i.e., a
turbine section which includes at least one rotor stage, for
example, having a plurality of shroud segments disposed
circumferentially one adjacent to another to form a shroud ring
surrounding a turbine rotor, and at least one stator vane stage
disposed immediately downstream and/or upstream of the rotor stage,
formed with outer and inner shrouds and a plurality of radial
stator vanes extending therebetween. Being exposed to very hot
gases, the rotor stage and the stator vane stage need to be cooled.
Hereintofore, efforts have been made in various approaches for
development of adequate cooling arrangements. Therefore, gas
turbine engine designers have been continuously seeking improved
configurations of turbine shroud segments which are not only
adapted for adequate cooling arrangement of a turbine shroud
assembly but also provide improved mechanical properties thereof,
as well as convenience of manufacture.
[0004] Accordingly, there is a need to provide improved turbine
shroud segments adapted for adequate cooling arrangement of a
turbine shroud assembly.
SUMMARY OF THE INVENTION
[0005] It is therefore an object of this invention to provide
turbine shroud segments adapted for adequate cooling arrangement of
the turbine shroud assembly.
[0006] One aspect of the present invention therefore provides a
turbine shroud segment of a turbine shroud of a gas turbine engine,
which comprises a platform having a hot gas path side and a back
side. The platform is axially defined between leading and trailing
ends thereof and is circumferentially defined between opposite
lateral sides thereof. The platform further defines a plurality of
axially extending transpiration holes with individual inlets on the
back side of the platform for transpiration cooling of the platform
of the turbine shroud segment.
[0007] Another aspect of the present invention provides a turbine
shroud of a gas turbine engine which comprises a plurality of
circumferentially adjoining shroud segments and an annular support
structure supporting the shroud segments together within an engine
casing. Each of the shroud segments includes a platform and also
includes front and rear legs to support the platform radially and
inwardly spaced apart from the support structure in order to define
an annular cavity between the front and rear legs. The platform
defines a plurality of transpiration cooling passages extending
therein and substantially axially therethrough. The transpiration
cooling passages have individual inlets defined in the outer
surface of the platform in fluid communication with the annular
cavity for intake of cooling air therefrom.
[0008] These and other aspects of the present invention will be
better understood with reference to preferred embodiments described
hereinafter.
DESCRIPTION OF THE DRAWINGS
[0009] Reference is now made to the accompanying figures depicting
aspects of the present invention, in which:
[0010] FIG. 1 is a schematic cross-sectional view of a gas turbine
engine;
[0011] FIG. 2 is an axial cross-sectional view of a turbine shroud
assembly used in the gas turbine engine of FIG. 1, in accordance
with one embodiment of the present invention;
[0012] FIG. 3 is a perspective view of a shroud segment used in the
turbine shroud assembly of FIG. 2; and
[0013] FIG. 4 is a perspective view of a shroud segment alternative
to the shroud segment of FIG. 3, according to another embodiment of
the present invention.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
[0014] Referring to FIG. 1, a turbofan gas turbine engine
incorporates an embodiment of the present invention, presented as
an example of the application of the present invention, and
includes a housing or a nacelle 10, a core casing 13, a low
pressure spool assembly seen generally at 12 which includes a fan
14, low pressure compressor 16 and low pressure turbine 18, and a
high pressure spool assembly seen generally at 20 which includes a
high pressure compressor 22 and a high pressure turbine 24. There
is provided a burner 25 for generating combustion gases. The low
pressure turbine 18 and high pressure turbine 24 include a
plurality of rotor stages 28 and stator vane stages 30.
[0015] Referring to FIGS. 1-3, each of the rotor stages 28 has a
plurality of rotor blades 33 encircled by a turbine shroud assembly
32 and each of the stator vane stages 30 includes a stator vane
assembly 34 which is positioned upstream and/or downstream of one
of the rotor stages 28, for directing combustion gases into or out
of an annular gas path 36 within a corresponding turbine shroud
assembly 32, and through the corresponding rotor stage 28.
[0016] The stator vane assembly 34, for example a first stage of a
low pressure turbine (LPT) vane assembly, is disposed, for example,
downstream of the shroud assembly 32 of one rotor stage 28, and
includes, for example a plurality of stator vane segments (not
indicated) joined one to another in a circumferential direction to
form a turbine vane outer shroud 38 which comprises a plurality of
axial stator vanes 40 (only a portion of one is shown) which divide
a downstream section of the annular gas path 36 relative to the
rotor stage 28, into sectoral gas passages for directing combustion
gas flow out of the rotor stage 28.
[0017] The shroud assembly 32 in the rotor stage 28 includes a
plurality of shroud segments 42 (only one shown) each of which
includes a platform 44 having front and rear radial legs 46, 48
with respective hooks. The shroud segments 42 are joined one to
another in a circumferential direction and thereby form the shroud
assembly 32.
[0018] The platform 44 of each shroud segment 42 has a back side 50
and a hot gas path side 52 and is defined axially between leading
and trailing ends 54, 56, and circumferentially between opposite
lateral sides 58, 60 thereof. The platforms 44 of the segments
collectively form a turbine shroud ring (not indicated) which
encircles the rotor blades 33 and in combination with the rotor
stage 28, defines a section of the annular gas path 36. The turbine
shroud ring is disposed immediately upstream of and abuts the
turbine vane outer shroud 38, to thereby form a portion of an outer
wall (not indicated) of the annular gas path 36.
[0019] The front and rear radial legs 46, 48 are axially spaced
apart and integrally extend from the back side 50 radially and
outwardly such that the hooks of the front and rear radial legs 46,
48 are conventionally connected with an annular shroud support
structure 62 which is formed with a plurality of shroud support
segments (not indicated) and is in turn supported within the core
casing 13. An annular cavity 64 is thus defined axially between the
front and rear legs 46, 48 and radially between the platforms 44 of
the shroud segments 42 and the annular shroud support structure 62.
The annular middle cavity is in fluid communication with a cooling
air source, for example bleed air from the low or high pressure
compressors 16, 22 and thus the cooling air under pressure is
introduced into and accommodated within the annular cavity 64.
[0020] The platform 44 of each shroud segment 42 preferably
includes a passage, for example a plurality of transpiration holes
66 extending axially within the platform 44 for directing cooling
air therethrough for transpiration cooling of the platform 44. In
prior art, for convenience of the hole drilling, a groove (not
shown) extending in a circumferential direction with opposite ends
closed is conventionally provided, for example, on the back side 50
of the platform 44 such that transpiration holes 66 can be drilled
from the trailing end 56 of the platform straightly and axially
towards and terminate at the groove. Thus, such a groove forms a
common inlet of the transpiration holes 66 for intake of cooling
air accommodated within the cavity 64. However, this type of groove
usually extends across almost the entire width of the platform 44
and has a depth of about a half the thickness of the platform 44.
Therefore, the groove unavoidably and significantly reduces the
strength of the platform 44 and thus the durability of shroud
segment 42.
[0021] In accordance with one embodiment of the present invention,
a plurality of individual inlets, preferably cast inlet cavities
68, instead of a conventional groove, are provided on the back side
50 of the platform 44, in order to overcome the shortcomings of the
prior art while providing convenience of manufacture for the
hole-making in the platform 44. The transpiration holes 66 can be
drilled from the trailing end 56 of the platform 44 axially towards
and terminate at the individual cast inlet cavities 68. The number
of cast inlet cavities 68 is equal to the number of the
transpiration holes 66. The dimension of the individual cast inlet
cavities 68 is preferably greater than the diameter of the
respective transpiration holes 66. For example, the individual cast
inlet cavities 68 may be shaped with a bell mouth profile which
provides convenience for the casting process of the platforms 44.
In contrast to the conventional groove as a common inlet of the
transpiration holes 66, the body portions of the platform 44
remaining between the adjacent cast inlet cavities 66, effectively
improve the strength of the platform 44 and thus the durability of
the shroud segment 42.
[0022] The individual cast inlet cavities 68 are in fluid
communication with the middle cavity 64 and thus cooling air
introduced into the cavity 64 is directed into and through the
axial transpiration holes 66 for effectively cooling the platform
44 of the shroud segments 42. The cooling air is then discharged at
the trailing end 56 of the platform 42, impinging on a downstream
engine part such as the turbine vane outer shroud 38, before
entering the gas path 36.
[0023] The individual cast inlet cavities 68 are preferably located
close to the front leg 46 such that the transpiration holes 66
extend through a major section of the entire axial length of the
platform 44 of the shroud segment 42, thereby efficiently cooling
the platform 44 of the shroud segment 42.
[0024] The transpiration holes 66 are preferably substantially
evenly spaced apart in a circumferential direction and are
preferably aligned with the turbine vane outer shroud. Thus, the
cooling air impinges on the leading end of the turbine vane outer
shroud 38. The number of transpiration holes 66 in each shroud
segment 42 is determined such that the cooling air discharged from
the transpiration holes 66 effectively cools the entire
circumference of the leading end of the turbine vane outer shroud
38.
[0025] The trailing end 56 of the platform 44 is conventionally
disposed in a very close or abutting relationship with the leading
end (not indicated) of the turbine vane outer shroud 38, in order
to prevent leakage of hot combustion gases flowing through the gas
path 36. It is therefore preferable to provide one or more outlets
in the trailing end 56 of the platform 44 for adequately
discharging cooling air from the transpiration holes 66, thereby
not only permitting the cooling air to flow through the
transpiration holes 66 without substantial blocking but also
directing the discharged cooling air to adequately cool the stator
vane assembly 34.
[0026] In this embodiment a plurality of individual outlets,
preferably individual cast outlet cavities 70, are provided in the
trailing end 56 of the platform 44 of each shroud segment 42. For
example, each cast outlet cavity 70 is configured as a groove
extending radially in the trailing end 56 of the platform 44, with
opposite ends: one end being closed and the other end opening onto
hot gas path side 52 of the platform 44. The transpiration holes 66
are in fluid communication with and terminate at the individual
grooves (the individual cast outlet cavities 70). Due to the
restriction by the closed end of the radial grooves, the cooling
air discharged from the transpiration holes 66 is directed to
impinge the leading end of the turbine vane outer shroud 38, and
upon impingement thereon is directed radially, inwardly and
rearwardly, thereby further film cooling a front portion of the
inner surface of the turbine vane outer shroud 38 and a portion of
the axial stator vanes 40, prior to being discharged into hot
combustion gases flowing through the gas path 36. In contrast to
the cross-section of the transpiration holes 66, the individual
cast outlet cavities 70 have an enlarged dimension which
advantageously reduces the contact surface of the trailing end 56
of the platform 44 with the leading end of the turbine vane outer
shroud 38, thereby minimizing fretting therebetween.
[0027] FIG. 4 illustrates another embodiment of the shroud segment
42 which is similar and alternative to the embodiment of FIG. 3 and
will not be redundantly described. The only difference therebetween
lies in that the individual cast outlet cavities 70 of FIG. 3 are
replaced by an elongate, preferably cast, recess 70 which is a
common outlet of the holes 66 and is provided in the trailing end
56 of the platform 44 with an opening defined on the hot gas path
side 52 of the platform 44. The elongate recess 70 will provide a
function generally similar to that of the individual outlets.
However, individual outlets are preferable to a common outlet
because cooling air streams discharged from the transpiration holes
66 through the individual outlets 70 will not interfere with one
another when approaching the leading end of the turbine vane outer
shroud 38 for impingement cooling thereof.
[0028] The above description is meant to be exemplary only, and one
skilled in the art will recognize that changes may be made to the
embodiments described without departure from the scope of the
invention disclosed. For example, the present invention can be
applicable in any type of gas turbine engine other than the
described turbofan gas turbine engine. The described individual
inlet and outlet cavities may be used either in combination or in a
separate manner in various configurations of turbine shroud
segments. Other modifications which fall within the scope of the
present invention will be apparent to those skilled in the art, in
light of a review of this disclosure, and such modifications are
intended to fall within the appended claims.
* * * * *