U.S. patent application number 13/668398 was filed with the patent office on 2014-05-08 for blade outer air seal.
This patent application is currently assigned to UNITED TECHNOLOGIES CORPORATION. The applicant listed for this patent is UNITED TECHNOLOGIES CORPORATION. Invention is credited to Brian Duguay, Russell E. Keene, Dmitriy A. Romanov.
Application Number | 20140127006 13/668398 |
Document ID | / |
Family ID | 50622522 |
Filed Date | 2014-05-08 |
United States Patent
Application |
20140127006 |
Kind Code |
A1 |
Romanov; Dmitriy A. ; et
al. |
May 8, 2014 |
BLADE OUTER AIR SEAL
Abstract
A blade outer air seal for a gas turbine engine includes a first
side surface, a second side surface, and a wall. The wall extends
between the first side surface and the second side surface and has
one or more holes formed therein. The holes are spaced from the
first side surface and/or the second side surface and have areas
between about 0.005% and 0.450% of a surface area of the blade
outer air seal.
Inventors: |
Romanov; Dmitriy A.; (Wells,
ME) ; Duguay; Brian; (Berwick, ME) ; Keene;
Russell E.; (Arundel, ME) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
UNITED TECHNOLOGIES CORPORATION |
Hartford |
CT |
US |
|
|
Assignee: |
UNITED TECHNOLOGIES
CORPORATION
Hartford
CT
|
Family ID: |
50622522 |
Appl. No.: |
13/668398 |
Filed: |
November 5, 2012 |
Current U.S.
Class: |
415/173.1 |
Current CPC
Class: |
F01D 9/04 20130101; F05D
2260/201 20130101; F05D 2240/11 20130101; Y02T 50/60 20130101; F01D
25/246 20130101; Y02T 50/6765 20180501; F01D 11/08 20130101; F05D
2260/221 20130101; Y02T 50/676 20130101 |
Class at
Publication: |
415/173.1 |
International
Class: |
F01D 11/08 20060101
F01D011/08 |
Claims
1. A blade outer air seal for a gas turbine engine, the blade outer
air seal comprising: a first side surface; a second side surface;
and a wall extending between the first side surface and the second
side surface, wherein one or more holes are formed in the wall and
are spaced from at least one of the first side surface or second
side surface, and wherein the one or more holes have areas between
about 0.005% and 0.450% of a surface area of the blade outer air
seal.
2. The blade outer air seal of claim 1, wherein the one or more
holes have areas between about 0.02% and 0.30% of a surface area of
the blade outer air seal
3. The blade outer air seal of claim 1, wherein the one or more
holes comprise six holes with three holes positioned adjacent the
first side surface and three holes positioned adjacent the second
side surface.
4. The blade outer air seal of claim 1, further comprising internal
passages that extend through the wall from the first side surface
to the second side surface, and wherein the one or more holes
communicate with internal passages.
5. The blade outer air seal of claim 4, wherein the one or more
holes comprise six holes with one hole for communicating with each
one of the internal passages
6. The blade outer air seal of claim 4, wherein each of the
internal passages has a radial height between 30% to 40% of a total
radial thickness the wall.
7. The blade outer air seal of claim 5, wherein the internal
passages together have an axial length that comprises between 75%
and 85% of the axial length of the wall.
8. The blade outer air seal of claim 1, further comprising one or
more forward hooks that extend from the wall, wherein at least one
of the forward hooks has a slot therein that is offset relative to
an axis of symmetry of the blade outer air seal.
9. The blade outer air seal of claim 8, wherein at least one of the
forward hooks has an angled wall that extends from an outer radial
surface of the at least one of the forward hooks to the wall.
10. The blade outer air seal of claim 1, wherein the wall has a
thermal barrier coating applied to an inner radial surface thereof,
wherein the thermal barrier coating has a radial thickness between
3% and 10% of a total radial thickness of the wall, wherein the
wall includes a bond coat, and wherein the bond coat has a radial
thickness between 3% and 10% of a total radial thickness of the
wall.
11. A blade outer air seal for a gas turbine engine, the blade
outer air seal comprising: a first side surface; a second side
surface; and a wall extending between the first side surface and
the second side surface, wherein the blade outer air seal has a
wall with a bond coat and a thermal barrier coating, and wherein
both the bond coat and the thermal barrier coating have a radial
thickness between 3% and 10% of the total radial thickness of the
wall.
12. The blade outer air seal of claim 11, one or more holes are
formed in the wall and are spaced from at least one of the first
side surface or second side surface, and wherein the one or more
holes have areas between about 0.005% and 0.450% of a surface area
of the blade outer air seal.
13. The blade outer air seal of claim 12, wherein the one or more
holes have areas between about 0.02% and 0.30% of a surface area of
the blade outer air seal.
14. The blade outer air seal of claim 13, wherein the one or more
holes comprise six holes with three holes positioned adjacent the
first side surface and three holes positioned adjacent the second
side surface.
15. The blade outer air seal of claim 11, further comprising
internal passages that extend through the wall from the first side
surface to the second side surface, and wherein the one or more
holes communicate with internal passages.
16. The blade outer air seal of claim 15, wherein each of the
internal passages has a radial height between 30% to 40% of a total
radial thickness the wall.
17. The blade outer air seal of claim 15, wherein the internal
passages together have an axial length that comprises between 75%
and 85% of the axial length of the wall.
18. The blade outer air seal of claim 11, further comprising one or
more forward hooks that extend from the wall, wherein at least one
of the forward hooks has a slot therein that is offset relative to
an axis of symmetry of the blade outer air seal, and wherein at
least one of the forward hooks has an angled wall that extends from
an outer radial surface of the at least one of the forward hooks to
the wall.
19. A blade outer air seal for a gas turbine engine, the blade
outer air seal comprising: a first side surface; a second side
surface; a wall extending between the first side surface and the
second side surface; and one or more forward hooks extending from
the wall, wherein at least one of the forward hooks has a slot
therein that is offset relative to an axis of symmetry of the blade
outer air seal.
20. The blade outer air seal of claim 19, wherein at least one of
the forward hooks has an angled wall that extends from an outer
radial surface of the at least one of the forward hooks to the
wall.
Description
BACKGROUND
[0001] The invention relates to gas turbine engines, and more
particularly to blade outer air seals (BOAS) for gas turbine
engines.
[0002] Gas turbine engines operate according to a continuous-flow,
Brayton cycle. A compressor section pressurizes an ambient air
stream, fuel is added and the mixture is burned in a central
combustor section. The combustion products expand through a turbine
section where bladed rotors convert thermal energy from the
combustion products into mechanical energy for rotating one or more
centrally mounted shafts. The shafts, in turn, drive the forward
compressor section, thus continuing the cycle. Gas turbine engines
are compact and powerful power plants, making them suitable for
powering aircraft, heavy equipment, ships and electrical power
generators. In power generating applications, the combustion
products can also drive a separate power turbine attached to an
electrical generator.
[0003] The BOAS as well as turbine vanes are exposed to
high-temperature combustion gases and must be cooled to extend
their useful lives. Cooling air is typically taken from the flow of
compressed air. Therefore, some of the energy that could be
extracted from the flow of combustion gases must instead be
expended to provide the compressed air used to cool the BOAS as
well as the turbine vanes. Energy expended on compressing air used
for cooling the BOAS and turbine vanes is not available to produce
power. Improvements in the efficient use of compressed air for
cooling the BOAS and turbine vanes and/or materials that can better
withstand the turbine operating environment can improve the total
efficiency of the turbine engine and extend the operational life of
the BOAS.
SUMMARY
[0004] A blade outer air seal for a gas turbine engine includes a
first side surface, a second side surface, and a wall. The wall
extends between the first side surface and the second side surface
and has one or more holes formed therein. The holes are spaced from
the first side surface and/or the second side surface and have
areas between about 0.005% and 0.450% of a surface area of the
blade outer air seal.
[0005] A turbine section of a gas turbine engine includes an engine
case, a support connected to the engine case, and a blade outer air
seal. The blade outer air seal is mounted to the support and has a
wall with a bond coat and a thermal barrier coating. Both the bond
coat and the thermal barrier coating have a radial thickness
between 3% and 10% of the total radial thickness of the wall.
[0006] A blade outer air seal for a gas turbine engine includes a
first side surface, a second side surface, a wall, and one or more
forward hooks. The one or more forward hooks extend from the wall
and at least one of the hooks has a slot therein that is offset
relative to an axis of symmetry of the blade outer air seal.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] FIG. 1 is a partial cross-sectional view of an exemplary gas
turbine engine.
[0008] FIG. 2 is an enlarged view of a turbine portion of the gas
turbine engine shown in FIG. 1 with a BOAS mounted therein.
[0009] FIG. 3 is a perspective view of one embodiment of the
BOAS.
[0010] FIG. 4 is a plane view of the outer radial surface of the
BOAS of FIG. 3.
[0011] FIG. 4A is a cross-sectional view of the BOAS of FIG. 4.
DETAILED DESCRIPTION
[0012] The present invention provides a BOAS design with higher
convective efficiency and with improved durability due to improved
corrosion and oxidation resistance. More particularly, the BOAS
described herein utilizes optimally sized holes in an outer
diameter surface of a wall and optimally sized passages within the
wall to better control cooling air flow through the BOAS and
thereby improve convective efficiency of the BOAS. These features
improve the operational longevity of the BOAS. Additionally, the
BOAS is adapted with features such as a non-symmetric slot and an
angled hook wall that extends radially and axially to aid in
assembly of the BOAS within a gas turbine engine.
[0013] An exemplary industrial gas turbine engine 10 is
circumferentially disposed about a central, longitudinal axis or
axial engine centerline axis 12 as illustrated in FIG. 1. The
engine 10 includes in series order from front to rear, low and high
pressure compressor sections 16 and 18, a central combustor section
20 and high and low pressure turbine sections 22 and 24. In some
examples, a free turbine section 26 is disposed aft of the low
pressure turbine 24. Although illustrated with reference to an
industrial gas turbine engine, this application also extends to
aero engines with a fan or gear driven fan, and engines with more
or fewer sections than illustrated.
[0014] As is well known in the art of gas turbines, incoming
ambient air 30 becomes pressurized air 32 in the compressors 16 and
18. Fuel mixes with the pressurized air 32 in the combustor section
20, where it is burned. Once burned, combustion gases 34 expand
through turbine sections 22, 24 and power turbine 26. Turbine
sections 22 and 24 drive high and low rotor shafts 36 and 38
respectively, which rotate in response to the combustion products
and thus the attached compressor sections 18, 16. Free turbine
section 26 may, for example, drive an electrical generator, pump,
or gearbox (not shown).
[0015] It is understood that FIG. 1 provides a basic understanding
and overview of the various sections and the basic operation of an
industrial gas turbine engine. It will become apparent to those
skilled in the art that the present application is applicable to
all types of gas turbine engines, including those with aerospace
applications.
[0016] FIG. 2 is an enlarged view of a turbine section 22 and/or 24
of gas turbine engine 10 shown in FIG. 1 with a blade outer air
seal (BOAS) 40 disposed adjacent a turbine rotor blade 46. FIG. 2
illustrates BOAS 40, an engine case 42, stator vanes 44A and 44B,
rotor blade 46, a BOAS support 48, and a band segment 50. Vanes 44A
and 44B include platforms 43A and 43B.
[0017] BOAS 40 comprises an arcuate shroud segment with an inner
diameter wall forming the outer diameter of the engine flow path
through which combustion gases 34 pass. As will be discussed
subsequently, passages (not numbered) extend through at least a
portion of wall of BOAS 40 to provide for cooling of BOAS 40 during
operation. BOAS 40 is mounted within engine case 42 by forward and
aft hooks, which engage BOAS support 48 and vane platform 43B,
respectively. BOAS support 48 and vane platforms 43A and 43B are in
turn connected to engine case 42. Band segment 50 is positioned
radially outward of BOAS 40 and extends between BOAS support 48 and
vane platform 43B. Conformal seals such as w-seals are disposed
between vane platform 43B, BOAS support 48, and BOAS 40.
[0018] Rotor blade 46 comprises a single blade in a rotor stage
disposed downstream of combustor section 20 (FIG. 1). The rotor
stage extends in a circumferential direction about engine
centerline axis 12 (FIG. 1) and has a plurality of rotor blades 46.
During operation, combustion gases 34 pass between adjacent rotor
blades 46 and pass downstream to stator vane 44B. Rotor blade 46 is
disposed radially inward of BOAS 40, with respect to engine
centerline axis 12 as shown in FIG. 1.
[0019] Stator vanes 44A and 44B are disposed axially forward and
rearward of BOAS 40, respectively. Like the rotor stage, the stator
stages (of which vanes 44A and 44B are a part) extend in a
circumferential direction about engine center line axis 12, and
each stage has a plurality of stator vanes. Vanes 44A and 44B
include outer diameter platforms 43A and 43B, respectively.
Platforms 43A and 43B include features that facilitate the mounting
stator vanes 44A and 44B to engine case 42.
[0020] In operation, the flow of combustion gases 34 impinges upon
vanes 44A and 44B and is aligned for a subsequent rotor stage. As
the flow of combustion gases 34 passes through turbine blades 46
between a blade platform (not shown) and BOAS 40 the flow of
combustion gases 40 impinges upon rotor blade 46 causing the rotor
stage to rotate about engine center line axis 12 (FIG. 1). BOAS 40
is mounted just radially outward from rotor blade 46 tip and
provides a seal against combustion gases 34 radially bypassing
rotor blade 46. The flow of combustion gases 34 exits rotor stage
and enters stator vane stage passing vane 44B.
[0021] Engine case 42 and other components including vane platforms
43A and 43B form plenums that can be used to communicate cooling
air A to various components including BOAS 40, and in some
embodiments, vanes 44A and 44B. Generally, cooling air A is
supplied to plenums from a source such as high-pressure stage 18
and/or intermediate pressure stage of compressor (FIG. 1). Cooling
air A passes through components such as BOAS 40 via passages (not
shown) to provide for convection cooling. Thus, cooling air A
provides desired cooling in order to increase the operational life
of BOAS 40.
[0022] FIG. 3 provides a perspective view of BOAS 40. BOAS 40
includes a wall 51, an outer diameter surface 52, an inner diameter
surface 54, a first side surface 56, a second side surface 58, a
forward surface 60, an aft surface 62, a forward hooks 64, an aft
hook 65, lateral hooks 66, holes 68A and 68B, a slot 72, an angled
wall 74 and a outer radial hook surface 75.
[0023] Wall 51 of BOAS 40 has outer diameter surface 52, which
extends between first side surface 56 and second side surface 58
and between forward hooks 64 and aft hooks 65. Wall 51 has a total
radial thickness T between outer diameter surface 52 and inner
diameter surface 54. Thickness T of wall 51 can vary from
embodiment to embodiment, and can include a bond coat and/or a
thermal barrier coating.
[0024] Inner diameter surface 54 is disposed on an opposing side of
wall 51 from outer diameter surface 52. Inner diameter surface 54
extends between first side surface 56 and second side surface 58
and between forward surface 60 and aft surface 62. When BOAS 40 is
installed in gas turbine engine 10 (FIG. 1), inner diameter surface
54 interfaces with and forms the outer diameter of the engine flow
path through which combustion gases 34 pass (FIGS. 1 and 2). As
will be discussed subsequently, in some embodiments inner diameter
surface 54 is formed by application of bond coat and thermal
barrier coating.
[0025] First and second side surfaces 56 and 58 are disposed to
either lateral side of BOAS 40. First and second side surfaces 56
and 58 intersect with forward surface 60. Forward surface 60 is
disposed axially forward (with respect to direction of flow of the
combustion gases 34 through engine flow path) of forward hooks
64.
[0026] First and second side surfaces 56 and 58 also intersect with
aft surface 62, which extends radially inward of aft hook 65. Aft
hook 65 extends from wall 51 and is adapted to be received in a
recess in vane platform 43B (FIG. 2). Similarly, forward hooks 64
extend radially outward and forward from wall 51 and are adapted to
be received in BOAS support 48 (FIG. 2). Lateral hooks 66 extend
radially outward from both wall 51 and forward hooks 64 over first
side surface 56. Lateral hooks 66 overlap adjacent BOAS (not shown)
when assembled.
[0027] Holes 68A and 68B (other holes not shown) are formed in wall
51 and extend through outer diameter surface 52 adjacent second
side surfaces 58. Holes 68A and 68B communicate with passages (FIG.
4A), which extend generally laterally through wall 51 from first
side surface 56 to second side surface 58. As will be discussed
subsequently, holes, including holes 68A and 68B, are sized to
allow for the passage of optimal amounts of cooling air A (FIG. 2)
into and through BOAS 40 in order to increase the operational life
of BOAS 40.
[0028] In the embodiment shown in FIG. 3, forward hooks 64 are
separated by slots including slot 72. Slot 72 is offset relative to
a lateral axis of symmetry A.sub.SM of BOAS 40. Thus, BOAS 40,
including forward hooks 64, has an asymmetric design in the lateral
direction. Once assembled in gas turbine engine 10 (FIG. 1), slot
72 receives an anti-rotation feature (not shown) of BOAS support 48
(FIG. 2). Slot 72 and anti-rotation feature prevent lateral
movement (movement circumferentially around rotor stage within
circumferential engine case 42) of BOAS 40.
[0029] Angled wall 74 extends radially and axially from outer
radial hook surface 75 to connect to outer diameter surface 52 of
wall 51. Angled wall 74 provides for ease of identification of BOAS
40 during assembly and disassembly processes.
[0030] FIG. 4 shows a plane view of the outer diameter of BOAS 40.
FIG. 4A shows a cross-sectional view of BOAS 40. As illustrated in
FIGS. 4 and 4A, BOAS 40 includes outer diameter surface 52, inner
diameter surface 54 (FIG. 4A), first side surface 56 (FIG. 4),
second side surface 58 (FIG. 4), forward surface 60, aft surface
62, forward hooks 64, aft hook 65, lateral hooks 66 (FIG. 4), slot
72 (FIG. 4), passages 70 (FIGS. 3 and 4A), angled wall 74 and outer
radial hook surface 75. In addition to holes 68A and 68B, FIG. 4
illustrates holes 68C, 68D, 68E, and 68F. FIG. 4A illustrates a
bond coat 76 and a thermal barrier coating 78.
[0031] As illustrated in FIG. 4, holes 68A, 68B, and 68C extend
through outer diameter surface 52 adjacent second side surface 58
and holes 68D, 68E, and 68F extend through outer diameter surface
52 adjacent first side surface 56. As discussed, holes 68A, 68B
68C, 68D, 68E, and 68F communicate with passages 70 (FIGS. 3 and
4A). Holes 68A, 68B 68C, 68D, 68E, and 68F having varying diameters
and are sized to allow for the passage of optimal amounts of
cooling air A (FIG. 2) into and through BOAS 40 in order to
increase the operational life of BOAS 40. Thus, in one embodiment
each hole 68A, 68B 68C, 68D, 68E, and 68F has an area between about
0.005% and 0.45% of the surface area of BOAS 40 (as measured along
a plane extending between first side surface 56, second side
surface 58, forward surface 60, and aft surface 62). In a further
embodiment each hole 68A, 68B 68C, 68D, 68E, and 68F has an area
between about 0.020% and 0.30% of the surface area of BOAS 40 (as
measured along a plane extending between first side surface 56,
second side surface 58, forward surface 60, and aft surface
62).
[0032] As shown in FIG. 4A, in one embodiment each passage 70 has a
radial height H.sub.1 that comprises between about 30% to 40% of
the total radial thickness T of wall 51. Passages 70 axial length L
in total (between all six passages) comprises between about 75% and
85% of the axial length of BOAS 40 between forward surface 60 and
aft surface 62.
[0033] FIG. 4A additionally shows bond coat 76, which is added to
the wall 51. In one embodiment, bond coat 76 comprises a metallic
coating that provides for increased oxidation and corrosion
resistance. Bond coat 76 can be a nickel alloy layer applied using
plasma or high velocity oxy-fuel deposition processes. In one
embodiment, bond coat 76 has a radial thickness H.sub.2 between
about 3% and 10% of the total radial thickness T of wall 51.
[0034] Thermal barrier coating 78 can be applied to bond coat 76 to
form inner radial surface 52. In one embodiment, thermal barrier
coating comprises a ceramic layer that simultaneously provides
thermal insulation and abradability and has a thickness H.sub.3
between about 3% and 10% of the total radial thickness T of wall
51. The thermal bearing coating 78 can be applied using plasma
deposition or other known methods.
[0035] BOAS 40 can be constructed of metallic material such as a
nickel base alloy that offers high temperature strength and hot
corrosion resistance. In one embodiment, BOAS 40 is formed of a
single crystal alloy that is cast and directionally solidified. The
alloy can additionally be heat treated at various temperature
ranges for varying durations as desired.
[0036] The present invention provides a BOAS design with higher
convective efficiency and with improved durability due to improved
corrosion and oxidation resistance. More particularly, the BOAS
described herein utilizes optimally sized holes in an outer
diameter surface of a wall and optimally sized passages within the
wall to better control cooling air flow through the BOAS and
thereby improve convective efficiency of the BOAS. These features
improve the operational longevity of the BOAS. Additionally, the
BOAS is adapted with features such as a non-symmetric slot and an
angled hook wall that extends radially and axially to aid in
assembly of the BOAS within a gas turbine engine.
DISCUSSION OF POSSIBLE EMBODIMENTS
[0037] The following are non-exclusive descriptions of possible
embodiments of the present invention.
[0038] A blade outer air seal for a gas turbine engine includes a
first side surface, a second side surface, and a wall. The wall
extends between the first side surface and the second side surface
and has one or more holes formed therein. The holes are spaced from
the first side surface and/or the second side surface and have
areas between about 0.005% and 0.450% of a surface area of the
blade outer air seal.
[0039] The blade outer air seal of the preceding paragraph can
optionally include, additionally and/or alternatively, any one or
more of the following features, configurations and/or additional
components.
[0040] The one or more holes have areas between about 0.02% and
0.30% of a surface area of the blade outer air seal.
[0041] The one or more holes comprise six holes with three holes
positioned adjacent the first side surface and three holes
positioned adjacent the second side surface.
[0042] Internal passages extend through the wall from the first
side surface to the second side surface, and wherein the one or
more holes communicate with the internal passages.
[0043] The six holes comprise one hole for each of the internal
passages.
[0044] Each of the internal passages has a radial height between
30% to 40% of an total radial thickness the wall.
[0045] The internal passages together have an axial length that
comprises between 75% and 85% of the axial length of the wall.
[0046] One or more forward hooks extend from the wall, and at least
one of the forward hooks has a slot therein that is offset relative
to an axis of symmetry of the blade outer air seal.
[0047] At least one of the forward hooks has an angled wall that
extends from an outer radial surface of the at least one of the
forward hooks to the wall.
[0048] The wall includes a bond coat, wherein the bond coat has a
radial thickness between 3% and 10% of the total radial thickness
of the wall.
[0049] The wall has a thermal barrier coating applied to an inner
radial surface thereof, wherein the thermal barrier coating has a
radial thickness between 3% and 10% of the total radial thickness
of the wall.
[0050] A blade outer air seal for a gas turbine engine includes a
first side surface, a second side surface, and a wall. The wall
extends between the first side surface and the second side surface
and has a bond coat and a thermal barrier coating. Both the bond
coat and the thermal barrier coating have a radial thickness
between 3% and 10% of the total radial thickness of the wall.
[0051] The blade outer air seal of the preceding paragraph can
optionally include, additionally and/or alternatively, any one or
more of the following features, configurations and/or additional
components.
[0052] One or more holes are formed in the wall and are spaced from
at least one of the first side surface or second side surface, and
the one or more holes have areas between about 0.005% and 0.450% of
a surface area of the blade outer air seal;
[0053] One or more holes are formed in the wall and are spaced from
at least one of the first side surface or second side surface, and
the one or more holes have areas between about 0.020% and 0.30% of
a surface area of the blade outer air seal;
[0054] The one or more holes comprise six holes with three holes
positioned adjacent the first side surface and three holes
positioned adjacent the second side surface.
[0055] Internal passages that extend through the wall from the
first side surface to the second side surface, and the one or more
holes communicate with the internal passages.
[0056] Each of the internal passages has a radial height between
30% to 40% of an total radial thickness the wall.
[0057] The internal passages together have an axial length that
comprises between 75% and 85% of the axial length of the wall.
[0058] One or more forward hooks extend from the wall, wherein at
least one of the forward hooks has a slot therein that is offset
relative to an axis of symmetry of the blade outer air seal;
and
[0059] At least one of the forward hooks has an angled wall extends
from an outer radial surface of the at least one of the forward
hooks to the wall.
[0060] A blade outer air seal for a gas turbine engine includes a
first side surface, a second side surface, a wall, and one or more
forward hooks. The one or more forward hooks extend from the wall
and at least one of the hooks has a slot therein that is offset
relative to an axis of symmetry of the blade outer air seal.
[0061] The blade outer air seal of the preceding paragraph can
optionally include, additionally and/or alternatively, any one or
more of the following features, configurations and/or additional
components.
[0062] At least one of the forward hooks has an angled wall that
extends from an outer radial surface of the at least one of the
forward hooks to the wall.
[0063] While the invention has been described with reference to an
exemplary embodiment(s), it will be understood by those skilled in
the art that various changes may be made and equivalents may be
substituted for elements thereof without departing from the scope
of the invention. In addition, many modifications may be made to
adapt a particular situation or material to the teachings of the
invention without departing from the essential scope thereof.
Therefore, it is intended that the invention not be limited to the
particular embodiment(s) disclosed, but that the invention will
include all embodiments falling within the scope of the appended
claims.
* * * * *