U.S. patent application number 13/597745 was filed with the patent office on 2014-03-06 for blade outer air seal.
The applicant listed for this patent is Russell E. Keene, Brian R. Pelletier, Dmitriy A. Romanov. Invention is credited to Russell E. Keene, Brian R. Pelletier, Dmitriy A. Romanov.
Application Number | 20140064969 13/597745 |
Document ID | / |
Family ID | 50184124 |
Filed Date | 2014-03-06 |
United States Patent
Application |
20140064969 |
Kind Code |
A1 |
Romanov; Dmitriy A. ; et
al. |
March 6, 2014 |
BLADE OUTER AIR SEAL
Abstract
An example blade outer air seal assembly for a gas turbine
engine includes a main body portion extending along an axis. The
main body portion has a radially inward facing surface and at least
one radially outward facing surface. A passage is provided in the
main body portion between the at least one radially inward facing
surface and the at least one radially outward facing surface. The
passage has a first portion and a second portion transverse to the
first portion, and at least one rib disposed along the axis
radially outward of the second portion.
Inventors: |
Romanov; Dmitriy A.; (Wells,
ME) ; Pelletier; Brian R.; (Berwick, ME) ;
Keene; Russell E.; (Arundel, ME) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Romanov; Dmitriy A.
Pelletier; Brian R.
Keene; Russell E. |
Wells
Berwick
Arundel |
ME
ME
ME |
US
US
US |
|
|
Family ID: |
50184124 |
Appl. No.: |
13/597745 |
Filed: |
August 29, 2012 |
Current U.S.
Class: |
416/174 ;
29/889.7 |
Current CPC
Class: |
Y10T 29/49336 20150115;
F05D 2230/90 20130101; F01D 11/08 20130101; Y02T 50/6765 20180501;
Y02T 50/676 20130101; F01D 5/20 20130101; Y02T 50/60 20130101; F05D
2260/22141 20130101 |
Class at
Publication: |
416/174 ;
29/889.7 |
International
Class: |
F01D 5/20 20060101
F01D005/20 |
Goverment Interests
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT
[0001] This disclosure was made with Government support under
N00019-12-D-0002-4Y01 awarded by The United States Navy. The
Government has certain rights in this disclosure.
Claims
1. A blade outer air seal assembly for a gas turbine engine,
comprising: a main body portion extending along an axis, the main
body portion having a radially inward facing surface and at least
one radially outward facing surface; a passage provided in the main
body portion between the at least one radially inward facing
surface and the at least one radially outward facing surface, the
passage having a first portion and a second portion transverse to
the first portion; and at least one rib disposed along the axis
radially outward of the second portion.
2. The assembly of claim 1, wherein the at least one rib is at
least partially in a cavity that opens in a radially outward facing
direction, the at least one rib comprising one of the radially
outward facing surfaces.
3. The assembly of claim 1, including at least one airflow inlet
provided by one of the radially outward facing surfaces.
4. The assembly of claim 3, wherein the at least one airflow inlet
includes a first airflow inlet opening to a cavity that opens in a
radially outward facing direction.
5. The assembly of claim 1, including at least one attachment
portion disposed at both a leading edge of the main body portion
and a trailing edge of the main body portion, each of the at least
one attachment portions having a flange extending axially aft.
6. The assembly of claim 5, wherein the at least one rib is
circumferentially aligned with one of the attachment portions
disposed at the leading edge.
7. The assembly of claim 1, wherein at least one fluid outlet is
disposed at a circumferential end of the main body portion radially
outward of the radially inward facing surface.
8. The assembly of claim 1, including at least one airflow inlet
provided by one of the radially outward facing surfaces, wherein
the at least one airflow inlet includes at least a first airflow
inlet having a cross-section greater than a second airflow
inlet.
9. The assembly of claim 1, wherein the main body portion comprises
a single crystal nickel alloy having a sulfur content that is equal
to or below 1 part per million.
10. The assembly of claim 1, wherein a thermal barrier coating is
disposed adjacent the radially inward facing surface, wherein a
thickness of the thermal barrier coating is substantially equal to
a thickness of an inner wall of the main body portion defined
between a floor of the passage and the radially inward facing
surface.
11. The assembly of claim 10, wherein the passage is entirely
radially outward of the thermal barrier coating.
12. The assembly of claim 1, wherein the second portion is radially
inward of the first portion.
13. The assembly of claim 1, wherein at least one of the radially
outward facing surfaces is exposed.
14. A gas turbine engine assembly, comprising: a plurality of blade
outer air seal assemblies each configured to attach to a casing and
disposed circumferentially about an engine axis, the blade outer
air seal assemblies each having a main body portion with a radially
inward facing surface and a radially outward facing surface,
wherein at least one passage is defined in the main body portion
between the radially inward facing surface and the radially outward
facing surface, the main body portion having an inner wall having a
first thickness defined between a floor of each of the at least one
passages and the radially inward facing surface; and a thermal
barrier coating adjacent to the radially inward facing surface, the
thermal barrier coating having a second thickness that is
substantially equal to the first thickness.
15. The gas turbine engine assembly of claim 14, wherein at least
one airflow inlet is opening to the radially outward facing
surface, the at least one of passage generally perpendicular to the
engine axis, the at least one passage in fluid communication with
the at least one inlet and configured to receive cooling air
flow.
16. The gas turbine engine assembly of claim 15, wherein the at
least one passage includes a first passage in fluid communication
with a first airflow inlet of the at least one airflow inlet having
airflow greater than a second passage in fluid communication with a
second airflow inlet.
17. The gas turbine engine assembly of claim 16, wherein the first
airflow inlet has a cross-section greater than the second airflow
inlet.
18. The gas turbine engine assembly of claim 15, wherein an airflow
source communicates compressor bleed air to the at least one
inlet.
19. The gas turbine engine assembly of claim 14, wherein each of
the plurality of blade outer air seal assemblies includes at least
one attachment portion disposed at each of a leading edge of the
main body portion and a trailing edge of the main body portion,
each of the at least one attachment portions having a flange
extending axially aft, each of the at least one attachment portions
aligned with corresponding receiving portion of the casing to
attach the plurality of blade outer air seal assemblies to the
casing.
20. The gas turbine engine assembly of claim 14, including at least
at least one rib is disposed along the engine axis radially outward
of the second portion.
21. A method of forming a blade outer air seal for a gas turbine
engine comprising: providing a main body portion having a radially
inward facing surface and a radially outward facing surface that
axially extend along an axis, wherein at least one passage is
defined in the main body portion between the radially inward facing
surface and the radially outward facing surface, the at least one
passage having a first portion extending generally radially and a
second portion extending generally circumferentially transverse to
the first portion, wherein the second portion has a floor, wherein
at least one rib is disposed along the axis on the radially outward
facing surface, the at least one rib radially outward of the
circumferential second portion; machining an inner wall defined
between the floor and the radially inward facing surface to reduce
a thickness of the inner wall; and depositing a thermal barrier
coating adjacent the radially inward facing surface, wherein the
thermal barrier coating is substantially equal to the reduced
thickness of the inner wall.
22. The method of claim 21, including the step of forming at least
one airflow inlet opening to the radially outward facing surface,
wherein the at least one passage is in fluid communication with the
at least one inlet and configured to receive cooling air flow.
Description
BACKGROUND
[0002] This disclosure relates generally to a blade outer air seal
and, more particularly, to enhancing the performance of a blade
outer air seal.
[0003] As known, gas turbine engines, and other turbomachines,
include multiple sections, such as a fan section, a compressor
section, a combustor section, a turbine section, and an exhaust
section. Air moves into the engine through the fan section. Airfoil
arrays in the compressor section rotate to compress the air, which
is then mixed with fuel and combusted in the combustor section. The
products of combustion are expanded to rotatably drive airfoil
arrays in the turbine section. Rotating the airfoil arrays in the
turbine section drives rotation of the fan and compressor
sections.
[0004] A blade outer air seal arrangement includes multiple blade
outer air seals circumferentially disposed about at least some of
the airfoil arrays. The tips of the blades within the airfoil
arrays seal against the blade outer air seals during operation.
Improving and maintaining the sealing relationship between the
blades and the blade outer air seals enhances performance of the
turbomachine. As known, the blade outer air seal environment is
exposed to temperature extremes and other harsh environmental
conditions, both of which can affect the integrity of the blade
outer air seal and the sealing relationship.
SUMMARY
[0005] An example blade outer air seal assembly for a gas turbine
engine includes a main body portion extending along an axis. The
main body portion has a radially inward facing surface and at least
one radially outward facing surface. A passage is provided in the
main body portion between the at least one radially inward facing
surface and the at least one radially outward facing surface. The
passage has a first portion and a second portion transverse to the
first portion, and at least one rib disposed along the axis
radially outward of the second portion.
[0006] In a further non-limiting embodiment according to the
previous assembly, the at least one rib is at least partially in a
cavity that opens in a radially outward facing direction. The at
least one rib comprises one of the radially outward facing
surfaces.
[0007] In a further non-limiting embodiment according to any of the
previous assemblies, at least one airflow inlet is provided by one
of the radially outward facing surfaces.
[0008] In a further non-limiting embodiment according to any of the
previous assemblies, the at least one airflow inlet includes a
first airflow inlet opening to the cavity.
[0009] In a further non-limiting embodiment according to any of the
previous assemblies, at least one attachment portion is disposed at
both a leading edge of the main body portion and a trailing edge of
the main body portion. Each of the at least one attachment portions
has a flange extending axially aft.
[0010] In a further non-limiting embodiment according to any of the
previous assemblies, the at least one rib is circumferentially
aligned with one of the attachment portions disposed at the leading
edge.
[0011] In a further non-limiting embodiment according to any of the
previous assemblies, at least one fluid outlet is disposed at a
circumferential end of the main body portion radially outward of
the radially inward facing surface.
[0012] In a further non-limiting embodiment according to any of the
previous assemblies, at least one airflow inlet is provided by one
of the radially outward facing surfaces. The at least a first
airflow inlet includes at least a first airflow inlet having a
cross-section greater than a second airflow inlet.
[0013] In a further non-limiting embodiment according to any of the
previous assemblies, the main body portion comprises a single
crystal nickel alloy having a sulfur content that is equal to or
below 1 part per million.
[0014] In a further non-limiting embodiment according to any of the
previous assemblies, a thermal barrier coating is disposed adjacent
the radially inward facing surface. A thickness of the thermal
barrier coating is substantially equal to a thickness of an inner
wall of the main body portion defined between a floor of each of
the at least one passages and the radially inward facing
surface.
[0015] In a further non-limiting embodiment according to any of the
previous assemblies, the passage is entirely radially outward of
the thermal barrier coating.
[0016] In a further non-limiting embodiment according to any of the
previous assemblies, the second portion is radially inward of the
first portion.
[0017] In a further non-limiting embodiment according to any of the
previous assemblies, at least one of the radially outward facing
surfaces is exposed.
[0018] An example gas turbine engine assembly includes a plurality
of blade outer air seal assemblies each configured to attach to a
casing and disposed circumferentially about an engine axis. The
blade outer air seal assemblies each have a main body portion with
a radially inward facing surface and a radially outward facing
surface. At least one passage is defined in the main body portion
between the radially inward facing surface and the radially outward
facing surface. The main body portion has an inner wall having a
first thickness defined between a floor of each of the at least one
passages and the radially inward facing surface. A thermal barrier
coating adjacent to the radially inward facing surface. The thermal
barrier coating has a second thickness that is substantially equal
to the first thickness.
[0019] In a further non-limiting embodiment according to the
previous gas turbine engine assembly, at least one airflow inlet is
opening to the radially outward facing surface. The at least one of
passage is generally perpendicular to the engine axis. The at least
one passage is in fluid communication with the at least one inlet
and configured to receive cooling air flow.
[0020] In a further non-limiting embodiment according to any of the
previous gas turbine engine assemblies, the at least one passage
includes a first passage in fluid communication with a first
airflow inlet of the at least one airflow inlets having airflow
greater than a second passage in fluid communication with the
second airflow inlet.
[0021] In a further non-limiting embodiment according to any of the
previous gas turbine engine assemblies, the first airflow inlet has
a cross-section greater than the second airflow inlet.
[0022] In a further non-limiting embodiment according to any of the
previous gas turbine engine assemblies, an airflow source
communicates compressor bleed air to the at least one inlet.
[0023] In a further non-limiting embodiment according to any of the
previous gas turbine engine assemblies, each of the plurality of
blade outer air seal assemblies includes at least one attachment
portion disposed at each of a leading edge of the main body portion
and a trailing edge of the main body portion. Each of the at least
one attachment portions have a flange extending axially aft. Each
of the at least one attachment portions aligned with corresponding
receiving portion of the casing attaches the plurality of BOAS
assemblies to the casing.
[0024] In a further non-limiting embodiment according to any of the
previous gas turbine engine assemblies, at least at least one rib
is disposed along the engine axis radially outward of the second
portion.
[0025] An example method of forming a blade outer air seal for a
gas turbine engine includes providing a main body portion having a
radially inward facing surface and a radially outward facing
surface that axially extend along an axis. At least one passage is
defined in the main body portion between the radially inward facing
surface and the radially outward facing surface. The at least one
passage has a first portion extending generally radially and a
second portion extending generally circumferentially transverse to
the first portion. The second portion has a floor. At least one rib
is disposed along the axis on the radially outward facing surface.
The at least one rib is radially outward of the circumferential
second portion. An inner wall defined between the floor and the
radially inward facing surface is machined to reduce a thickness of
the inner wall. A thermal barrier coating is deposited adjacent the
radially inward facing surface. The thermal barrier coating is
substantially equal to the reduced thickness of the inner wall.
[0026] In a further non-limiting embodiment according to the
previous method, at least one airflow inlet opening to the radially
outward facing surface is formed. The at least one passage is in
fluid communication with the at least one inlet and configured to
receive cooling air flow.
[0027] These and other features of the disclosed examples can be
best understood from the following specification and drawings, the
following of which is a brief description.
BRIEF DESCRIPTION OF THE FIGURES
[0028] FIG. 1 shows a cross-section of an example turbomachine.
[0029] FIG. 2 shows a cross-section of an example turbine section
of the turbomachine of FIG. 1.
[0030] FIG. 3 shows a perspective view of a blade outer air seal
assembly of the turbine section of FIG. 2.
[0031] FIG. 4 shows a top view of the blade outer air seal assembly
of the turbine section of FIG. 2.
[0032] FIG. 5 shows another perspective view of the blade outer air
seal assembly of the turbine section of FIG. 2.
[0033] FIG. 6 shows top cross-sectional view of the blade outer air
seal assembly of the turbine section of FIG. 2.
[0034] FIG. 7 shows method of forming a blade outer air seal
assembly.
DETAILED DESCRIPTION
[0035] Referring to FIG. 1, a gas turbine engine 20 is
schematically illustrated. The gas turbine engine 20 is disclosed
herein as a two-spool turbofan that generally incorporates a fan
section 22, a compressor section 24, a combustor section 26 and a
turbine section 28. Alternative engines might include an augmentor
section (not shown) among other systems or features. The fan
section 22 drives air along a bypass flowpath while the compressor
section 24 drives air along a core flowpath for compression and
communication into the combustor section 26 then expansion through
the turbine section 28. Although depicted as a turbofan gas turbine
engine in the disclosed non-limiting embodiment, it should be
understood that the concepts described herein are not limited to
use with turbofans as the teachings may be applied to other types
of turbine engines including three-spool architectures, non-geared
turbine engines, and land-based turbines.
[0036] The engine 20 generally includes a first spool 30 and a
second spool 32 mounted for rotation about an engine central axis A
relative to an engine static structure 36 via several bearing
systems 38. It should be understood that various bearing systems 38
at various locations may alternatively or additionally be
provided.
[0037] The first spool 30 generally includes a first shaft 40 that
interconnects a fan 42, a first compressor 43 and a first turbine
46. The first shaft 40 is connected to the fan 42 through a gear
assembly of a fan drive gear system 48 to drive the fan 42 at a
lower speed than the first spool 30. The second spool 32 includes a
second shaft 49 that interconnects a second compressor 52 and
second turbine 55. The first spool 30 runs at a relatively lower
pressure than the second spool 32. It is to be understood that "low
pressure" and "high pressure" or variations thereof as used herein
are relative terms indicating that the high pressure is greater
than the low pressure. An annular combustor 57 is arranged between
the second compressor 52 and the second turbine 55. The first shaft
40 and the second shaft 49 are concentric and rotate via bearing
systems 38 about the engine central axis A which is collinear with
their longitudinal axes.
[0038] The core airflow is compressed by the first compressor 43
then the second compressor 52, mixed and burned with fuel in the
annular combustor 57, then expanded over the second turbine 55 and
first turbine 46. The first turbine 46 and the second turbine 55
rotationally drive, respectively, the first spool 30 and the second
spool 32 in response to the expansion.
[0039] The engine 20 is a high-bypass geared aircraft engine that
has a bypass ratio that is greater than about six (6), with an
example embodiment being greater than ten (10), the gear assembly
of the fan drive gear system 48 is an epicyclic gear train, such as
a planetary gear system or other gear system, with a gear reduction
ratio of greater than about 2.3:1 and the first turbine 46 has a
pressure ratio that is greater than about 5. The first turbine 46
pressure ratio is pressure measured prior to inlet of first turbine
46 as related to the pressure at the outlet of the first turbine 46
prior to an exhaust nozzle. The first turbine 46 has a maximum
rotor diameter and the fan 42 has a fan diameter such that a ratio
of the maximum rotor diameter divided by the fan diameter is less
than 0.6. It should be understood, however, that the above
parameters are only exemplary.
[0040] A significant amount of thrust is provided by the bypass
flow B due to the high bypass ratio. The fan section 22 of the
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet. The flight
condition of 0.8 Mach and 35,000 feet, with the engine at its best
fuel consumption. To make an accurate comparison of fuel
consumption between engines, fuel consumption is reduced to a
common denominator, which is applicable to all types and sizes of
turbojets and turbofans. The term is thrust specific fuel
consumption, or TSFC. This is an engine's fuel consumption in
pounds per hour divided by the net thrust. The result is the amount
of fuel required to produce one pound of thrust. The TSFC unit is
pounds per hour per pounds of thrust (lb/hr/lb Fn). When it is
obvious that the reference is to a turbojet or turbofan engine,
TSFC is often simply called specific fuel consumption, or SFC. "Low
fan pressure ratio" is the pressure ratio across the fan blade
alone, without a Fan Exit Guide Vane system. The low fan pressure
ratio as disclosed herein according to one non-limiting embodiment
is less than about 1.45. "Low corrected fan tip speed" is the
actual fan tip speed in feet per second divided by an industry
standard temperature correction of
[(Tram.degree.R)/(518.7.degree.R)].sup.0.5. The "Low corrected fan
tip speed" as disclosed herein according to one non-limiting
embodiment is less than about 1150 feet per second.
[0041] Referring to FIG. 2, an example blade outer air seal (BOAS)
assembly 50 is attached to an inner engine case structure 44 of the
gas turbine engine 10 by a receiving portion 68 of the inner engine
case structure 44. In this example, the BOAS assembly 50 is located
within the turbine section 28 of the gas turbine engine 20. The
BOAS assembly 50 faces turbine blade 51 to define a radial tip
clearance 53 between the turbine blade 51 and the BOAS assembly 50.
Although only one BOAS assembly 50 is shown, a number of BOAS
assembly 50 are arranged circumferentially about engine axis A to
form a shroud. Alternatively, the BOAS assemblies 50 may be formed
as a unitary BOAS structure, with the same features described
herein.
[0042] Referring to FIGS. 3-6, the BOAS assembly 50 includes a main
body portion 54 that extends generally axially from a leading edge
portion 56 to a trailing edge portion 58 and from a radially
outward facing surface 62 to a radially inward facing surface 64.
The BOAS assembly 50 also includes at least one leading attachment
portion 60a ("attachment portions 60a") disposed at or near the
leading edge portion 56 and at least one trailing attachment
portion 60b ("attachment portions 60b") disposed at or near the
trailing edge portion 58. Each of the attachment portions 60a, 60b
define a flange 66 extending in an axially aft direction. Each
axially extending flange 66 corresponds to the receiving portion 68
of the inner engine case structure 44 to support and attach the
BOAS assembly 50 (Shown in FIG. 2). In this example, the attachment
portions 60a may be circumferentially offset, circumferentially
aligned, or a combination of both, from the attachment portions 60b
in response to BOAS assembly 50 parameters.
[0043] In this example, the BOAS assembly 50 includes a cavity 70
opening to the radially outward facing surface 62 between the
attachment portions 60a and the attachment portions 60b. It is
understood that other configurations of cavity 70 are contemplated
by this disclosure.
[0044] In this example, the main body portion 54 establishes at
least one rib 72 circumferentially aligned with corresponding
attachment portion 60a and circumferentially offset from attachment
portions 60b. The at least one rib 72 is disposed at least
partially within the cavity 70 and extends axially from
corresponding attachment portion 60a within the cavity 70 to the
trailing edge portion 58 adjacent attachment portions 60b. In
another example, the at least one rib 72 is circumferentially
aligned with corresponding attachment portions 60a and attachment
portion 60b and extends axially from attachment portion 60a to
corresponding attachment portion 60b. A ratio of the height 74 in
the radial direction of the at least one rib 72 to the width 76 in
the circumferential direction of the at least one rib 72 is between
1:1 and 1:10.
[0045] In this example, a plurality of fluid inlets 80 open to the
radially outward facing surface 62 near a first circumferential end
82 and a second circumferential end 84 of the main body portion 54.
The fluid inlets 80 may be located in the cavity 70 or
alternatively at other positions on the radially outward facing
surface 62. The fluid inlets 80 are varied in size based on
pre-determined cooling parameters. Fluid inlets 80a with a larger
surface area provide a greater amount of cooling airflow than fluid
inlets 80b with a smaller surface area, thus providing cooling flow
distribution based on thermal load distribution. In this way,
portions of the BOAS assembly 50 subject to relatively higher
thermal loads compared to other portion of the BOAS assembly 50
receive greater cooling by receiving cooling airflow through
relatively larger fluid inlets 80.
[0046] In this example, each of the fluid inlets 80 is in fluid
communication with a corresponding cooling passage 86 defined in
the main body portion 54 radially inwards of fluid inlet 80 and
cavity 70. Each cooling passage 86 includes a first portion 86a
generally transverse to a second portion 86b. The first portion 86a
extends in a generally circumferential direction from the first
circumferential end 82 to the second circumferential end 84 of the
main body portion 54 to provide cooling. The second portion 86b
extends in a generally radial direction to provide fluid
communication between inlet 80 and the first portion 86a. In this
example, each cooling passage 86 is defined entirely within the
main body portion 54.
[0047] The BOAS assembly 50 is in fluid communication with an
airflow source 90 (shown schematically), such as an upstream
compressor 24 or other source, such that fluid inlets 80 receive
cooling airflow, such as bleed compressor air. The cooling airflow
passes through fluid inlets 80 and is communicated to the first
portion 86a of corresponding cooling passage 86 via second portion
86b for cooling the BOAS assembly 50. The cooling airflow passes
through the cooling passage 86 from the fluid inlet 80 to a fluid
outlet 92 located at the circumferential end 82, 84 of the cooling
passage 86 opposite fluid inlet 80. The amount of cooling airflow
communicated to each cooling passage 86 is determined by the size
of fluid inlet 80. The BOAS assembly 50 also include a pressure
gradient which determines the amount of cooling airflow
communicated to each cooling passage 86. In this example, the fluid
inlets 80 have a cross sectional area 94 between about 0.00028 in.
2 (about 0.00181 cm 2) and about 0.0078 in. 2 (about 0.00503 cm 2).
Cooling airflow is communicated through each cooling passage 86 in
a single circumferential direction in this example. Plugs (not
shown) are inserted to close the cooling passage 86 at the
circumferential end 82, 84 corresponding to the fluid inlet 80 of
each passage.
[0048] In this example, the BOAS assembly 50 is made of a material
having sulfer levels at or below 1 part per million (PPM), such as
a single crystal nickel alloy, but other examples may include other
types of material. The thermal barrier coating 110 is a metallic or
ceramic based material in this example.
[0049] In this example, the BOAS assembly 50 includes a thermal
barrier coating 110 disposed on the radially inward facing surface
64 of the main body portion 54. The thermal barrier coating 110
includes a thermal layer 112 and a bond layer 114. A thickness 116
of the thermal barrier coating 110 is substantially equal to a
thickness 118 of an inner wall 120 of the main body portion 54
defined between a floor 122 of the plurality of cooling passages 86
and the radially inward facing surface 64. The term substantially
equal conveys that one of ordinary skill in the art would consider
the measurements to be the same within recognized tolerances.
[0050] The thickness 116 of the thermal barrier coating 110 and the
thickness 118 of the inner wall 120 are between 0.025 (0.0635 cm)
and 0.030 inches (0.0762 cm). In this example, the thickness 116 of
the thermal barrier coating 110 and the thickness 118 of the inner
wall 120 is about 0.027 inches (0.0686 cm).
[0051] During gas turbine engine 20 operation, the BOAS assembly 50
is subjected to different thermal loads and environmental
conditions. Cooling air flow from the airflow source 90 is provided
to the various fluid inlets 80, which communicate the cooling
airflow to the cooling passages 86 to provide varying levels of
cooling to different areas of the BOAS assembly 50 and effectively
communicate thermal energy away from the BOAS assembly 50 and the
tip of the rotating blade 51. The thermal barrier coating 110 is
provided on the radially inner facing surface 64 of the main body
portion 54 to provide additional protection from the thermal loads.
The at least one rib 72 provides stability to the BOAS assembly 50
to prevent axial deformation due to the reduction in BOAS assembly
50 material due to the use of cooling features and thermal bond
coating, as described in this disclosure.
[0052] Referring to FIG. 6, in another example, adjacent first
portions 86a of cooling passages 86 provided in the main body
portion 54 are connected at either the first circumferential end 82
or the second circumferential end 84 to form a serpentine passage
96 (shown schematically in phantom in FIG. 6) in fluid
communication with a single fluid inlet 80 or multiple fluid inlets
80 via one or more second portions 86b (as shown in FIG. 3).
[0053] Referring to FIG. 7, with continued reference to FIGS. 2-6,
a method of forming a BOAS assembly 200 includes providing a main
body portion along an axis A having a radially inward facing
surface and a radially outward facing surface that extend axially
between a leading edge portion and a trailing edge portion 202. At
least one passage is disposed in the main body portion between the
radially inward facing surface and the radially outward facing
surface 202. The at least one passage has a radial first portion
and an axial second portion transverse to the radial first portion.
The second portion defines a floor 202. At least one rib is
provided along the axis A 202. The at least one rib is radially
outward of the circumferential second portion 202. An inner wall
defined between the floor and the radially inward facing surface is
machined to reduce a thickness of the inner wall 204. A thermal
barrier coating is deposited adjacent the radially inward facing
surface 206. A thickness of the thermal barrier coating is
substantially equal to the reduced thickness of the inner wall
206.
[0054] The preceding description is exemplary rather than limiting
in nature. Variations and modifications to the disclosed examples
may become apparent to those skilled in the art that do not
necessarily depart from the essence of this disclosure. Thus, the
scope of legal protection given to this disclosure can only be
determined by studying the following claims.
* * * * *