U.S. patent application number 13/190559 was filed with the patent office on 2013-01-31 for blade outer air seal with passage joined cavities.
The applicant listed for this patent is Anne-Marie B. Thibodeau. Invention is credited to Anne-Marie B. Thibodeau.
Application Number | 20130028704 13/190559 |
Document ID | / |
Family ID | 46551421 |
Filed Date | 2013-01-31 |
United States Patent
Application |
20130028704 |
Kind Code |
A1 |
Thibodeau; Anne-Marie B. |
January 31, 2013 |
BLADE OUTER AIR SEAL WITH PASSAGE JOINED CAVITIES
Abstract
A blade outer air seal assembly includes a body that defines a
first cavity separated from a second cavity by a circumferential
rib. The circumferential rib includes at least one passage which
provides communication between the first cavity and the second
cavity.
Inventors: |
Thibodeau; Anne-Marie B.;
(Winslow, ME) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Thibodeau; Anne-Marie B. |
Winslow |
ME |
US |
|
|
Family ID: |
46551421 |
Appl. No.: |
13/190559 |
Filed: |
July 26, 2011 |
Current U.S.
Class: |
415/1 ;
415/173.1 |
Current CPC
Class: |
F05D 2260/201 20130101;
F01D 11/24 20130101; Y02T 50/60 20130101; F05D 2260/205 20130101;
F05D 2240/11 20130101 |
Class at
Publication: |
415/1 ;
415/173.1 |
International
Class: |
F01D 11/08 20060101
F01D011/08 |
Claims
1. A blade outer air seal assembly comprising: a body which defines
a first cavity separated from a second cavity by a circumferential
rib, said circumferential rib includes at least one passage which
provides communication between said first cavity and said second
cavity.
2. The blade outer air seal assembly as recited in claim 1, further
comprising an impingement plate which encloses said first cavity
and said second cavity.
3. The blade outer air seal assembly as recited in claim 1, wherein
said at least one passage comprises three passages.
4. The blade outer air seal assembly as recited in claim 3, wherein
said three passages removes approximately 3.5% of said rib.
5. The blade outer air seal assembly as recited in claim 1, wherein
said at least one passage removes approximately 3.5% of said
rib.
6. The blade outer air seal assembly as recited in claim 1, wherein
said first cavity is a forward cavity and said second cavity is
axially aft of said first cavity.
7. The blade outer air seal assembly as recited in claim 6, wherein
said at least one passage comprises three passages.
8. The blade outer air seal assembly as recited in claim 7, wherein
said three passages removes approximately 3.5% of said rib.
10. A method of communicating a secondary cooling airflow within a
gas turbine engine comprising: segregating a first cavity from a
second cavity by a circumferential rib, the circumferential rib
having at least one passage; and communicating secondary cooling
airflow between the first cavity and the second cavity through the
at least one passage.
11. The method as recited in claim 10, communicating the secondary
cooling airflow into the first cavity and the second cavity through
a single impingement plate.
12. The method as recited in claim 11, communicating the secondary
cooling airflow from the first cavity and the second cavity through
a multiple of edge holes to a core flow.
Description
BACKGROUND
[0001] The present application relates to a blade outer air seal
(BOAS) and more particularly to a multi-cavity blade outer air seal
(BOAS).
[0002] Gas turbine engines generally include fan, compressor,
combustor and turbine sections along an engine axis of rotation.
The fan, compressor, and turbine sections each include a series of
stator and rotor blade assemblies. A rotor and an axially adjacent
array of stator assemblies may be referred to as a stage. Each
stator vane assembly increases efficiency through the direction of
core gas flow into or out of the rotor assemblies.
[0003] An outer case, including a multiple of blade outer air seals
(BOAS), provides an outer radial flow path boundary. A multiple of
BOAS are typically provided to accommodate thermal and dynamic
variation typical in a high pressure turbine (HPT) section of the
gas turbine engine. The BOAS are subjected to relatively high
temperatures and receive a secondary cooling airflow for
temperature control. The secondary cooling airflow is communicated
into the BOAS then through cooling channels within the BOAS for
temperature control.
SUMMARY
[0004] A blade outer air seal assembly according to an exemplary
aspect of the present disclosure includes a body that defines a
first cavity separated from a second cavity by a circumferential
rib. The circumferential rib includes at least one passage which
provides communication between the first cavity and the second
cavity.
[0005] A method of communicating a secondary cooling airflow within
a gas turbine engine according to an exemplary aspect of the
present disclosure includes segregating a first cavity from a
second cavity by a circumferential rib, the circumferential rib
having at least one passage. Communicating secondary cooling
airflow between the first cavity and the second cavity through the
at least one passage.
BRIEF DESCRIPTION OF THE DRAWINGS
[0006] The various features and advantages of this invention will
become apparent to those skilled in the art from the following
detailed description of the disclosed non-limiting embodiment. The
drawings that accompany the detailed description can be briefly
described as follows:
[0007] FIG. 1 is a general sectional diagrammatic view of a gas
turbine engine HPT section;
[0008] FIG. 2 is a perspective exploded view of a BOAS segment;
and
[0009] FIG. 3 is a chart of pressures within the BOAS and axial
distance from a leading edge thereof;
DETAILED DESCRIPTION OF THE DISCLOSED EMBODIMENT
[0010] FIG. 1 schematically illustrates a gas turbine engine 20,
illustrated partially herein as a High Pressure Turbine (HPT)
section 22 disposed along a common engine longitudinal axis A. The
engine 20 includes a Blade Outer Air Seal (BOAS) assembly 24 to
provide an outer core gas path seal for the turbine section 22. It
should be understood that although a BOAS assembly for a HPT of a
gas turbine engine is disclosed in the illustrated embodiment, the
BOAS assembly may be utilized in any section of a gas turbine
engine. The BOAS segment may find beneficial use in many industries
including aerospace, industrial, electricity generation, naval
propulsion, pumping sets for gas and oil transmission, aircraft
propulsion, vehicle engines, and stationary power plants.
[0011] The HPT section 22 generally includes a rotor assembly 26
disposed between forward and aft stationary vane assemblies 28, 30.
Outer vane supports 28A, 30A attach the respective vane assemblies
to an engine case 32 (illustrated schematically). The rotor
assembly 26 generally includes a multiple of airfoils 34
circumferentially disposed around a disk 36. The distal end of each
airfoil 34 may be referred to as an airfoil tip 34T which rides
adjacent to the BOAS assembly 24.
[0012] The BOAS assembly 24 is generally disposed in an annulus
radially between the engine case 32 and the airfoil tips 34T. The
BOAS assembly 24 generally includes a blade outer air seal (BOAS)
support 38 and a multiple of blade outer air seal (BOAS) segments
40 mountable thereto (also see FIG. 2). The BOAS support 38 is
mounted within the engine case 32 to define forward and aft flanges
42, 44 to receive the BOAS segments 40. The forward flanges 42 and
the aft flanges 44 may be circumferentially segmented to receive
the BOAS segments 40 in a circumferentially rotated and locked
arrangement as generally understood. It should be understood that
various interfaces and BOAS assemblies may alternatively be
provided.
[0013] Each BOAS segment 40 includes a body 46 which defines a
forward interface 48 and an aft interface 50. The forward interface
48 and the aft interface 50 respectively engage the flanges 42, 44
to secure each individual BOAS segment 40 thereto.
[0014] With reference to FIG. 2, each BOAS segment 40 includes at
least two cavities 52A, 52B to receive a secondary cooling airflow
S. Each cavity 52A, 52B may be formed through, for example, an
investment casting process then closed by a single impingement
plate 54.
[0015] In the disclosed non-limiting embodiment, the cavity 52A is
axially forward of cavity 52B but separated therefrom by a
circumferential rib 56. That is, the circumferential rib 56
essentially surrounds the engine longitudinal axis A. Secondary
cooling air S flows through the plate 54, impinges in the BOAS
cavities 52A, 52B then flows out to the core gaspath flow through a
multiple of edge holes 60. The circumferential rib 56 and plate 54
isolates the secondary cooling air allocated to a specific cavity
52A, 52B. It should be understood that various alternative cavity
and passageway arrangements may be provided.
[0016] The circumferential rib 56 includes at least one passage 58
which provides for secondary cooling air S to flow aftward from the
forward cavity 52A to the aft cavity 52B. It should be understood
that the term "passage" as utilized herein may include various
slots, apertures, openings, holes and paths. In the disclosed
non-limiting embodiment, three passages 58 are provided which
removes approximately 3.5% of the rib 56. Since the percentage of
material removed is minimal and since the removal of the material
is from a circumferential member rather than an axial member,
minimal, if any, structural impact is experienced by the BOAS
segment 40.
[0017] The passages 58 allows some of the forward cavity 52A
secondary cooling air S to be reused in the aft cavity 52B which
results in lower temperatures and relatively lower cooling flow
requirements for the BOAS segment 40. The passages 58 also permits
at least some reuse of the secondary cooling air S with but the
single plate 54 which need not be welded to the rib 56. The single
plate 54 facilitates manufacture with minimal brazing filler metal
(BFM) and minimizes undesired leakage.
[0018] In the disclosed non-limiting embodiment, the secondary
cooling air S gaspath pressure within the BOAS segment 40 is lower
axially aft of airfoil tips 34T (FIG. 1). The forward cavity 52A
thus has a somewhat higher static pressure than the aft cavity 52B
(FIG. 3) due to the direction of primary core flow. The higher
static pressure in cavity 52A also results in increased axial
crossflow heat transfer coefficient (Hc) in the forward cavity 52A
which results in, for example, lower temperatures, and, likewise,
longer operational life of the BOAS segment 40 as represented in
the chart below:
TABLE-US-00001 RELATED ART 2-Cavity with 3 2-Cavity Total Flow
passages - Total Flow % Wae 0.69% 0.59% Tmet, mx (max BOAS
2159.degree. F. 2151.degree. F. temperature) Nominal/min back
13.2%/5% 12.6% flow margin (BFM)
[0019] It should be understood that like reference numerals
identify corresponding or similar elements throughout the several
drawings. It should also be understood that although a particular
component arrangement is disclosed in the illustrated embodiment,
other arrangements will benefit herefrom.
[0020] Although particular step sequences are shown, described, and
claimed, it should be understood that steps may be performed in any
order, separated or combined unless otherwise indicated and will
still benefit from the present invention.
[0021] The foregoing description is exemplary rather than defined
by the limitations within. Many modifications and variations of the
present invention are possible in light of the above teachings. The
disclosed embodiments of this invention have been disclosed,
however, one of ordinary skill in the art would recognize that
certain modifications would come within the scope of this
invention. It is, therefore, to be understood that within the scope
of the appended claims, the invention may be practiced otherwise
than as specifically described. For that reason the following
claims should be studied to determine the true scope and content of
this invention.
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