U.S. patent application number 17/647235 was filed with the patent office on 2022-04-28 for pressure regulated piston seal for a gas turbine combustor liner.
The applicant listed for this patent is General Electric Company. Invention is credited to Kirk Douglas Gallier, Jason Paul Hoppa, Robert Proctor, Stephen Gerard Schadewald, Jinjie Shi, Christopher Edward Wolfe.
Application Number | 20220128236 17/647235 |
Document ID | / |
Family ID | |
Filed Date | 2022-04-28 |
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United States Patent
Application |
20220128236 |
Kind Code |
A1 |
Shi; Jinjie ; et
al. |
April 28, 2022 |
PRESSURE REGULATED PISTON SEAL FOR A GAS TURBINE COMBUSTOR
LINER
Abstract
A seal assembly to seal a gas turbine hot gas path flow at an
interface of a combustor liner and a downstream component, such as
a stage one turbine nozzle, in a gas turbine. The seal assembly
including a piston ring seal housing, defining a cavity, and a
piston ring disposed within the cavity. The piston ring disposed
circumferentially about the combustor liner. The piston ring is
responsive to a regulated pressure to secure sealing engagement of
the piston ring and outer surface of the combustor liner. The seal
assembly includes at least one of one or more sectional
through-slots, bumps or channel features to provide for a flow
therethrough of a high-pressure (P.sub.high) bypass airflow exiting
a compressor to the cavity. The high-pressure (P.sub.high) bypass
airflow exerting a radial force on the piston ring.
Inventors: |
Shi; Jinjie; (Clifton Park,
NY) ; Proctor; Robert; (Mason, OH) ; Hoppa;
Jason Paul; (West Chester, OH) ; Schadewald; Stephen
Gerard; (Clifton Park, NY) ; Wolfe; Christopher
Edward; (Niskayuna, NY) ; Gallier; Kirk Douglas;
(Liberty Township, OH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Appl. No.: |
17/647235 |
Filed: |
January 6, 2022 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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16907013 |
Jun 19, 2020 |
11221140 |
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17647235 |
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15493549 |
Apr 21, 2017 |
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16907013 |
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International
Class: |
F23R 3/00 20060101
F23R003/00; F16J 15/16 20060101 F16J015/16; F23R 3/60 20060101
F23R003/60; F16J 15/44 20060101 F16J015/44; F01D 9/02 20060101
F01D009/02 |
Claims
1. A seal assembly to seal a gas turbine hot gas path flow at an
interface of a combustor liner and a downstream component in a gas
turbine, the seal assembly comprising: a piston ring seal housing
having defined therein a cavity; and a piston ring disposed within
the cavity of the piston ring seal housing and circumferentially
about the combustor liner, the piston ring being responsive to a
regulated pressure to secure sealing engagement of the piston ring
and an outer surface of the combustor liner, the piston ring
including at least one arcuate seal ring segment, wherein the
piston ring seal housing comprises one or more bumps protruding
from at least one of an upstream surface of the piston ring and a
front wall surface of the piston ring seal housing, to guide a flow
therethrough of a high-pressure (P.sub.high) bypass airflow exiting
an upstream component and to exert a pressurized force on an
outermost radial surface of the piston ring in a direction
perpendicular to a main gas flow regardless of an angular position
of the piston ring relative to the piston ring seal housing.
2. The seal assembly according to claim 1, wherein the one or more
bumps protruding from the at least one of the upstream surface of
the piston ring and the front wall surface of the piston ring seal
housing define a gap between the upstream surface of the piston
ring and the front wall surface of the piston ring seal
housing.
3. The seal assembly according to claim 1, wherein the downstream
component is a stage one turbine nozzle.
4. The seal assembly according to claim 1, wherein one of the
piston ring seal housing or the piston ring is rotated relative to
the other of the piston ring seal housing or the piston ring.
5. The seal assembly according to claim 1, wherein the piston ring
further comprises a cockle spring disposed circumferentially
thereabout and exerting a radially inward force on the piston
ring.
6. The seal assembly according to claim 1, wherein the one or more
bumps are configured in a plurality of columns and rows.
7. A gas turbine comprising: a combustor liner; a stage one nozzle
disposed downstream of the combustor liner; a piston seal assembly
defined at an interface of the combustor liner and the stage one
nozzle to seal a gas turbine hot gas path flow, the piston seal
assembly comprising: a piston ring seal housing having defined
therein a cavity; and a piston ring disposed within the cavity of
the piston ring seal housing and circumferentially about the
combustor liner, the piston ring being responsive to a regulated
pressure to secure sealing engagement of the piston ring and an
outer surface of the combustor liner, the piston ring including at
least one arcuate seal ring segment, wherein the piston ring seal
housing comprises one or more bumps protruding from at least one of
an upstream surface of the piston ring and a front wall surface of
the piston ring seal housing, to guide a flow therethrough of a
high-pressure (P.sub.high) bypass airflow exiting an upstream
component and to exert a pressurized force on an outermost radial
surface of the piston ring in a direction perpendicular to a main
gas flow regardless of an angular position of the piston ring
relative to the piston ring seal housing.
8. The gas turbine according to claim 7, wherein one of the piston
ring seal housing or the piston ring is rotated relative to the
other of the piston ring seal housing or the piston ring during a
takeoff condition.
9. The gas turbine according to claim 7, wherein the one or more
bumps protruding from the at least one of the upstream surface of
the piston ring and the front wall surface of the piston ring seal
housing define a gap between the upstream surface of the piston
ring and the front wall surface of the piston ring seal
housing.
10. The gas turbine according to claim 7, wherein the piston ring
further includes a cockle spring disposed circumferentially
thereabout and exerting a radially inward force on the piston
ring.
11. The gas turbine according to claim 7, wherein the one or more
bumps are configured in a plurality of columns and rows.
12. A seal assembly to seal a gas turbine hot gas path flow at an
interface of a combustor liner and a downstream component in a gas
turbine, the seal assembly comprising: a piston ring seal housing
having defined therein a cavity; and a piston ring disposed within
the cavity of the piston ring seal housing and circumferentially
about the combustor liner, the piston ring being responsive to a
regulated pressure to secure sealing engagement of the piston ring
and an outer surface of the combustor liner, the piston ring
including at least one arcuate seal ring segment, wherein the
piston ring seal housing comprises an opening in at least one of a
back wall of the piston ring seal housing and an outermost radial
face of the piston ring seal housing, to guide a flow therethrough
of a high-pressure (P.sub.high) bypass airflow exiting an upstream
component and to exert a pressurized force on an outermost radial
surface of the piston ring in a direction perpendicular to a main
gas flow regardless of an angular position of the piston ring
relative to the piston ring seal housing.
13. The seal assembly according to claim 12, wherein the opening
includes one or more sectional through-slots.
14. The seal assembly according to claim 12, wherein the opening
consists of a single continuous through-slot.
15. The seal assembly according to claim 12, wherein the downstream
component is a stage one turbine nozzle.
16. The seal assembly according to claim 12, wherein one of the
piston ring seal housing or the piston ring is rotated relative to
the other of the piston ring seal housing or the piston ring.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application is a divisional of U.S. patent application
Ser. No. 16/907,013 filed Jun. 19, 2020, which is a divisional of
U.S. patent application Ser. No. 15/493,549 filed Apr. 21, 2017,
the contents of which are hereby incorporated by reference in their
entireties.
BACKGROUND
[0002] This disclosure relates generally to turbine engine
combustors and, more particularly, to a piston seal assembly for a
combustor liner.
[0003] Gas turbine engines feature combustors as components. Air
enters the engine and passes through a compressor. The compressed
air is routed through one or more combustors. Within a combustor
are one or more nozzles that serve to introduce fuel into a stream
of air passing through the combustor. The resulting fuel-air
mixture is ignited in the combustor by igniters to generate hot,
pressurized combustion gases in the range of about 1100.degree. C.
to 2000.degree. C. that expand through a turbine nozzle. The burned
air-fuel mixture is routed out of the combustor through the turbine
nozzle, which directs the flow to downstream high and low-pressure
turbine stages. In these stages, the expanded hot gases exert
forces upon turbine blades, thus providing additional rotational
energy, for example, to drive a power-producing generator.
[0004] Turbine engine operators desire high efficiency while also
achieving low emissions. At least some known turbine engines
include a plurality of seal assemblies in a fluid flow path to
facilitate increasing the operating efficiency of the turbine. For
example, some known seal assemblies are coupled between a
stationary component of the engine and a rotary component of the
engine to provide sealing between a high-pressure area and a
low-pressure area. In at least some known gas turbine engines,
seals are provided between static components in adjacent stages, or
between components within a stage.
[0005] Of particular interest is a combustor of a turbine engine,
and more specifically, a combustor liner, such as ceramic matrix
composite (CMC), an adjacent stage one turbine nozzle and a piston
seal formed therebetween. Typically, the combustor liner includes a
seal housing support on a liner aft end where it joins an adjacent
stage one turbine nozzle. The piston seal is formed therebetween to
provide sealing and control of the cooling bypass flow flowing
between the combustor liner and the stage one turbine nozzle. The
amount of cooling flow through the piston seal plays an important
role in cooling the mechanical parts along the flow path.
Sufficient cooling flows are necessary to assure acceptable
lifetime, while too much cooling flow results in waste of
compressor air.
[0006] Conventional combustor liner piston seals are typically
comprised of a piston ring seal housing and a piston ring that
provides for sealing between the outer surface of the combustor
liner and the piston ring seal housing. Conventional piston seals
often fail when the piston ring seal housing is tilted or rotated.
More particularly, conventional piston seal designs typically allow
for a minimal degree of rotation (tilt) of the piston ring seal
housing relative to the piston ring during a takeoff condition. A
relative rotation between the piston ring seal housing and piston
ring of a greater degree causes the piston ring to block a gap that
is required between the front wall of the piston ring seal housing
and the piston ring. The gap allows for the passage therethrough of
a portion of a high-pressure (P.sub.high) compressor airflow from
an upstream compressor as a high-pressure (P.sub.high) bypass flow.
The high-pressure (P.sub.high) bypass flow ensures sufficient force
acting on the piston ring to engage the piston ring on the
combustor liner outer surface and on the aft wall of the piston
ring seal housing, and form a seal therebetween each. In response
to this blockage of the gap, the piston seal no longer seals
adequately and leakage from the seal is significantly higher than
desired. The lack of adequate sealing, not only wastes compressor
air that passes therethrough, but may also change the heat transfer
design point of the turbine engine.
[0007] Accordingly, it is desired to provide an improved piston
seal for sealing between stages of a turbine, and more particularly
between a combustor liner and a stage one turbine nozzle of a
turbine engine. It is desired that the piston seal provide for
leakage control during all flight conditions. More particularly, it
is desired to provide for a piston seal responsive to relative
rotation and motion of the piston ring seal housing to the piston
ring or vice versa during takeoff conditions.
BRIEF DESCRIPTION
[0008] Various embodiments of the disclosure include a piston seal
for a gas turbine engine, including a means for controlling such
seal by pressure regulation.
[0009] In accordance with one exemplary embodiment, disclosed is a
seal assembly to seal a gas turbine hot gas path flow at an
interface of a combustor liner and a downstream component in a gas
turbine. The seal assembly includes a piston ring seal housing and
a piston ring. The piston ring seal housing has defined therein a
cavity. The piston ring is disposed within the cavity of the piston
ring seal housing and circumferentially about the combustor liner.
The piston ring is responsive to a regulated pressure to secure
sealing engagement of the piston ring and an outer surface of the
combustor liner. The piston ring includes at least one arcuate seal
ring segment.
[0010] In accordance with another exemplary embodiment, disclosed
is a gas turbine including a combustor liner, a stage one nozzle
disposed downstream of the combustor liner and a piston seal
assembly defined at an interface of the combustor liner and the
stage one nozzle to seal a gas turbine hot gas path flow. The
piston seal assembly including a piston ring seal housing and a
piston ring. The piston ring seal housing has defined therein a
cavity. The piston ring is disposed within the cavity of the piston
ring seal housing and circumferentially about the combustor liner.
The piston ring is responsive to a regulated pressure to secure
sealing engagement of the piston ring and an outer surface of the
combustor liner. The piston ring includes at least one arcuate seal
ring segment.
[0011] In accordance with yet another exemplary embodiment,
disclosed is a gas turbine system including a compressor section, a
combustor section, a turbine section and a piston seal assembly.
The combustor section is coupled to the compressor section and
comprising an annular combustor liner defining an annular
combustion chamber coaxial with a longitudinal axis. The turbine
section is coupled to the combustor section and comprising a stage
one turbine nozzle positioned at the downstream end of the annular
combustor liner. The piston seal assembly is defined at an
interface of the annular combustor liner and the stage one nozzle
to seal a gas turbine hot gas path flow. The piston seal assembly
including a piston ring seal housing and a piston ring. The piston
ring seal housing has defined therein a cavity. The piston ring is
disposed within the cavity of the piston ring seal housing and
circumferentially about the combustor liner. The piston ring is
responsive to a regulated pressure to secure sealing engagement of
the piston ring and an outer surface of the combustor liner, the
piston ring including at least one arcuate seal ring segment.
[0012] Other objects and advantages of the present disclosure will
become apparent upon reading the following detailed description and
the appended claims with reference to the accompanying drawings.
These and other features and improvements of the present
application will become apparent to one of ordinary skill in the
art upon review of the following detailed description when taken in
conjunction with the several drawings and the appended claims.
DRAWINGS
[0013] These and other features, aspects, and advantages of the
present disclosure will become better understood when the following
detailed description is read with reference to the accompanying
drawings in which like characters represent like parts throughout
the drawings, wherein:
[0014] FIG. 1 is a cross sectional illustration of an aviation gas
turbine engine, in accordance with one or more embodiments shown or
described herein;
[0015] FIG. 2 is a cross sectional view of a portion of the engine
of FIG. 1, including a pressure regulated piston seal assembly, in
accordance with one or more embodiments shown or described
herein;
[0016] FIG. 3 is an enlarged cross sectional view of a portion of a
known piston ring seal housing, having a known piston ring disposed
therein, and under the influence of a high-pressure flow;
[0017] FIG. 4 is an enlarged cross sectional view of the piston
ring seal housing of FIG. 3, indicating rotation (tilt) of the
piston ring seal housing relative to the piston ring at a takeoff
condition;
[0018] FIG. 5 is a schematic view of an embodiment of a pressure
regulated piston seal assembly for a combustor liner, wherein the
piston ring seal housing is rotated relative to the piston ring and
wherein one or more sectional through-slots are located on an
upstream portion of the piston ring seal housing, in accordance
with one or more embodiments shown or described herein;
[0019] FIG. 6 is a schematic view of the pressure regulated piston
seal assembly of FIG. 5, wherein the piston ring is rotated
relative to the piston ring seal housing and wherein one or more
sectional through-slots are located on an upstream portion of the
piston ring seal housing, in accordance with one or more
embodiments shown or described herein;
[0020] FIG. 7 is a schematic view of the pressure regulated piston
seal assembly of FIG. 5, wherein the piston ring seal housing is
rotated relative to the piston ring and wherein one or more
sectional through-slots are located on a downstream portion of the
piston ring seal housing, in accordance with one or more
embodiments shown or described herein;
[0021] FIG. 8 is a schematic view of the pressure regulated piston
seal assembly of FIG. 5, wherein the piston ring is rotated
relative to the piston ring seal housing and wherein one or more
sectional through-slots are located on an outermost radial face of
the piston ring seal housing, in accordance with one or more
embodiments shown or described herein;
[0022] FIG. 9 is a schematic view of a portion of the pressure
regulated piston seal assembly of FIG. 5, wherein the piston ring
seal housing is rotated relative to the piston ring, in accordance
with one or more embodiments shown or described herein;
[0023] FIG. 10 is a schematic view of a portion of the pressure
regulated piston seal assembly of FIG. 6, wherein the piston ring
is rotated relative to the piston ring seal housing, in accordance
with one or more embodiments shown or described herein;
[0024] FIG. 11 is a schematic view of another embodiment of a
pressure regulated piston seal assembly for a combustor liner,
wherein the piston ring seal housing is rotated relative to the
piston ring, in accordance with one or more embodiments shown or
described herein;
[0025] FIG. 12 is a schematic view of the pressure regulated piston
seal assembly of FIG. 11, wherein the piston ring is rotated
relative to the piston ring seal housing, in accordance with one or
more embodiments shown or described herein;
[0026] FIG. 13 is a schematic view of a portion of the pressure
regulated piston seal assembly of FIG. 11, wherein the piston ring
seal housing is rotated relative to the piston ring, in accordance
with one or more embodiments shown or described herein;
[0027] FIG. 14 is a schematic view of a portion of the pressure
regulated piston seal assembly of FIG. 12, wherein the piston ring
is rotated relative to the piston ring seal housing, in accordance
with one or more embodiments shown or described herein;
[0028] FIG. 15 is a schematic view of an embodiment of a pressure
regulated piston seal assembly for a combustor liner, wherein the
piston ring seal housing is rotated relative to the piston ring, in
accordance with one or more embodiments shown or described
herein;
[0029] FIG. 16 is a schematic view of the pressure regulated piston
seal assembly of FIG. 15, wherein the piston ring is rotated
relative to the piston ring seal housing, in accordance with one or
more embodiments shown or described herein;
[0030] FIG. 17 is a schematic view of a piston ring of FIGS. 15 and
16, in accordance with one or more embodiments shown or described
herein;
[0031] FIG. 18 is a schematic view of another embodiment of the
piston ring of FIGS. 15 and 16, in accordance with one or more
embodiments shown or described herein;
[0032] FIG. 19 is a schematic view of an embodiment of a pressure
regulated piston seal assembly for a combustor liner, wherein the
piston ring seal housing is rotated relative to the piston ring, in
accordance with one or more embodiments shown or described
herein;
[0033] FIG. 20 is a schematic view of the pressure regulated piston
seal assembly of FIG. 19, wherein the piston ring is rotated
relative to the piston ring seal housing, in accordance with one or
more embodiments shown or described herein;
[0034] FIG. 21 is a schematic view of a portion of the pressure
regulated piston seal assembly of FIG. 20, wherein the piston ring
is rotated relative to the piston ring seal housing, in accordance
with one or more embodiments shown or described herein;
[0035] FIG. 22 is a schematic view of a portion of the pressure
regulated piston seal assembly of FIG. 20, wherein the piston ring
is rotated relative to the piston ring seal housing, in accordance
with one or more embodiments shown or described herein;
[0036] FIG. 23 is a schematic view of another embodiment of a
pressure regulated piston seal assembly for a combustor liner,
wherein the piston ring seal housing is rotated relative to the
piston ring, in accordance with one or more embodiments shown or
described herein;
[0037] FIG. 24 is a schematic view of the pressure regulated piston
seal assembly of FIG. 23, wherein the ring is rotated relative to
the piston ring seal housing, in accordance with one or more
embodiments shown or described herein;
[0038] FIG. 25 is a schematic view of a portion of the pressure
regulated piston seal assembly of FIG. 24, wherein the piston ring
is rotated relative to the piston ring seal housing, in accordance
with one or more embodiments shown or described herein;
[0039] FIG. 26 is a schematic view of a portion of the pressure
regulated piston seal assembly of FIG. 24, wherein the piston ring
is rotated relative to the piston ring seal housing, in accordance
with one or more embodiments shown or described herein;
[0040] FIG. 27 is a schematic view of an embodiment of a pressure
regulated piston seal assembly for a combustor liner, wherein the
piston ring seal housing is rotated relative to the piston ring, in
accordance with one or more embodiments shown or described
herein;
[0041] FIG. 28 is a schematic view of the pressure regulated piston
seal assembly of FIG. 27, wherein the ring is rotated relative to
the piston ring seal housing, in accordance with one or more
embodiments shown or described herein;
[0042] FIG. 29 is a schematic view of a portion of the pressure
regulated piston seal assembly of FIG. 28, wherein the piston ring
is rotated relative to the piston ring seal housing, in accordance
with one or more embodiments shown or described herein;
[0043] FIG. 30 is a schematic view of an embodiment of a pressure
regulated piston seal assembly for a combustor liner, in accordance
with one or more embodiments shown or described herein;
[0044] FIG. 31 is an enlarged cross-sectional view of a portion of
the pressure regulated piston seal assembly of FIG. 30, along
section 31-31, in accordance with one or more embodiments shown or
described herein;
[0045] FIG. 32 is a schematic view of the pressure regulated piston
seal assembly of FIG. 30, wherein the piston ring seal housing is
rotated relative to the piston ring, in accordance with one or more
embodiments shown or described herein; and
[0046] FIG. 33 is a schematic view of the pressure regulated piston
seal assembly of FIG. 30, wherein the piston ring is rotated
relative to the piston ring seal housing, in accordance with one or
more embodiments shown or described herein.
[0047] Unless otherwise indicated, the drawings provided herein are
meant to illustrate features of embodiments of this disclosure.
These features are believed to be applicable in a wide variety of
systems comprising one or more embodiments of this disclosure. As
such, the drawings are not meant to include all conventional
features known by those of ordinary skill in the art to be required
for the practice of the embodiments disclosed herein.
[0048] It is noted that the drawings as presented herein are not
necessarily to scale. The drawings are intended to depict only
typical aspects of the disclosed embodiments, and therefore should
not be considered as limiting the scope of the disclosure. In the
drawings, like numbering represents like elements between the
drawings.
DETAILED DESCRIPTION
[0049] In the following specification and the claims, reference
will be made to a number of terms, which shall be defined to have
the following meanings.
[0050] The singular forms "a", "an", and "the" include plural
references unless the context clearly dictates otherwise.
[0051] Approximating language, as used herein throughout the
specification and claims, is applied to modify any quantitative
representation that could permissibly vary without resulting in a
change in the basic function to which it is related. Accordingly, a
value modified by a term or terms, such as "about",
"approximately", and "substantially", are not to be limited to the
precise value specified. In at least some instances, the
approximating language may correspond to the precision of an
instrument for measuring the value. Here and throughout the
specification and claims, range limitations are combined and
interchanged. Such ranges are identified and include all the
sub-ranges contained therein unless context or language indicates
otherwise.
[0052] Referring now to the drawings wherein like numerals
correspond to like elements throughout, attention is directed
initially to FIG. 1 which depicts in diagrammatic form an exemplary
gas turbine engine 10 utilized with aircraft having a longitudinal
or axial centerline axis 12 therethrough for reference purposes.
Engine 10 preferably includes a core gas turbine engine generally
identified by numeral 14 and a fan section 16 positioned upstream
thereof. Core engine 14 typically includes a generally tubular
outer casing 18 that defines an annular inlet 20. Outer casing 18
further encloses and supports a booster compressor 22 for raising
the pressure of the air that enters core engine 14 to a first
pressure level. A high pressure, multi-stage, axial-flow compressor
24 receives pressurized air from booster 22 and further increases
the pressure of the air. The pressurized air flows to a combustor
26, where fuel is injected into the pressurized air stream to raise
the temperature and energy level of the pressurized air. The high
energy combustion products flow from combustor 26 to a first
(high-pressure) turbine 28 for driving high-pressure compressor 24
through a first (high-pressure) drive shaft (not shown), and then
to a second (low pressure) turbine 32 for driving booster
compressor 22 and fan section 16 through a second (low pressure)
drive shaft (not shown) that is coaxial with first drive shaft.
After driving each of turbines 28 and 32, the combustion products
leave core engine 14 through an exhaust nozzle 36.
[0053] Fan section 16 includes a rotatable, axial-flow fan rotor 38
and a plurality of fan rotor blades 44 that are surrounded by an
annular fan casing 40. It will be appreciated that fan casing 40 is
supported from core engine 14 by a plurality of substantially
radially-extending, circumferentially-spaced outlet guide vanes 42.
In this way, fan casing 40 encloses fan rotor 38 and the plurality
of fan rotor blades 44.
[0054] From a flow standpoint, it will be appreciated that an
initial airflow, represented by arrow 50, enters gas turbine engine
10 through an inlet 52. Airflow 50 passes through fan blades 44 and
splits into a first compressed airflow (represented by arrow 54)
that moves through the fan casing 40 and a second compressed
airflow (represented by arrow 56) which enters booster compressor
22. The pressure of second compressed airflow 56 is increased and
enters high-pressure compressor 24, as represented by arrow 57. A
high-pressure (P.sub.high) compressor airflow 58 exiting the
upstream compressor 24 flows in a downstream direction towards the
combustor 26, as a high-pressure (P.sub.high) airflow 59. After
mixing with fuel and being combusted in the combustor 26,
combustion products 61 exit combustor 26 and flow through first
turbine 28. The combustion products 61 then flow through second
turbine 32 and exit exhaust nozzle 36 to provide thrust for gas
turbine engine 10.
[0055] Referring now to FIG. 2, illustrated is an enlargement of a
portion of the gas turbine engine 10, as indicated by dashed line
in FIG. 1. The combustor 26 includes an annular combustion chamber
62 that is coaxial with longitudinal axis 12 (FIG. 1), as well as
an inlet (generally illustrated at 64) and an outlet 66. The
combustion chamber 62 is housed within engine outer casing 18 (FIG.
1) and defined by an annular combustor liner 68, and more
specifically, an annular combustor outer liner 69 and a
radially-inwardly positioned annular combustor inner liner 70.
Liners 69 and 70 may include those manufactured from and in a
process for CMC (Ceramic Matrix Composite). As noted above,
combustor 26 receives an annular stream of pressurized air, and
more particularly, the high-pressure (P.sub.high) airflow 59 via a
high-pressure compressor discharge outlet (not shown). This
high-pressure (P.sub.high) airflow 59 flows into the combustor 26,
where fuel is also injected from a fuel nozzle (not shown) and
mixes with the high-pressure (P.sub.high) airflow 59 from the
compressor to form a fuel-air mixture that is provided to the
combustion chamber 62 for combustion. Ignition of the fuel-air
mixture is accomplished by a suitable igniter 72, and the resulting
combustion gases 61 flow in an axial direction toward and into an
annular, stage one turbine nozzle 78 positioned at the downstream
end of the annular combustor outer liner 69 and the annular
combustor inner liner 70. The stage one turbine nozzle 78 is
defined by an annular flow channel that includes a plurality of
radially-extending, circularly-spaced nozzle vanes 80 that turn the
combustion gases 61 so that they flow angularly and impinge upon
the first stage turbine blades of first turbine 28. The stage one
turbine nozzle 78 includes a pair of flanges 82 and 84 to which the
downstream end of the annular combustor outer liner 69 and the
annular combustor inner liner 70, respectively, are mounted. As
illustrated, a portion of the high-pressure (P.sub.high) compressor
airflow 58 flows from the compressor 24 (FIG. 1) in a downstream
direction outside of the combustion liner 68, referred to herein as
a high-pressure (P.sub.high) bypass airflow 60. This high-pressure
(P.sub.high) bypass airflow 60 is intentionally meant to bypass the
combustor 26 (FIG. 1) to feed a piston assembly 86 (as indicated by
dotted lines in FIG. 2 and presently described) and the annular,
first stage turbine nozzle 78 and provide blade cooling. More
particularly, as best illustrated in FIG. 2, a portion of the
high-pressure (P.sub.high) bypass airflow 60 flows as an inner flow
60a and acts on the piston seal assembly 86. In addition, another
portion of the high-pressure (P.sub.high) bypass airflow 60 flows
as an outer flow 60b, radially outward the flow 60a, around the
piston seal assembly 86 to feed the turbine cooling.
[0056] The seal assembly 86 is comprised of integrated pressurized
piston seals for enhanced sealing of the annular combustor outer
liner 69, the annular combustor inner liner 70 and the downstream
stage one turbine nozzle 78, is disclosed and described herein.
[0057] Referring now to FIGS. 3 and 4, illustrated is an enlarged
portion of a known gas turbine engine, each labeled prior art,
generally similar to gas turbine 10, and including a combustor
liner 88, a stage one turbine nozzle 90, and a seal 86 formed
therebetween and in sealing engagement with the combustor liner 88
and the stage one turbine nozzle 90. More particularly, illustrated
in FIG. 3 is a seal assembly 86 during a first flight condition 92
and illustrated in FIG. 4 is the seal assembly 86 during a second
flight condition 94, such as during a takeoff condition. It should
be understood that only an outer radial portion of the seal
assembly is illustrated and described. As illustrated, the
pressurized seal assembly 86 is defined by a portion of the stage
one turbine nozzle 90 and more particularly a flange 96, forming a
hanger-type piston ring seal housing 98. The piston ring seal
housing 98 defining a cavity 102 therebetween a front wall 104 of
the piston ring seal housing 98 and a back wall 106 of the piston
ring seal housing 98. A piston ring 100 is disposed therein the
cavity 102 and confined, both radially and axially, by the
hanger-type piston ring seal housing 98. The piston ring seal
housing 98, and more particularly the cavity 102, includes an axial
dimension "a" that is greater than an axial dimension "b" of the
piston ring 100. The piston ring 100 is generally configured of a
single seal segment, or multiple arcuate seal segments, and more
particularly as a 360-degree circular ring that encloses an outer
surface 108 of the combustor liner 88 and defines the seal 86 at
the downstream end of the combustor liner 88. As used herein, the
term "arcuate" may refer to a member, component, part, etc. having
a curved or partially curved shape.
[0058] During the first flight condition 92, a high-pressure
(P.sub.high) bypass airflow 110 of compressor air, similar to the
high-pressure (No) bypass airflow 60a of FIG. 2, pushes the piston
ring 100 against the back wall 106 of the piston ring seal housing
98. As illustrated, in that the piston ring seal housing 98 is
wider than the piston ring 100, a front gap 112 is formed between
the front wall 104 of the piston ring seal housing 98 and the
piston ring 100 when the piston ring 100 is pushed against the back
wall 106 of the piston ring seal housing 98. As the high-pressure
(No) bypass airflow 110 enters the piston ring seal housing 98 via
the front gap 112 it results in a pressure drop cross the piston
ring 100 in the radial direction. This radial pressure drop on the
piston ring 100 assures the seal of a leakage path 116, as
indicated by dashed line, between the piston ring 100 and the outer
surface 108 of the combustor liner 88. In addition, the
high-pressure (No) bypass airflow 110 pushes the piston ring 100
against the back wall 106 sealing the contact between the piston
ring 100 and the back wall 106 of the piston ring seal housing 98,
assuring the seal of a leakage path 114, as indicated by dashed
line.
[0059] During the second flight condition 94, such as during a
takeoff stage of operation, the piston ring seal housing 98 is
rotated relative to the longitudinal axis 12 (FIG. 1), and more
particularly the combustor liner 88, due to high thermal gradient
across the piston ring seal housing 98 in radial direction. In the
illustrated embodiment of FIG. 4, the piston ring seal housing 98
is rotated at a degree "0", wherein 0.noteq.0 degrees. The rotation
of the piston ring seal housing 98 results in a blockage of the gap
112 resulting in a lack of pressure drop across the piston ring 100
in the radial direction. The main reason for piston seal failure is
due to this blockage of the gap 112 between the piston ring 100 and
the front wall 104 of the piston ring seal housing 98, resulting in
the high-pressure (P.sub.high) bypass airflow 110 not being present
on the top of the piston ring 100 and lack of sufficient
radially-inward force to engage the piston ring 100 on the outer
surface 108 of the combustor liner 88. As illustrated, the rotation
of the piston ring seal housing 98 and resulting closure of the gap
112 will result in the piston ring 100 no longer engaging with
combustor liner 88 causing the seal to fail. In an embodiment,
where the tilt is minimal, the piston ring 100 will not block the
gap 112 and the high-pressure (P.sub.high) bypass airflow 110 from
entering an upper portion 118 of the cavity 102.
[0060] Referring now to FIGS. 5-31, illustrated are embodiments of
a pressure-regulated piston seal assembly according to this
disclosure. It is again noted, that like numbers represent like
elements throughout the embodiments. As described, the seal
assembly will provide for sealing engagement during all flight
conditions, and in particular, during takeoff flight conditions
whereby the piston ring seal housing and the piston ring are
rotated, relative to one another. In the illustrated embodiments,
each pressure-regulated seal assembly configuration is illustrated
wherein the piston ring seal housing is rotated, relative to the
longitudinal axis, and wherein the piston ring is rotated relative
to the longitudinal axis.
[0061] Furthermore, in the embodiments of FIGS. 5-31, the
pressure-regulated piston seal assembly is defined by a portion of
the stage one turbine nozzle 78, the hanger-type piston ring seal
housing 122 and the piston ring 124. The piston ring seal housing
122 defining a cavity 132 therebetween a front wall 126 of the
piston ring seal housing 122 and a back wall 128 of the piston ring
seal housing 122. The piston ring 124 is disposed therein the
cavity 132 and confined, both radially and axially, by the
hanger-type piston ring seal housing 122. As previously described
with regard to FIGS. 3 and 4, in an embodiment the piston ring seal
housing 122, and more particularly the cavity 132, includes an
axial dimension "a" that is greater than an axial dimension "b" of
the piston ring 124. In some embodiments, the cavity 132 has an
axial dimension "a" of approximately 1 millimeter to approximately
10 millimeters. In some embodiments, the piston ring 124 has an
axial dimension "b" of approximately 1 millimeter to approximately
10 millimeters. The piston ring 124 is generally configured as a
single component or segment, or multiple arcuate segments, and more
particularly as a 360-degree circular ring that encloses an outer
surface 136 of the combustor liner 68 and defines a seal at the
downstream end of the combustor liner 68. As used herein, the term
"arcuate" may refer to a member, component, part, etc. having a
curved or partially curved shape. As disclosed herein, in a
preferred embodiment the combustor liner 68 is fabricated with
ceramic matrix composites (CMCs).
[0062] As previously indicated, in advanced gas path (AGP) heat
transfer designs for gas turbine engines, the arcuate components,
and in particular the shrouds, nozzles, and the like, are
fabricated with ceramic matrix composites (CMCs). Similar to seal
assemblies used in conventional designs, the AGP components utilize
static seals of various types of construction (e.g. solid,
laminate, shaped, etc.). The static seals are typically made of a
high temperature metal material, such as nickel alloy. The CMC
material that forms many of the AGP components has a lower
co-efficient of thermal expansion (CTE) compared to the static
seals formed of the high temperature metal.
[0063] Referring specifically to FIGS. 5-8, illustrated are
embodiments of a pressure-regulated seal assembly, generally
referenced 120. The seal assembly 120 is generally defined by a
portion of the stage one nozzle 78, as previously introduced in
FIG. 2, a piston ring seal housing 122 and a piston ring 124.
Similar to the previously described known art of FIGS. 3 and 4, the
piston ring seal housing 122 is wider than the piston ring 124, and
defines a front gap (not shown) between a front wall 126 of the
piston ring seal housing 122 and the piston ring 124 when the
piston ring 124 is pushed against a back wall 128 of the piston
ring seal housing 122.
[0064] During a first flight condition (not shown), as a portion of
the high-pressure (P.sub.high) compressor airflow 58 from the
compressor 24 (FIG. 1), and more particularly an inner flow of a
high-pressure (P.sub.high) bypass airflow 60a enters the piston
ring seal housing 122 via the gap, generally similar to gap 112 of
FIGS. 3 and 4, it results in a pressure drop cross the piston ring
124 in the radial direction, thereby engaging the piston ring 124
with the combustor liner 68 and the piston ring seal housing 122,
as previously described. In contrast, during a second operating
condition, as best illustrated in FIG. 5, such as during a takeoff
condition, the piston ring seal housing 122 is rotated at an angle
".beta..sub.1" relative to the piston ring 124, and more
specifically the longitudinal axis 12, due to the high thermal
gradient across the piston ring seal housing 98 in the radial
direction. In an embodiment, .beta..sub.1.noteq.0 degrees. In this
instance, the piston ring 124 blocks the gap (not shown) and thus
the pathway allowing the inner flow of the high-pressure
(P.sub.high) bypass airflow 60a to enter an upper portion 130 of a
cavity 132 defined between the front wall 126 and the back wall 128
of the piston ring seal housing 122. The piston ring 124 no longer
engages with combustor liner 68 due to the lack of the inner flow
of the high-pressure (P.sub.high) bypass airflow 60a into the
cavity 132, and the seal fails. To address such rotation of the
piston ring seal housing 122 during the second flight condition,
one or more sectional through-slots 134 are provided in the front
wall 126 of the piston ring seal housing 122, following a
360-degree circular contour. The one or more sectional
through-slots 134 may be configured a single continuous
through-slot 134, as best illustrated in FIG. 9 or as a plurality
of through-slots 134 as best illustrated in FIG. 10. The one or
more sectional through-slots 134 provide an opening to the upper
portion 130 of the cavity 132. The one or more sectional
through-slots 134 guide the inner flow of the high-pressure
(P.sub.high) bypass airflow 60a directly to a top of the piston
ring 124 and assure engagement of the piston ring 124 with an outer
surface 136 of the combustor liner 68 regardless of the degree of
rotation of the piston ring seal housing 122. In an embodiment, the
one or more sectional through-slots 134 may be manufactured
together with the front wall 104 of the piston ring seal housing
122 and no further changes are required for the parts assembly.
[0065] Referring now to FIG. 6, in an embodiment the piston ring
124 may rotate relative to the piston ring seal housing 122, and
more particularly the longitudinal axis 12, such as during the
second operating condition. In this particular embodiment, the
piston ring 124 is rotated at an angle "a" relative to the piston
ring seal housing 122, and more specifically the longitudinal axis
12, due to thermal stresses. In an embodiment .alpha..noteq.0
degrees. Similar to the embodiment of FIG. 6, in this instance, the
piston ring 124 will block the pathway of the inner flow of the
high-pressure (P.sub.high) bypass airflow 60a from entering an
upper portion 130 of the cavity 132 defined between the front wall
126 and the back wall 128 of the piston ring seal housing 122. The
lack of inner flow of the high-pressure (P.sub.high) bypass airflow
60a into the cavity 132 causes the piston ring 124 to no longer
engage with combustor liner 68 and the seal fails. To address such
rotation of the piston ring 124 during the second flight condition,
one or more sectional through-slots 134 provided in the front wall
126 of the piston ring seal housing 122, similar to the embodiment
of FIG. 5, are provided following a 360-degree circular contour and
configured as previously described. The one or more sectional
through-slots 134 provide an opening to the upper portion 130 of
the cavity 132. The one or more sectional through-slots 134 guide
the inner flow of the high-pressure (P.sub.high) bypass airflow 60a
directly to the top of the piston ring 124 and assure engagement of
the piston ring 124 with the outer surface 136 of the combustor
liner 68 regardless of the degree of rotation of the piston ring
124.
[0066] Referring now to FIGS. 7 and 8, in an embodiment, as
previously described, a portion of the high-pressure (P.sub.high)
bypass airflow 60 flows through a window (not shown) in a support
leg that supports the piston ring seal housing 122. As a result, an
outer flow of the high-pressure (P.sub.high) bypass airflow 60b
passes through one or more sectional through-slots 134 formed on
the aft face, or back wall 128, of the piston ring seal housing
122, as best illustrated in FIG. 7. In the embodiment of FIG. 8,
the outer flow of the high-pressure (P.sub.high) bypass airflow 60b
passes through one or more sectional through-slots 134 formed on an
outermost radial face 129 of the piston ring seal housing 122. It
is noted that the configurations of FIGS. 7 and 8 may be
advantageous over the embodiment of FIGS. 5 and 6 where there is
limited room for the one or more sectional through-slots 134 on the
forward face 126.
[0067] Referring now to FIGS. 11-18, illustrated is a second
embodiment of a pressure-regulated seal assembly, generally
referenced 140. The seal assembly 140 is generally defined by a
portion of the stage one nozzle 78, as previously introduced in
FIG. 2, a piston ring seal housing 122 and a piston ring 124.
Similar to the previously described known art of FIGS. 3 and 4, the
piston ring seal housing 122 is wider than the piston ring 124, and
defines a front gap (not shown) between a front wall 126 of the
piston ring seal housing 122 and the piston ring 124 when the
piston ring 124 is pushed against a back wall 128 of the piston
ring seal housing 122.
[0068] During a first flight condition (not shown), as the inner
flow of the high-pressure (P.sub.high) bypass airflow 60a enters
the piston ring seal housing 122 via the gap, it results in a
pressure drop cross the piston ring 124 in the radial direction,
thereby engaging the piston ring 124 with the combustor liner 68
and the piston ring seal housing 122, as previously described. In
contrast, during a second operating condition, as best illustrated
in FIGS. 11, 13 and 15, such as during a takeoff condition, the
piston ring seal housing 122 is rotated at an angle relative to the
piston ring 124, and more specifically the longitudinal axis 12,
due to the high thermal gradient across the piston ring seal
housing 98 in the radial direction. As best illustrated in FIG. 11,
the piston ring seal housing 122 is rotated at an angle
".beta..sub.2" relative to the piston ring 124, and more
specifically the longitudinal axis 12. In this particular
embodiment, it is noted the piston ring seal housing 122 is rotated
in generally an opposed direction to that of the previous
embodiment. In an embodiment, .beta..sub.2.noteq.0 degrees. In the
embodiments of FIGS. 13 and 15, the piston ring seal housing 122 is
rotated at an angle ".beta..sub.1" relative to the piston ring 124,
and more specifically the longitudinal axis 12. It is additionally
noted the piston ring seal housing 122 may have a direction of
rotation of either .beta..sub.1 or .beta..sub.2 in each of the
disclosed embodiments.
[0069] As illustrated, the piston ring 124 will block the inner
flow of the high-pressure (P.sub.high) bypass airflow 60a from the
compressor 24 (FIG. 1) from entering an upper portion 130 of a
cavity 132 defined between the front wall 126 and the back wall 128
of the piston ring seal housing 122. The lack of inner flow of the
high-pressure (P.sub.high) bypass airflow 60a into the cavity 132
causes the piston ring 124 to no longer engage with combustor liner
68 and the seal fails. To address such rotation of the piston ring
seal housing 122 during the second flight condition, a plurality of
local bumps 142 are provided on the front wall 126 of the piston
ring seal housing 122, as best illustrated in FIGS. 11-14, or on an
upstream surface 144 of the piston ring 124 as best illustrated in
FIGS. 15-18. As illustrated in FIGS. 11-14, the plurality of local
bumps 142 extend or protrude therefrom the front wall 126 of the
piston ring seal housing 122. The plurality of local bumps 142
provides a flow opening therebetween to the upper portion 130 of
the cavity 132. The plurality of local bumps 142 thus guide the
inner flow of the high-pressure (P.sub.high) bypass airflow 60a
directly to top of the piston ring 124 and assure engagement of the
piston ring 124 with an outer surface 136 of the combustor liner 68
regardless of the degree of rotation of the piston ring seal
housing 122. In an embodiment, the one or more local bumps 142 may
be configured in a plurality of columns and rows, as best
illustrated in FIGS. 13 and 17. In an embodiment, the one or more
local bumps 142 may be configured in a plurality of columns and
offset rows, as best illustrated in FIGS. 14 and 18. In an
embodiment, the plurality of local bumps 142 may be manufactured
together with the front wall 104 of the piston ring seal housing
122 and/or the upstream surface 144 of the piston ring 124 and no
further changes are required for the parts assembly.
[0070] Referring more particularly to FIGS. 12, 14 and 16, in an
embodiment the piston ring 124 may rotate relative to the piston
ring seal housing 122, and more particularly the longitudinal axis
12, such as during the second operating condition. In this
particular embodiment, the piston ring 124 is rotated at an angle
"a" relative to the piston ring seal housing 122, and more
specifically the longitudinal axis 12, due to thermal stresses. As
previously indicated a 0 degrees. Similar to the embodiment of FIG.
6, in this instance, the piston ring 124 will block the inner flow
of the high-pressure (P.sub.high) bypass airflow 60a from entering
an upper portion 130 of a cavity 132 defined between the front wall
126 and the back wall 128 of the piston ring seal housing 122. The
lack of inner flow of the high-pressure (P.sub.high) bypass airflow
60a into the cavity 132 causes the piston ring 124 to no longer
engage with the combustor liner 68 and the seal fails. To address
such rotation of the piston ring 124 during the second flight
condition, the plurality of local bumps 142 are configured to
extend therefrom the upstream surface 144 of the piston ring 124.
The plurality of local bumps 142 provide a flow opening to the
upper portion 130 of the cavity 132. The plurality of local bumps
142 guide inner flow of the high-pressure (P.sub.high) bypass
airflow 60a directly to top of the piston ring 124 and assure
engagement of the piston ring 124 with an outer surface 136 of the
combustor liner 68 regardless of the degree of rotation of the
piston ring 124.
[0071] Referring now to FIGS. 19-29, illustrated is a third
embodiment of a pressure-regulated seal assembly, generally
referenced 150. The seal assembly 150 is generally defined by a
portion of the stage one nozzle 78, as previously introduced in
FIG. 2, a piston ring seal housing 122 and a piston ring 124.
Similar to the previously described known art of FIGS. 3 and 4, the
piston ring seal housing 122 is wider than the piston ring 124, and
defines a front gap (not shown) between a front wall 126 of the
piston ring seal housing 122 and the piston ring 124 during when
the piston ring 124 is pushed against a back wall 128 of the piston
ring seal housing 122.
[0072] During a first flight condition (not shown), as the inner
flow of the high-pressure (P.sub.high) bypass airflow 60a enters
the piston ring seal housing 122 via the gap, it results in a
pressure drop cross the piston ring 124 in the radial direction,
thereby engaging the piston ring 124 with the combustor liner 68
and the piston ring seal housing 122, as previously described. In
contrast, during a second operating condition, as best illustrated
in FIGS. 19, 23 and 27, such as during a takeoff condition, the
piston ring seal housing 122 is rotated at an angle "Pi" relative
to the piston ring 124, and more specifically the longitudinal axis
12, due to the high thermal gradient across the piston ring seal
housing 98 in the radial direction. As previously indicated, in an
embodiment .beta..sub.1.noteq.0 degrees. In this instance, the
piston ring 124 will block the inner flow of the high-pressure
(P.sub.high) bypass airflow 60a from entering an upper portion 130
of a cavity 132 defined between the front wall 126 and the back
wall 128 of the piston ring seal housing 122. The lack of inner
flow of the high-pressure (P.sub.high) bypass airflow 60a into the
cavity 132 causes the piston ring 124 to no longer engage with
piston seal liner 124 and the seal fails. In this particular
embodiment, to address such rotation of the piston ring seal
housing 122 during the second flight condition, a plurality of
channels 152 are provided on the front wall 126 of the piston ring
seal housing 122, as best illustrated in FIGS. 19-22, on the
upstream surface 144 of the piston ring 124 as best illustrated in
FIGS. 23-26, or on both the front wall 126 of the piston ring seal
housing 122 and the upstream surface 144 of the piston ring 124 as
best illustrated in FIGS. 27-29. As illustrated in FIGS. 19-22, the
plurality of channels 152 are configured to extend into the surface
125 of the front wall 126 of the piston ring seal housing 122. The
plurality of channels 152 provides a plurality of high pressure gas
flow through conduits to the upper portion 130 of the cavity 132.
More specifically, the plurality of channels 152 guide the inner
flow of the high-pressure (P.sub.high) bypass airflow 60a directly
to top of the piston ring 124 and assure engagement of the piston
ring 124 with an outer surface 136 of the combustor liner 68
regardless of the degree of rotation of the piston ring seal
housing 122. In an embodiment, the plurality of channels 152 may be
manufactured together with the front wall 104 of the piston ring
seal housing 122 and no further changes are required for the parts
assembly.
[0073] Referring now to FIGS. 20-22, 24-26, 28 and 29, in an
embodiment the piston ring 124 may rotate relative to the piston
ring seal housing 122, and more particularly the longitudinal axis
12, such as during the second operating condition. In this
particular embodiment, the piston ring 124 is rotated at an angle
"a" relative to the piston ring seal housing 122, and more
specifically the longitudinal axis 12, due to thermal stresses. As
previously indicated, in an embodiment .alpha..noteq.0 degrees.
Similar to the embodiment of FIG. 6, in this instance, the piston
ring 124 will block the pathway of the inner flow of the
high-pressure (P.sub.high) bypass airflow 60a from the compressor
24 (FIG. 1) from entering an upper portion 130 of a cavity 132
defined between the front wall 126 and the back wall 128 of the
piston ring seal housing 122. The lack of inner flow of the
high-pressure (P.sub.high) bypass airflow 60a into the cavity 132
causes the piston ring 124 to no longer engage with the combustor
liner 68 and the seal fails. To address such rotation of the piston
ring 124 during the second flight condition, a plurality of
channels 152 are configured to extend into the upstream surface 144
of the piston ring 124. The plurality of channels 152 provides a
plurality of high pressure gas flow through conduits to the upper
portion 130 of the cavity 132. More specifically, the plurality of
channels 152 guide the inner flow of the high-pressure (P.sub.high)
bypass airflow 60a directly to top of the piston ring 124 and
assure engagement of the piston ring 124 with an outer surface 136
of the combustor liner 68 regardless of the degree of rotation of
the piston ring 124. In an embodiment, the plurality of channels
152 may be manufactured together with upstream surface 144 of the
piston ring 124 and no further changes are required for the parts
assembly. As best illustrated in FIG. 29, in an embodiment, the
plurality of channels 152 may be configured to extend both into the
front wall 126 of the piston ring seal housing 122 and into the
upstream surface 144 of the piston ring 124.
[0074] Referring now to FIGS. 30-33, illustrated is yet another
embodiment of a pressure-regulated seal assembly, generally
referenced 160. The seal assembly 160 is generally defined by a
portion of the stage one nozzle 78, as previously introduced in
FIG. 2, a piston ring seal housing 122 and a piston ring 124.
Similar to the previously described known art of FIGS. 3 and 4, the
piston ring seal housing 122 is wider than the piston ring 124, and
defines a front gap (not shown) between a front wall 126 of the
piston ring seal housing 122 and the piston ring 124 during when
the piston ring 124 is pushed against a back wall 128 of the piston
ring seal housing 122.
[0075] During a first flight condition, as best illustrated in FIG.
30, as the inner flow of the high-pressure (P.sub.high) bypass
airflow 60a from the compressor 24 (FIG. 1) enters the piston ring
seal housing 122 via the front gap 112, it results in a pressure
drop cross the piston ring 124 in the radial direction, thereby
engaging the piston ring 124 with the combustor liner 68 and the
piston ring seal housing 122, as previously described. In this
particular embodiment, a cockle spring 162 is provided having an
inner circumferential dimension "D" less than an outer
circumferential dimension "d" of the downstream end of the
combustor liner 26, as indicated in FIG. 31. The cockle spring 162
exerts a radially-inward force on the piston ring 124, as indicated
by arrows in FIG. 30, to enhance engagement of the piston ring 124
with the combustor liner 68.
[0076] In contrast, during a second operating condition, as best
illustrated in FIG. 32, such as during a takeoff condition, the
piston ring seal housing 122 is rotated at an angle ".beta..sub.1"
relative to the piston ring 124, and more specifically the
longitudinal axis 12, due to the high thermal gradient across the
piston ring seal housing 98 in the radial direction. As previously
indicated, .beta..sub.1.noteq.0 degrees. In this instance, the
piston ring 124 will block the inner flow of the high-pressure
(P.sub.high) bypass airflow 60a from entering an upper portion 130
of a cavity 132 defined between the front wall 126 and the back
wall 128 of the piston ring seal housing 122. The lack of inner
flow of the high-pressure (P.sub.high) bypass airflow 60a into the
cavity 132 causes the piston ring 124 to no longer engage with
piston seal liner 124 and the seal fails. In this particular
embodiment, in addition to the inclusion of one or more sectional
through-slots 134, the cockle spring 162 addresses such rotation of
the piston ring seal housing 122 during the second flight condition
by enhancing the radially-inward force on the piston ring 124, as
indicated by arrows in FIG. 32, to enhance engagement of the piston
ring 124 with the combustor liner 68.
[0077] Referring now to FIG. 33, in another embodiment the piston
ring 124 may rotate relative to the piston ring seal housing 122,
and more particularly the longitudinal axis 12, such as during the
second operating condition. In this particular embodiment, the
piston ring 124 is rotated at an angle "a" relative to the piston
ring seal housing 122, and more specifically the longitudinal axis
12, due to the thermal stresses. As previously indicated, in an
embodiment .alpha..noteq.0 degrees. Similar to the embodiment of
FIG. 32, in this instance, the rotation of the piston ring 124 will
cause it to block the inner flow of the high-pressure (P.sub.high)
bypass airflow 60a from entering an upper portion 130 of a cavity
132 defined between the front wall 126 and the back wall 128 of the
piston ring seal housing 122. The lack of inner flow of the
high-pressure (P.sub.high) bypass airflow 60a into the cavity 132
causes the piston ring 124 to no longer engage with the combustor
liner 68 and the seal fails. In this particular embodiment, to
address such rotation of the piston ring 124 during the second
flight condition, in addition to the inclusion of a plurality of
local bumps 142 on the front wall 126 of the piston ring seal
housing 122, the cockle spring 162 further addresses such rotation
of the piston ring 124. As previously described, the cockle spring
162 provides a radially-inward force on the piston ring 124, as
indicated by arrows in FIG. 33, to assure engagement of the piston
ring 124 with the combustor liner 68 regardless of the degree of
rotation of the piston ring 124.
[0078] As described herein, in known piston seal assemblies, the
main reason for piston seal failure is due to the blockage of a gap
that is present between the piston ring and the front wall of the
piston ring seal housing, so that an upstream high-pressure
(P.sub.high) flow exiting the compressor is not present on the top
of the piston ring and there is a lack of sufficient
radially-inward force to engage the piston ring on an outer surface
of the combustor liner. In an embodiment, where tilting of the
piston ring seal housing or piston ring is present, the piston ring
will block the gap and more particularly the high-pressure
(P.sub.high) bypass airflow exiting the compressor from entering an
upper portion of the cavity. The blocking of the gap results in a
lack of pressure drop across the piston ring in the radial
direction. As illustrated, the relative rotating of the piston ring
seal housing and closure of the gap will result in the piston ring
no longer engaging with combustor liner causing the seal to
fail.
[0079] Accordingly, disclosed is a pressure regulated piston seal
for sealing between a combustor liner and a downstream stage one
turbine nozzle. The amount of cooling flow through the seal plays
an important role in cooling the mechanical parts along the flow
path. Sufficient cooling flows are necessary to assure acceptable
lifetime of the seal, while too much cooling flow results in waste
of compressor air. Meanwhile the extra cooling flow through the
piston seal would affect the exit temperature of the combustor
liner and further lower the turbine efficiency.
[0080] Conventional piston seals typically applied on metal
combustor liners were not sensitive to thermal gradient, in that
the piston ring seal housing and combustor liner were both formed
of a metal and rotated simultaneously. CMC combustor liners
experience large relative rotation between the piston ring seal
housing and the combustor liner, and result in failure of the seal.
The seal assembly disclosed herein provides a solution for CMC
combustor liners and associated seal assemblies. The proposed seal
assembly assures the piston seal functions well in all flight
conditions and thus a controllable cooling flow through the
interface of the combustor liner and the stage one turbine nozzle.
The manufacturing of the proposed seal assembly does not increase
cost and the assembly procedure of the pressure regulated seal
disclosed herein is essentially the same as that of conventional
piston seals.
[0081] Exemplary embodiments of a pressure-regulated seal assembly
are described above in detail. The systems and methods are not
limited to the specific embodiments described herein, but rather,
operations of the methods and components of the systems may be
utilized independently and separately from other operations or
components described herein. For example, the systems, methods, and
apparatus described herein may have other industrial or consumer
applications and are not limited to practice as described herein.
Rather, one or more embodiments may be implemented and utilized in
connection with other industries.
[0082] Although specific features of various embodiments of the
disclosure may be shown in some drawings and not in others, this is
for convenience only. In accordance with the principles of the
disclosure, any feature of a drawing may be referenced and/or
claimed in combination with any feature of any other drawing.
[0083] This written description uses examples to disclose the
embodiments, including the best mode, and enables any person
skilled in the art to practice the embodiments, including making
and using any devices or systems and performing any incorporated
methods. The patentable scope of the disclosure is defined by the
claims, and may include other examples that occur to those skilled
in the art. Such other examples are intended to be within the scope
of the claims if they have structural elements that do not differ
from the literal language of the claims, or if they include
equivalent structural elements with insubstantial differences from
the literal language of the claims.
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