U.S. patent application number 17/142357 was filed with the patent office on 2022-07-07 for system for controlling blade clearances within a gas turbine engine.
The applicant listed for this patent is General Electric Company. Invention is credited to Bradley W. Fintel, Steven Douglas Johnson, Brandon Wayne Miller, Julius John Montgomery, Robert Proctor, Jeffrey Douglas Rambo.
Application Number | 20220213802 17/142357 |
Document ID | / |
Family ID | 1000005384782 |
Filed Date | 2022-07-07 |
United States Patent
Application |
20220213802 |
Kind Code |
A1 |
Johnson; Steven Douglas ; et
al. |
July 7, 2022 |
SYSTEM FOR CONTROLLING BLADE CLEARANCES WITHIN A GAS TURBINE
ENGINE
Abstract
A system for controlling blade clearances within a gas turbine
engine includes a rotor disk and a rotor blade coupled to the rotor
disk. Additionally, the system includes an outer turbine component
positioned outward of the rotor blade such that a clearance is
defined between the rotor blade and the outer turbine component.
Furthermore, the system includes a heat exchanger configured to
receive a flow of cooling air bled from the gas turbine engine and
transfer heat from the received flow of the cooling air to a flow
of coolant to generate cooled cooling air. Moreover, the system
includes a valve configured to control the flow of the coolant to
the heat exchanger. In this respect, the cooled cooling air is
supplied to at least one of the rotor disk or the rotor blade to
adjust the clearance between the rotor blade and the outer turbine
component.
Inventors: |
Johnson; Steven Douglas;
(Milford, OH) ; Montgomery; Julius John; (Mason,
OH) ; Miller; Brandon Wayne; (Liberty Township,
OH) ; Proctor; Robert; (Mason, OH) ; Fintel;
Bradley W.; (West Chester, OH) ; Rambo; Jeffrey
Douglas; (Mason, OH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Family ID: |
1000005384782 |
Appl. No.: |
17/142357 |
Filed: |
January 6, 2021 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F05D 2220/323 20130101;
F05D 2240/11 20130101; F05D 2260/30 20130101; F01D 11/24 20130101;
F05D 2240/55 20130101; F01D 25/14 20130101; F02C 6/08 20130101;
F01D 25/12 20130101; F05D 2260/20 20130101; F01D 11/18 20130101;
F02C 7/185 20130101 |
International
Class: |
F01D 11/24 20060101
F01D011/24; F01D 11/18 20060101 F01D011/18 |
Claims
1. A system for controlling blade clearances within a gas turbine
engine, the gas turbine engine defining an axial centerline and a
radial direction extending orthogonal to the axial centerline, the
system comprising: a rotor disk; a rotor blade coupled to the rotor
disk; an outer turbine component positioned outward of the rotor
blade in the radial direction such that a clearance is defined
between the rotor blade and the outer turbine component; a heat
exchanger configured to receive a flow of cooling air bled from the
gas turbine engine and transfer heat from the received flow of the
cooling air to a flow of coolant to generate cooled cooling air;
and a valve configured to control the flow of the coolant to the
heat exchanger, wherein the cooled cooling air is supplied to at
least one of the rotor disk or the rotor blade to adjust the
clearance between the rotor blade and the outer turbine
component.
2. The system of claim 1, further comprising: a shaft coupled to
the rotor disk such that rotation of the rotor disk and the rotor
blade rotates the shaft; a combustor positioned outward in the
radial direction from the shaft; and a conduit at least partially
positioned between the shaft and the combustor in the radial
direction such that the cooled cooling air flows through the
conduit to the at least one of the rotor disk or the rotor
blade.
3. The system of claim 2, further comprising: an inducer configured
to direct the cooled cooling air flowing through the conduit toward
the rotor disk.
4. The system of claim 3, wherein the inducer narrows as the
inducer extends from the conduit toward the rotor disk.
5. The system of claim 3, further comprising: a seal positioned
upstream of the rotor disk along the axial centerline relative to a
direction of flow through the gas turbine engine, wherein the
inducer directs the cooled cooling air such that the cooled cooling
air flows between the rotor disk and the seal.
6. The system of claim 5, wherein the seal corresponds to an outer
seal, the system further comprising: an inner seal positioned
inward along the radial direction relative to the outer seal such
that a gap is defined between the inner and outer seals through
which the cooled cooling air flows from the inducer toward the
rotor disk.
7. The system of claim 2, wherein the conduit includes a first
portion extending along the radial direction from the heat
exchanger and a second portion extending along the axial centerline
from the first portion toward the rotor disk.
8. The system of claim 7, wherein the first portion of the conduit
is positioned upstream of the combustor relative to a direction of
flow through the gas turbine engine.
9. The system of claim 7, wherein the heat exchanger is positioned
outward along the radial direction from the combustor.
10. The system of claim 2, further comprising: a compressor
discharge casing at least partially surrounding the combustor, the
compressor discharge casing defining a compressor discharge plenum
configured to supply compressed air to the combustor, wherein the
cooling air received by the heat exchanger is bled from the
compressor discharge plenum.
11. The system of claim 1, further comprising: a turbine case
coupled to the outer turbine components, wherein the cooled cooling
air is supplied to the turbine case to adjust the clearance between
the rotor blade and the outer turbine component.
12. The system of claim 1, further comprising: a bypass conduit
fluidly coupled to the valve such that the bypass conduit is
configured to permit at least a portion of the coolant to bypass
the heat exchanger.
13. The system of claim 1, wherein the cooled cooling air is
discharged into a hot gas path at least partially defined by the
rotor blade and the outer turbine component after being supplied to
the at least one of the rotor disk or the rotor blade.
14. The system of claim 1, wherein the coolant comprises
supercritical carbon dioxide.
15. The system of claim 1, wherein the outer turbine component
comprises a shroud or an outer rotating drum.
16. A system for controlling blade tip clearances within a gas
turbine engine, the gas turbine engine defining an axial centerline
and a radial direction extending orthogonal to the axial
centerline, the system comprising: an inner rotor configured to
rotate in a first direction; an inner rotor blade coupled to the
inner rotor; an outer rotating drum configured to rotate in a
second direction opposite of the first direction; an outer rotor
blade coupled to the outer rotating drum; a heat exchanger
configured to receive a flow of cooling air bled from the gas
turbine engine and transfer heat from the received flow of the
cooling air to a flow of coolant to generate cooled cooling air; a
first air valve configured to direct a first portion of the cooled
cooling air to the outer rotating drum and a second portion of the
cooled cooling air to cool the inner rotor; and a second air valve
configured to direct a first portion of the cooling air to the
outer rotating drum and a second portion of the cooling air to cool
the inner rotor, wherein the cooled cooling air is supplied to at
least one of the outer rotating drum or the inner rotor to adjust a
first clearance defined between the inner rotor blade and the outer
rotating drum and a second clearance between the outer rotor blade
and the inner rotor.
17. The system of claim 16, where the first portion of the cooled
cooling air is introduced to the outer rotating drum through an
angled nozzle such that a tangential component of a velocity of the
first portion of the cooled cooling air is in the second
direction.
18. The system of claim 16, where the second portion of the cooled
cooling air is introduced to the inner rotor through an angled
nozzle such that a tangential component of a velocity of the second
portion of the cooled cooling air is in the first direction.
19. The system of claim 16, where the first portion of the cooling
air is introduced to outer rotating drum through an angled nozzle
such that a tangential component of a velocity of the first portion
of the cooling air is in the first direction.
20. The system of claim 16, where the second portion of the cooling
air is introduced to inner rotor through an angled nozzle such that
a tangential component of a velocity of the second portion of the
cooling air is in the second direction.
Description
FIELD
[0001] The present disclosure generally pertains to gas turbine
engines and, more particularly, to a system for controlling blade
clearances within a gas turbine engine.
BACKGROUND
[0002] A gas turbine engine generally includes a compressor
section, a combustion section, and a turbine section. During
operation, the compressor section progressively increases the
pressure of air entering the engine and supplies this compressed
air to the combustion section. The compressed air and a fuel mix
within the combustion section and burn within a combustion chamber
to generate high-pressure and high-temperature combustion gases.
The combustion gases flow through a hot gas path defined by the
turbine section before exiting the engine. In this respect, the
turbine section converts energy from the combustion gases into
rotational energy. Specifically, the turbine section includes a
plurality of rotor blades, which extract kinetic energy and/or
thermal energy from the combustion gases flowing therethrough. The
extracted rotational energy is, in turn, used to rotate one or more
shafts, thereby driving the compressor section and/or a fan
assembly of the gas turbine engine
[0003] In general, it desirable to minimize the clearance between
the outer tips of the rotor blades and the adjacent shrouds or drum
to maximize the amount of energy extracted by the rotor blades.
However, the rotor blades expand and contract relative to the
shrouds/drum during thermal cycling of the engine. As such, the
clearance between the rotor blades and the shrouds/drum generally
decreases as the engine heats up. In this respect, when the
clearance between the blade tips and the shrouds/drum is minimized
during cold operation of the engine, the blade tips may contact the
shrouds/drum when the engine heats up. Conversely, when the
clearance between the blade tips and the shroud/drum is optimized
for hot operation, such clearance may be sufficiently large to
reduce the efficiency of the energy extraction during cold
operation of the engine.
[0004] Accordingly, an improved system for controlling blade
clearances within a gas turbine engine would be welcomed in the
technology.
BRIEF DESCRIPTION
[0005] Aspects and advantages of the invention will be set forth in
part in the following description, or may be obvious from the
description, or may be learned through practice of the
invention.
[0006] In one aspect, the present subject matter is directed to a
system for controlling blade clearances within a gas turbine
engine. The gas turbine engine defines an axial centerline and a
radial direction extending orthogonal to the axial centerline. The
system includes a rotor disk and a rotor blade coupled to the rotor
disk. Additionally, the system includes an outer turbine component
positioned outward of the rotor blade in the radial direction such
that a clearance is defined between the rotor blade and the outer
turbine component. Furthermore, the system includes a heat
exchanger configured to receive a flow of cooling air bled from the
gas turbine engine and transfer heat from the received flow of the
cooling air to a flow of coolant to generate cooled cooling air.
Moreover, the system includes a valve configured to control the
flow of the coolant to the heat exchanger. In this respect, the
cooled cooling air is supplied to at least one of the rotor disk or
the rotor blade to adjust the clearance between the rotor blade and
the outer turbine component.
[0007] In another aspect, the present subject matter is directed to
a system for controlling blade tip clearances within a gas turbine
engine. The gas turbine engine defines an axial centerline and a
radial direction extending orthogonal to the axial centerline. The
system includes an inner rotor configured to rotate in a first
direction and an inner rotor blade coupled to the inner rotor.
Additionally, the system includes an outer rotating drum configured
to rotate in a second direction opposite of the first direction and
an outer rotor blade coupled to the outer rotating drum.
Furthermore, the system includes heat exchanger configured to
receive a flow of cooling air bled from the gas turbine engine and
transfer heat from the received flow of the cooling air to a flow
of coolant to generate cooled cooling air. In addition, the system
includes a first air valve configured to direct a first portion of
the cooled cooling air to the outer rotating drum and a second
portion of the cooled cooling air to cool the inner rotor and a
second air valve configured to direct a first portion of the
cooling air to the outer rotating drum and a second portion of the
cooling air to cool the inner rotor. As such, the cooled cooling
air is supplied to at least one of the outer rotating drum or the
inner rotor to adjust a first clearance defined between the inner
rotor blade and the outer rotating drum and a second clearance
between the outer rotor blade and the inner rotor.
[0008] These and other features, aspects and advantages of the
present invention will become better understood with reference to
the following description and appended claims. The accompanying
drawings, which are incorporated in and constitute a part of this
specification, illustrate embodiments of the invention and,
together with the description, serve to explain the principles of
the invention.
BRIEF DESCRIPTION OF THE DRAWINGS
[0009] A full and enabling disclosure of the present invention,
including the best mode thereof, directed to one of ordinary skill
in the art, is set forth in the specification, which makes
reference to the appended figures, in which:
[0010] FIG. 1 is a schematic cross-sectional view of one embodiment
of a gas turbine engine of an aircraft;
[0011] FIG. 2 is a schematic cross-sectional view of another
embodiment of a gas turbine engine of an aircraft;
[0012] FIG. 3 is a schematic view of one embodiment of a system for
controlling blade clearances within a gas turbine engine;
[0013] FIG. 4 is an enlarged, partial schematic view of the system
for controlling blade clearances within a gas turbine engine shown
in FIG. 3, particularly illustrating a rotor disk and a rotor blade
of the gas turbine engine;
[0014] FIG. 5 is a cross-sectional side view of one embodiment of a
turbine section of a gas turbine engine;
[0015] FIG. 6 is a schematic view of another embodiment of a system
for controlling blade clearances within a gas turbine engine;
and
[0016] FIG. 7 is another schematic view of the system shown in FIG.
6.
[0017] Repeat use of reference characters in the present
specification and drawings is intended to represent the same or
analogous features or elements of the present invention.
DETAILED DESCRIPTION
[0018] Reference now will be made in detail to exemplary
embodiments of the presently disclosed subject matter, one or more
examples of which are illustrated in the drawings. Each example is
provided by way of explanation and should not be interpreted as
limiting the present disclosure. In fact, it will be apparent to
those skilled in the art that various modifications and variations
can be made in the present disclosure without departing from the
scope or spirit of the present disclosure. For instance, features
illustrated or described as part of one embodiment can be used with
another embodiment to yield a still further embodiment. Thus, it is
intended that the present disclosure covers such modifications and
variations as come within the scope of the appended claims and
their equivalents.
[0019] As used herein, the terms "first", "second", and "third" may
be used interchangeably to distinguish one component from another
and are not intended to signify location or importance of the
individual components.
[0020] Furthermore, the terms "upstream" and "downstream" refer to
the relative direction with respect to fluid flow in a fluid
pathway. For example, "upstream" refers to the direction from which
the fluid flows, and "downstream" refers to the direction to which
the fluid flows.
[0021] Additionally, the terms "low," "high," or their respective
comparative degrees (e.g., lower, higher, where applicable) each
refer to relative parameter magnitudes (e.g., speeds, pressures, or
temperatures) within an engine, unless otherwise specified. For
example, a "low-pressure turbine" operates at a pressure generally
lower than a "high-pressure turbine." Alternatively, unless
otherwise specified, the aforementioned terms may be understood in
their superlative degree. For example, a "low-pressure turbine" may
refer to the lowest maximum pressure turbine within a turbine
section, and a "high-pressure turbine" may refer to the highest
maximum pressure turbine within the turbine section.
[0022] In general, the present subject matter is directed to a
system for controlling blade clearances within a gas turbine
engine. As will be described below, the gas turbine engine includes
a shaft, a rotor disk coupled to the shaft, and a rotor blade
coupled to the rotor disk (e.g., via a dovetail connection) such
that the rotor blade extends outward from the disk along a radial
direction of the engine. Additionally, the gas turbine engine
includes an outer turbine component, such as a shroud or a
counter-rotating outer drum, positioned outward of the rotor blade
in the radial direction. As such, a clearance is defined between
the outer tip of the rotor blade and the outer turbine
component.
[0023] The disclosed system is configured to supply cooled cooling
air to the rotor disk and/or the rotor blade to adjust the
clearance between the rotor blade and the outer turbine component.
Specifically, the system includes a heat exchanger configured to
receive a flow of cooling air bled from the gas turbine engine. For
example, in one embodiment, the cooling air is bled from a
compressor discharge plenum of the engine. As such, the heat
exchanger is configured to transfer heat from the received flow of
cooling air to a flow of coolant (e.g., supercritical carbon
dioxide) to generate cooled cooling air. Additionally, the system
includes a valve configured to control the flow of the coolant to
the heat exchanger to adjust the temperature of the cooled cooling
air. The cooled cooling air is, in turn, routed to the rotor disk
and/or the rotor blade to adjust the clearance between the rotor
blade and the outer turbine component. For example, in some
embodiments, the cooled cooling air flows from the heat exchanger
to the rotor disk and/or the rotor blade through a conduit at least
partially positioned between the shaft and a combustor of the
engine.
[0024] The disclosed system provides one or more technical
advantages. For example, as described above, the disclosed system
supplies cooled cooling air to the rotor disk and/or the rotor
blade. Such cooled cooling air reduces the amount that rotor blade
expands as the engine heats up, thereby controlling clearance
between the rotor blade and outer turbine component. Furthermore,
as mentioned above, the temperature of the cooled cooling air may
be controlled by the valve. In this respect, increasing the amount
of and/or decreasing the temperature of the cooled cooling air
supplied to the rotor blade and/or the rotor disk may shrink the
rotor blade and/or the disk, thereby increasing the clearance via a
reduction in the blade tip radius. Conversely, the clearance may be
decreased by reducing the amount of and/or increasing the
temperature of cooling air supplied to the rotor blade and/or the
rotor disk. Moreover, the disclosed system allows the thermal
expansion/contraction of the rotor blade and/or the disk to be
controlled independently of the thermal expansion/contraction of
the outer turbine component.
[0025] Referring now to the drawings, FIG. 1 is a schematic
cross-sectional view of one embodiment of a gas turbine engine 10.
In the illustrated embodiment, the engine 10 is configured as a
high-bypass turbofan engine. However, in alternative embodiments,
the engine 10 may be configured as a propfan engine, a turbojet
engine, a turboprop engine, a turboshaft gas turbine engine, or any
other suitable type of gas turbine engine.
[0026] As shown in FIG. 1, the engine 10 defines a longitudinal
direction L, a radial direction R, and a circumferential direction
C. In general, the longitudinal direction L extends parallel to an
axial centerline 12 of the engine 10, the radial direction R
extends orthogonally outward from the axial centerline 12, and the
circumferential direction C extends generally concentrically around
the axial centerline 12.
[0027] In general, the engine 10 includes a fan 14, a low-pressure
(LP) spool 16, and a high pressure (HP) spool 18 at least partially
encased by an annular nacelle 20. More specifically, the fan 14 may
include a fan rotor 22 and a plurality of fan blades 24 (one is
shown) coupled to the fan rotor 22. In this respect, the fan blades
24 are spaced apart from each other along the circumferential
direction C and extend outward from the fan rotor 22 along the
radial direction R. Moreover, the LP and HP spools 16, 18 are
positioned downstream from the fan 14 along the axial centerline 12
(i.e., in the longitudinal direction L). As shown, the LP spool 16
is rotatably coupled to the fan rotor 22, thereby permitting the LP
spool 16 to rotate the fan 14. Additionally, a plurality of outlet
guide vanes or struts 26 spaced apart from each other in the
circumferential direction C extend between an outer casing 28
surrounding the LP and HP spools 16, 18 and the nacelle 20 along
the radial direction R. As such, the struts 26 support the nacelle
20 relative to the outer casing 28 such that the outer casing 28
and the nacelle 20 define a bypass airflow passage 30 positioned
therebetween.
[0028] The outer casing 28 generally surrounds or encases, in
serial flow order, a compressor section 32, a combustion section
34, a turbine section 36, and an exhaust section 38. For example,
in some embodiments, the compressor section 32 may include a
low-pressure (LP) compressor 40 of the LP spool 16 and a
high-pressure (HP) compressor 42 of the HP spool 18 positioned
downstream from the LP compressor 40 along the axial centerline 12.
Each compressor 40, 42 may, in turn, include one or more rows of
stator vanes 44 interdigitated with one or more rows of compressor
rotor blades 46. Moreover, in some embodiments, the turbine section
36 includes a high-pressure (HP) turbine 48 of the HP spool 18 and
a low-pressure (LP) turbine 50 of the LP spool 16 positioned
downstream from the HP turbine 48 along the axial centerline 12.
Each turbine 48, 50 may, in turn, include one or more rows of
stator vanes 52 interdigitated with one or more rows of turbine
rotor blades 54.
[0029] Additionally, the LP spool 16 includes the low-pressure (LP)
shaft 56 and the HP spool 18 includes a high pressure (HP) shaft 58
positioned concentrically around the LP shaft 56. In such
embodiments, the HP shaft 58 rotatably couples the rotor blades 54
of the HP turbine 48 and the rotor blades 46 of the HP compressor
42 such that rotation of the HP turbine rotor blades 54 rotatably
drives HP compressor rotor blades 46. As shown, the LP shaft 56 is
directly coupled to the rotor blades 54 of the LP turbine 50 and
the rotor blades 46 of the LP compressor 40. Furthermore, the LP
shaft 56 is coupled to the fan 14 via a gearbox 60. In this
respect, the rotation of the LP turbine rotor blades 54 rotatably
drives the LP compressor rotor blades 46 and the fan blades 24.
[0030] In several embodiments, the engine 10 may generate thrust to
propel an aircraft. More specifically, during operation, air
(indicated by arrow 62) enters an inlet portion 64 of the engine
10. The fan 14 supplies a first portion (indicated by arrow 66) of
the air 62 to the bypass airflow passage 30 and a second portion
(indicated by arrow 68) of the air 62 to the compressor section 32.
The second portion 68 of the air 62 first flows through the LP
compressor 40 in which the rotor blades 46 therein progressively
compress the second portion 68 of the air 62. Next, the second
portion 68 of the air 62 flows through the HP compressor 42 in
which the rotor blades 46 therein continue progressively
compressing the second portion 68 of the air 62. The compressed
second portion 68 of the air 62 is subsequently delivered to the
combustion section 34. In the combustion section 34, the second
portion 68 of the air 62 mixes with fuel and burns to generate
high-temperature and high-pressure combustion gases 70. Thereafter,
the combustion gases 70 flow through the HP turbine 48 which the HP
turbine rotor blades 54 extract a first portion of kinetic and/or
thermal energy therefrom. This energy extraction rotates the HP
shaft 58, thereby driving the HP compressor 42. The combustion
gases 70 then flow through the LP turbine 50 in which the LP
turbine rotor blades 54 extract a second portion of kinetic and/or
thermal energy therefrom. This energy extraction rotates the LP
shaft 56, thereby driving the LP compressor 40 and the fan 14 via
the gearbox 60. The combustion gases 70 then exit the engine 10
through the exhaust section 38.
[0031] FIG. 2 is a schematic cross-sectional view of another
embodiment of a gas turbine engine 10 of an aircraft. Like the
embodiment of the engine 10 shown in FIG. 1, the embodiment of the
engine 10 shown in FIG. 2 includes an LP turbine 50. However,
unlike the embodiment of the engine 10 shown in FIG. 1, in the
embodiment of the engine 10 shown in FIG. 2, the LP turbine 50 is a
counter-rotating turbine. Specifically, in such an embodiment, the
LP turbine 50 includes an inner rotor 72 configured to rotate in a
first direction (e.g., one of the clockwise or counter-clockwise
directions) and one or more rows of inner rotor blades 74 coupled
to and extending outward from the inner rotor 72 in the radial
direction R. Furthermore, in such an embodiment, the LP turbine 50
includes an outer rotating drum 76 configured to rotate in a second
direction opposite of the first direction (e.g., the other of
clockwise or counter-clockwise directions) and one or more rows of
outer rotor blades 78 extending inward from the drum 102 toward the
axial centerline 12 in the radial direction R. As shown, the rows
of outer rotor blades 78 are interdigitated with the rows of inner
rotor blades 74. In addition, the LP shaft 24 may be coupled to the
outer rotor 76 of the LP turbine 50 via a gearbox 80.
[0032] The configuration of the gas turbine engine 10 described
above and shown in FIG. 1 is provided only to place the present
subject matter in an exemplary field of use. Thus, the present
subject matter may be readily adaptable to any manner of gas
turbine engine configuration, including other types of
aviation-based gas turbine engines, marine-based gas turbine
engines, and/or land-based/industrial gas turbine engines.
[0033] FIG. 3 illustrates one embodiment of a system 100 for
controlling blade clearances within a gas turbine engine. In
general, the system 100 will be discussed in the context of the gas
turbine engine 10 described above and shown in FIGS. 1 and 2.
However, the disclosed system 100 may be implemented within any gas
turbine engine having any other suitable configuration.
[0034] As shown, in several embodiments, the combustion section 34
of the gas turbine engine 10 includes one or more combustors 102.
In general, the combustor(s) 102 is positioned outward from the
shafts 56, 58 along the radial direction R Each combustor 102
includes a liner 104 defining a combustion chamber 106 therein.
Moreover, each combustor 102 includes one or more fuel nozzles 108,
which supply a mixture of fuel and compressed air (e.g., the
compressed, the second portion 68 of the air 62) to the combustion
chamber 106. The fuel and air mixture burns within the combustion
chamber 106 to generate the high-temperature and high-pressure
combustion gases 70. Although FIG. 3 illustrates a single combustor
102, the combustion section 34 may, in other embodiments, include a
plurality of combustors 102.
[0035] Additionally, in several embodiments, the combustion section
34 includes a compressor discharge casing 110. In such embodiments,
the compressor discharge casing 110 at least partially surrounds or
otherwise encloses the combustor(s) 102 in the circumferential
direction C. In this respect, a compressor discharge plenum 112 is
defined between the compressor discharge casing 110 and the liner
104. The compressor discharge plenum 112 is, in turn, configured to
supply compressed air to the combustor(s) 102. Specifically, as
shown, the compressed air exiting the HP compressor 42 is directed
into the compressor discharge plenum 112 by an inlet guide vane
113. The compressed air within the compressor discharge plenum 112
will be referred to as compressed air 114. A portion of the
compressed air 114 is supplied to the combustion chamber(s) 106 of
the combustor(s) 102 by the fuel nozzle(s) 108 for use in
combusting the fuel. As will be described below, in some
embodiments, another portion of the compressed air 114 is used for
cooling components of the HP turbine 48 of the gas turbine engine
10.
[0036] As shown, the system 100 includes a heat exchanger 116. More
specifically, the heat exchanger 116 is configured to receive a
flow of cooling air (indicated by arrow 118) bled from the gas
turbine engine 10 and a flow of coolant (indicated by arrows 120).
In this respect, the heat exchanger 116 is configured to transfer
heat from the flow of the cooling air 118 to the flow of coolant
120. Such heat transfer cools the received cooling air 118, thereby
generating cooled cooling air (indicated by arrows 122). As will be
described below, the temperature of the cooled cooling air 122 may
be adjusted by controlling the volume of the coolant 120 flowing
through the heat exchanger 116. Thereafter, the cooled cooling air
122 is routed to the turbine section 36 to control the blade tip
clearances therein.
[0037] In several embodiments, the heat exchanger 116 is configured
to receive the cooling air 118 from the compressor discharge plenum
112. Specifically, in such embodiments, a portion of the compressed
air 114 is bled from the compressor discharge plenum 112 and routed
to the heat exchanger 116. For example, in one embodiment, the
system 100 includes a conduit 124 that conveys the compressed air
114 from the compressor discharge plenum 112 to the heat exchanger
116. Although not shown in FIG. 3, a suitable valve(s) may be
provided in associated with the conduit 124 to control the flow of
the compressed air 114 from the compressor discharge plenum 112 to
the heat exchanger 116. However, in alternative embodiments, the
cooling air 118 received by the heat exchanger 116 may be bled from
any other suitable location on the gas turbine engine 10, such as
the compressor section 32.
[0038] The heat exchanger 116 may be positioned at any suitable
location within the gas turbine engine 10. For example, as shown,
in one embodiment, the heat exchanger 116 is positioned outward
along the radial direction R from the combustor(s) 102.
[0039] Additionally, the flow of coolant 120 received by the heat
exchanger 116 may be formed from any suitable type of coolant. For
example, in one embodiment, the flow of coolant 120 may be a flow
of supercritical carbon dioxide.
[0040] As mentioned above, in some embodiments, the temperature of
the cooled cooling air 122 may be adjusted by controlling the flow
of the coolant 120 to the heat exchanger 116. In such embodiments,
the system 100 includes a valve 126 configured to control the flow
of the coolant 120 to the heat exchanger 116 and a bypass conduit
128. More specifically, the valve 126 is configured to adjust the
volume of the coolant 120 supplied to the heat exchanger 116 by
allowing a portion of the coolant to bypass the heat exchanger 116
via the bypass conduit 128. For example, the valve 126 may increase
the volume of the coolant 120 supplied to the heat exchanger 116 by
allowing less coolant 120 (or no coolant 120) to flow into the
bypass conduit 128. Such an increase in the volume of the coolant
120 supplied to the heat exchanger 116 decreases the temperature of
the cooled cooling air 122. Conversely, the valve 126 may decrease
the volume of the coolant 120 supplied to the heat exchanger 116 by
allowing more coolant 120 to flow into the bypass conduit 128. Such
a decrease in the volume of the coolant 120 supplied to the heat
exchanger 116 increases the temperature of the cooled cooling air
122.
[0041] Referring now to FIGS. 3 and 4, the cooled cooling air 122
is routed to the turbine section 36 to control the blade tip
clearances therein. In several embodiments, the cooled cooling air
122 may be used to control the blade tip clearances of a first
stage 130 of the HP turbine 48. However, in alternative
embodiments, the cooled cooling air 122 may be used to control the
blade tip clearances of any other blade tips within the turbine
section 36.
[0042] In general, the first stage 130 includes a row of
circumferentially arranged stator vanes 52 (one is shown) and a row
of circumferentially arranged rotor blades 54 (one is shown). As
shown, the stator vanes 52 are positioned downstream from the
combustion chamber 106 relative to the direction of the flow of the
combustion gases 70. As such, the stator vanes 52 define a
downstream end of the compressor discharge plenum 112. Furthermore,
the rotor blades 54 are positioned downstream from the stator vanes
52 in the direction of the flow of the combustion gases 70. In this
respect, the stator vanes 52 and rotor blades 52 partially form a
hot gas path 132 along which the combustion gases 70 flow through
the turbine section 36. More specifically, each stator vane 52
includes inner and outer bands 134, 136 respectively forming the
inner and outer boundaries of the hot gas path 132 in the radial
direction R. Each stator vane 54 also includes an airfoil 138
extending through the hot gas path 132 along the radial direction R
between the inner and outer bands 134, 136. Moreover, each rotor
blade 54 includes a base portion 140 and an airfoil 142 extending
outward in the radial direction R from the base portion 140 into
the hot gas path 132. The base portion 140 of each rotor blade 54
is coupled to a rotor disk 144 (e.g., via a dovetail connection, a
fir tree-type connection, etc.), with the rotor disk 144, in turn,
being coupled to the HP shaft 58. As such, rotation of the rotor
disk 144 and the rotor blades 54 rotate the HP shaft 58, which, in
turn, drives the compressor 32 as described above.
[0043] Moreover, in some embodiments, one or more seals may be
positioned adjacent of the rotor disk 144. For example, as shown in
FIG. 4, inner and outer seals 143, 145 are positioned upstream of
the rotor disk 144 along the axial centerline 12 relative to the
direction of the flow of the combustion gases 70 through the gas
turbine engine 10. In such an embodiment, the inner seal 143 is
positioned inward along the radial direction R of the outer seal
145 such that a gap 147 is defined therebetween. As will be
described below, the cooled cooling air 122 may flow through the
gap 147 toward the rotor disk 144 and then outward along the radial
direction R between the outer seal 145 and the rotor disk 144,
thereby cooling the rotor disk 144 and the rotor blade 54.
[0044] Additionally, the first stage 130 of the HP turbine 48
includes one or more outer turbine components 146 partially
defining the hot gas path 132. In general, the outer turbine
component(s) 146 is positioned outward of airfoil 142 of the rotor
blade 54 in the radial direction R such that the component(s) 146
define an outer boundary of the hot gas path 132 in the radial
direction R. As shown, a clearance (indicated by arrow 148) is
defined between the tips 150 of the airfoils 142 of the rotor
blades 54 and an inner radial surface(s) 152 of the outer turbine
component(s) 146. As will be described below, the clearance 148 may
be controlled by the cooled cooling air 122 supplied to the first
stage 130. For example, in the illustrated embodiment, the outer
turbine component(s) 146 is a shroud 154 enclosing the rotor blades
54. However, in alternative embodiments, the outer turbine
component(s) 146 may be any other suitable component(s), such as a
counter-rotating drum (e.g., the outer rotating drum 76) or a
shroud attached to a counter-rotating drum.
[0045] Furthermore, in several embodiments, the system 100 includes
a conduit 156. In general, the conduit 156 is configured to supply
the cooled cooling air 122 from the heat exchanger 116 to the rotor
blades 54 and the rotor disk 144 of the first stage 130. As such,
in some embodiments, the conduit 156 is at least partially
positioned between the combustor(s) 102 (and, more specifically,
the inner portion of the compressor discharge casing 110 in the
radial direction R) and the HP shaft 58 in the radial direction R.
Additionally, in some embodiments, the system 100 includes an
inducer 157 configured to direct the cooled cooling air flowing
through the conduit 156 toward the rotor disk 144. For example, as
shown, in one embodiment, the inducer 157 narrows as the inducer
157 extends from the downstream end of the conduit 156 toward the
rotor disk 144 to direct the cooled cooling air 122 through the gap
147.
[0046] The conduit 156 may have any suitable configuration for
routing the cooled cooling air 122 to the rotor disk 144 and/or the
rotor blade 54. For example, as mentioned above, in the illustrated
embodiment, the heat exchanger 116 is positioned outward from the
combustor(s) 102 in the radial direction R. In such an embodiment,
the conduit includes a first portion 158 extending along the radial
direction R from the heat exchanger 116 inward toward the axial
centerline 12. In one embodiment, the first portion 158 of the
conduit 156 is positioned upstream of the combustor(s) 102 relative
to the direction of flow of the combustion gases 70 through the gas
turbine engine 10. In addition, in several embodiments, the system
100 includes a valve 159 configured to control the flow of the
cooled cooling air 122 from the heat exchanger 116 to the cooling
passage 156. Furthermore, the conduit 156 includes a second portion
161 extending from the downstream end of the first portion 158
along the axial centerline 12 between the HP shaft 58 and the
combustor 102 toward the rotor disk 144. As indicated above, the
inducer 157 is positioned at the downstream end of the second
portion 161 to direct the cooled cooling air 121 exiting the
conduit 156 through the gap 147 and toward the rotor disk 144.
However, in alternative embodiments, the conduit 156 may have any
other suitable configuration.
[0047] In some embodiments, the flow of cooled cooling air 122
supplied to the rotor disk 144 and/or the rotor blade 54 by the
conduit 156 is supplemented with additional compressed air 114 from
the compressor discharge plenum 112. More specifically, as shown in
FIG. 4, in such embodiments, the inner radial side of the
compressor discharge casing 110 defines a bleed port 160 fluidly
coupling the compressor discharge plenum 112 and the cooling
passage 156. In this respect, a portion of the compressed air 114
from the compressor discharge plenum 112 flows through the bleed
port 160 and directly into the cooling passage 156. This additional
compressed air 114 may increase the volume of the cooling air 122
supplied to the turbine section 36, thereby increasing the cooling
capacity of such air 122 without increasing the size of the heat
exchanger 116. In one embodiment, a valve (not shown) may control
the flow the additional compressed air 114 through the bleed port
160.
[0048] Referring particularly to FIG. 4, in several embodiments,
the cooled cooling air 122 flowing through the conduit 156 is
supplied to the rotor disk 144 and the rotor blades 54 of the first
stage 130 of the HP turbine 48. More specifically, the cooled
cooling air 122 flows inward along the radial direction R from the
heat exchanger 116 through the first portion 158 of the conduit 156
and subsequently downstream relative to the direction of flow of
the combustion gases 70 through the second portion 161 of the
conduit 156. The inducer 157 then directs the cooled cooling air
122 exiting the conduit 156 through the gap 147 between the seals
143, 145 and onto the rotor disk 144 of the first stage 130. The
cooled cooling air 122 then flows outward in the radial direction R
between the outer seal 145 and a forward or upstream surface 162 of
the rotor disk 144 such that the cooled cooling air 122 cools the
disk 144. Thereafter, the cooled cooling air 122 flows along
forward or upstream surfaces 164 of the the base portions 140 of
the first stage rotor blades 54. In one embodiment, a portion of
the cooled cooling air 122 flows through passages 166 (one is
shown) defined by the base portions 140 of the first stage rotor
blades 54, thereby cooling the interiors of the rotor blades
54.
[0049] As indicated above, the cooled cooling air 122 allows the
clearance 148 between the rotor blade tips 150 and the outer
turbine component(s) 146 to be controlled. More specifically, the
cooling of the first stage rotor disk 144 and rotor blades 54
provided by the cooled cooling air 122 causes the disk 144 and the
rotor blades 54 to shrink in the radial direction R. In this
respect, increasing the amount of and/or decreasing the temperature
(e.g., by controlling the valve 126) of the cooled cooling air 122
supplied to the rotor disk 144 and the rotor blades 54 increases
the amount such components shrink, thereby increasing the clearance
152. Conversely, decreasing the amount of and/or increasing the
temperature (e.g., by controlling the valve 126) of the cooled
cooling air 122 supplied to the rotor disk 144 and the rotor blades
54 causes the components to grow, thereby decreasing the clearance
152. As such, the disclosed system 100 allows the clearance 148 to
be minimized as the temperature of the gas turbine engine 10 varies
during operation.
[0050] After cooling the first stage rotor disk 144 and rotor
blades 54, the spent cooled cooling air 122 may be exhausted into
the hot gas path 132. For example, in some embodiments, at least a
portion of the spent cooled cooling air 122 may flow along the
upstream surfaces 164 of the rotor blades 54 and be exhausted in
the hot gas path through a clearance 168. The clearance 168 is, in
turn, defined between the inner bands 134 of the stator vanes 52
and the platforms of the rotor blades 54. Moreover, in some
embodiments, at least a portion of the spent cooled cooling air 122
may flow through the passages 166 in the base portions 140 of the
first stage rotor blades 54 and be exhausted in the hot gas path
through an outlet 170. However, in alternative embodiments, the
spent cooled cooling air 122 may be exhausted into the hot gas path
132 in any other suitable manner.
[0051] In several embodiments, the first stage outer turbine
component(s) 146 are cooled in a controlled manner to further
control the size of the clearance 148 between the outer turbine
component(s) 146 and the rotor blade tips 150. Specifically, in
such embodiments, compressed air 114 from the compressor discharge
plenum 112 is supplied to the outer turbine component(s) 146 to
cool this component(s) 146, thereby shrinking the component(s) 146.
Shrinking the outer turbine component(s) 146, in turn, decreases
the clearance 148. For example, in one embodiment, the shroud 154
(e.g., a 360-degree shroud) defines a passage 172 through which the
compressed air 114 flows to the cool the shroud 154. However, in
other embodiments, the compressed air 114 may be simply directed at
the outer radial side the outer turbine component(s) 146.
Furthermore, the compressed air 114 may be supplied to a turbine
case 173 to which the outer turbine component(s) 146 is coupled to
adjust the clearance between the rotor blade tip(s) 150 and the
outer turbine component(s) 146. After cooling the outer turbine
component(s) 146, the spent compressed air 114 may be exhausted
into the hot gas path 132. In other embodiments, the air supplied
to cool the outer turbine components, such as a 360-degree ring
shroud or a counter-rotating drum (with or without attached
segmented shrouds), may be cooled cooling air cooled by an
independent heat exchanger having a coolant (e.g., supercritical
CO2) which is controlled by an independent valve. This cooled
cooling air 122 used to cool the outer turbine components may also
be controlled or metered using an in-line air valve, such as the
valve 159.
[0052] The flow of the cooled cooling air 122 to the first stage
rotor disk 144 and the rotor blades 54 may be controlled
independently of the flow of the compressed air 114 or cooled
cooling air 122 to the outer turbine component(s) 146. As such, the
clearance 148 between the outer turbine component(s) 146 and rotor
blade tips 150 may be adjusted by controlling the flow and
temperature of the cooled cooling air 122 to the first stage rotor
disk 144 and the rotor blades 54, the flow of the compressed air
114 or the flow and temperature of the cooled cooling air 122 to
the outer turbine component(s) 146, or both.
[0053] Additionally, in some embodiments, the system 100 may be
used to control the sizes of the clearances in counter-rotating
turbines. More specifically, as shown in FIG. 5, in such a turbine
(e.g., the LP turbine 50 shown in FIG. 2), a first clearance
(indicated by arrow 174) is defined between the tips 176 of the
airfoils of the inner rotor blades 74 and an inner radial
surface(s) 178 of the outer rotating drum 76. Moreover, a second
clearance (indicated by arrow 180) is defined between the tips 182
of the airfoils of the outer rotor blades 78 and an outer radial
surface(s) 184 of the inner rotor 72.
[0054] FIG. 6 is a schematic view of another embodiment of a system
100 for controlling blade clearances within a gas turbine engine.
Like the embodiment of the system 100 shown in FIGS. 3 and 4, the
system 100 shown in FIG. 6 includes a heat exchanger 116 configured
to receive and cool cooling air 118 to generate cooled cooling air
122. However, unlike the embodiment of the system 100 shown in
FIGS. 3 and 4, the system 100 shown in FIG. 6 includes a first air
valve 186 in fluid communication with the heat exchanger 116. In
this respect, the first air valve 186 is configured to direct or
otherwise route a first portion 188 of the cooled cooling air 122
from the heat exchanger 116 to the outer rotating drum 76 and a
second portion 190 of the cooled cooling air 122 from the heat
exchanger 116 to cool the inner rotor 72. Furthermore, unlike the
embodiment of the system 100 shown in FIGS. 3 and 4, the system 100
shown in FIG. 6 includes a second air valve 192 configured to route
a first portion of cooling air 118 (e.g., cooling air 118 bled from
the compressor discharge plenum 112, but not delivered to the heat
exchanger 116) to the outer rotating drum 76 and a second portion
196 of the cooling air 118 to cool the inner rotor 72.
[0055] As indicated above, the cooling air 118 and the cooled
cooling air 122 allow the first and second clearances 174 and 180
to be controlled. More specifically, the cooling of the inner rotor
72 and the outer rotating drum 76 provided by the cooling air 118
and the cooled cooling air 122 causes the inner rotor 72 and the
outer rotating drum 76 to shrink in the radial direction R. In this
respect, increasing the amount of cooled cooling air 122 and the
decreasing the amount of cooled air 118 (e.g., by controlling the
valves 186, 192) supplied to the inner rotor 72 and the outer
rotating drum 76 increases the amount such components shrink,
thereby increasing the clearance 174, 180. Conversely, decreasing
the amount of cooled cooling air 122 and the increasing the amount
of cooled air 118 (e.g., by controlling the valves 186, 192)
supplied to the inner rotor 72 and the outer rotating drum 76
causes the components to grow, thereby decreasing the clearance
152. As such, the disclosed system 100 allows the clearances, 174,
180 to be minimized as the temperature of the gas turbine engine 10
varies during operation.
[0056] As shown in FIGS. 6 and 7, the cooled cooling air 122 and
the cooling air 118 are delivered to the outer rotating drum 76 and
the inner rotor 72 via angled nozzles 198 to further affect the
cooling of the outer rotating drum 76 and the inner rotor 72. More
specifically, the first portion 188 of the cooled cooling air 122
may be introduced to or otherwise directed at the outer rotating
drum 76 through one or more angled nozzles 198 such that a
tangential component of the velocity of the first portion 188 of
the cooled cooling air 122 is in the second direction (i.e., the
direction in which the outer rotating drum 76 rotates).
Furthermore, the second portion 190 of the cooled cooling air 122
may be introduced to or otherwise directed at the inner rotor 72
through one or more angled nozzles 198 such that a tangential
component of the velocity of the second portion 190 of the cooled
cooling air 122 is in the first direction (i.e., the direction in
which the inner rotor 72 rotates). Directing the cooled cooling air
122 in the same direction as the rotation of the inner rotor 72 and
the outer rotating drum 76 increases the cooling that the cooled
cooling air 122 provides. Conversely, the first portion 194 of the
cooling air 118 may be introduced to or otherwise directed at the
outer rotating drum 76 through one or more angled nozzles 198 such
that a tangential component of the velocity of the first portion
194 of the cooling air 118 is in the first direction (i.e., the
opposite direction to which the outer rotating drum 76 rotates).
Moreover, the second portion 196 of the cooling air 118 may be
introduced to or otherwise directed at the inner rotor 72 through
one or more angled nozzles 198 such that a tangential component of
the velocity of the second portion 196 of the cooling air 118 is in
the second direction (i.e., the opposite direction in which the
inner rotor 72 rotates). Directing the cooled cooling air 122 in
the opposite direction as the rotation of the inner rotor 72 and
the outer rotating drum 76 decreases the cooling that the cooled
cooling air 122 provides. In this respect, the controlling the
amount of cooling air 118 and the cooled cooling air 122 (e.g.,
with the valves 186, 192) and its direction of flow relative to the
inner rotor 72 and the outer rotating drum 76 (e.g., via the
nozzles 198), the clearances, 174, 180 to be minimized as the
temperature of the gas turbine engine 10 varies during
operation.
[0057] This written description uses examples to disclose the
invention, including the best mode, and also to enable any person
skilled in the art to practice the invention, including making and
using any devices or systems and performing any incorporated
methods. The patentable scope of the invention 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 include 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.
[0058] Further aspects of the invention are provided by the subject
matter of the following clauses:
[0059] A system for controlling blade clearances within a gas
turbine engine, the gas turbine engine defining an axial centerline
and a radial direction extending orthogonal to the axial
centerline, the system comprising: a rotor disk; a rotor blade
coupled to the rotor disk; an outer turbine component positioned
outward of the rotor blade in the radial direction such that a
clearance is defined between the rotor blade and the outer turbine
component; a heat exchanger configured to receive a flow of cooling
air bled from the gas turbine engine and transfer heat from the
received flow of the cooling air to a flow of coolant to generate
cooled cooling air; and a valve configured to control the flow of
the coolant to the heat exchanger, wherein the cooled cooling air
is supplied to at least one of the rotor disk or the rotor blade to
adjust the clearance between the rotor blade and the outer turbine
component.
[0060] The system of one or more of these clauses, further
comprising: a shaft coupled to the rotor disk such that rotation of
the rotor disk and the rotor blade rotates the shaft; a combustor
positioned outward in the radial direction from the shaft; and a
conduit at least partially positioned between the shaft and the
combustor in the radial direction such that the cooled cooling air
flows through the cooling passage from the heat exchanger to the at
least one of the rotor disk or the rotor blade.
[0061] The system of one or more of these clauses, further
comprising: an inducer configured to direct the cooled cooling air
flowing through the conduit toward the rotor disk.
[0062] The system of one or more of these clauses, wherein the
inducer narrows as the inducer extends from the conduit toward the
rotor disk.
[0063] The system of one or more of these clauses, further
comprising: a seal positioned upstream of the rotor disk along the
axial centerline relative to a direction of flow through the gas
turbine engine, wherein the inducer directs the cooled cooling air
such that the cooled cooling air flows between the rotor disk and
the seal.
[0064] The system of one or more of these clauses, wherein the seal
corresponds to an outer seal, the system further comprising: an
inner seal positioned inward along the radial direction relative to
the outer seal such that a gap is defined between the inner and
outer seals through which the cooled cooling air flows from the
inducer toward the rotor disk.
[0065] The system of one or more of these clauses wherein the
conduit includes a first portion extending along the radial
direction from the heat exchanger and a second portion extending
along the axial centerline from the first portion toward the rotor
disk.
[0066] The system of one or more of these clauses, wherein the
first portion of the conduit is positioned upstream of the
combustor relative to a direction of flow through the gas turbine
engine.
[0067] The system of one or more of these clauses, wherein the heat
exchanger is positioned outward along the radial direction from the
combustor.
[0068] The system of one or more of these clauses, further
comprising: a compressor discharge casing at least partially
surrounding the combustor, the compressor discharge casing defining
a compressor discharge plenum configured to supply compressed air
to the combustor, wherein the cooling air received by the heat
exchanger is bled from the compressor discharge plenum.
[0069] The system of one or more of these clauses, further
comprising: a turbine case coupled to the outer turbine components,
wherein the cooled cooling air is supplied to the turbine case to
adjust the clearance between the rotor blade and the outer turbine
component.
[0070] The system of one or more of these clauses, further
comprising: a bypass conduit fluidly coupled to the valve such that
the bypass conduit is configured to permit at least a portion of
the coolant to bypass the heat exchanger.
[0071] The system of one or more of these clauses, wherein the
cooled cooling air is discharged into a hot gas path at least
partially defined by the rotor blade and the outer turbine
component after being supplied to the at least one of the rotor
disk or the rotor blade.
[0072] The system of one or more of these clauses, wherein the
coolant comprises supercritical carbon dioxide.
[0073] The system of one or more of these clauses, wherein the
outer turbine component comprises a shroud or an outer rotating
drum.
[0074] A system for controlling blade tip clearances within a gas
turbine engine, the gas turbine engine defining an axial centerline
and a radial direction extending orthogonal to the axial
centerline, the system comprising: an inner rotor configured to
rotate in a first direction; an inner rotor blade coupled to the
inner rotor; an outer rotating drum configured to rotate in a
second direction opposite of the first direction; an outer rotor
blade coupled to the outer rotating drum; a heat exchanger
configured to receive a flow of cooling air bled from the gas
turbine engine and transfer heat from the received flow of the
cooling air to a flow of coolant to generate cooled cooling air; a
first air valve configured to direct a first portion of the cooled
cooling air to the outer rotating drum and a second portion of the
cooled cooling air to cool the inner rotor; and a second air valve
configured to direct a first portion of the cooling air to the
outer rotating drum and a second portion of the cooling air to cool
the inner rotor, wherein the cooled cooling air is supplied to at
least one of the outer rotating drum or the inner rotor to adjust a
first clearance defined between the inner rotor blade and the outer
rotating drum and a second clearance between the outer rotor blade
and the inner rotor.
[0075] The system of one or more of these clauses, where the first
portion of the cooled cooling air is introduced to the outer
rotating drum through an angled nozzle such that a tangential
component of a velocity of the first portion of the cooled cooling
air is in the second direction.
[0076] The system of one or more of these clauses, where the second
portion of the cooled cooling air is introduced to the inner rotor
through an angled nozzle such that a tangential component of a
velocity of the second portion of the cooled cooling air is in the
first direction.
[0077] The system of one or more of these clauses, where the first
portion of the cooling air is introduced to outer rotating drum
through an angled nozzle such that a tangential component of a
velocity of the first portion of the cooling air is in the first
direction.
[0078] The system of one or more of these clauses, where the second
portion of the cooling air is introduced to inner rotor through an
angled nozzle such that a tangential component of a velocity of the
second portion of the cooling air is in the second direction.
* * * * *