U.S. patent application number 14/132024 was filed with the patent office on 2015-06-18 for adjustable clearance control system for airfoil tip in gas turbine engine.
The applicant listed for this patent is John A. Orosa. Invention is credited to John A. Orosa.
Application Number | 20150167488 14/132024 |
Document ID | / |
Family ID | 52278820 |
Filed Date | 2015-06-18 |
United States Patent
Application |
20150167488 |
Kind Code |
A1 |
Orosa; John A. |
June 18, 2015 |
ADJUSTABLE CLEARANCE CONTROL SYSTEM FOR AIRFOIL TIP IN GAS TURBINE
ENGINE
Abstract
An airfoil system for use in a gas turbine engine having an
adjustable clearance control system including an axially adjustable
ring segment releasably coupled to the stationary turbine component
whereby the axially adjustable ring segment may be controlled
independently of other airfoil stages is disclosed The adjustable
clearance control system may thus control the flow of hot gases
passing one particular airfoil stage while the flow passing other
airfoil stages within the component of the turbine engine remains
unchanged The adjustable clearance control system may control the
size of the gap between the axially adjustable ring segment and the
tip of an airfoil through axial movement of the axially adjustable
ring segment. The axially adjustable ring segment may include a
radially inward contact surface that is positioned nonparallel and
nonorthogonal relative to a direction of movement of the axial
adjustable ring segment.
Inventors: |
Orosa; John A.; (Palm Beach
Gardens, FL) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Orosa; John A. |
Palm Beach Gardens |
FL |
US |
|
|
Family ID: |
52278820 |
Appl. No.: |
14/132024 |
Filed: |
December 18, 2013 |
Current U.S.
Class: |
415/173.2 |
Current CPC
Class: |
F05D 2220/3215 20130101;
F05D 2250/283 20130101; F01D 11/22 20130101; F05D 2240/55
20130101 |
International
Class: |
F01D 11/22 20060101
F01D011/22 |
Claims
1. An airfoil system for use in a gas turbine engine, comprising: a
rotor assembly having a plurality of airfoils extending radially
therefrom and aligned axially to form a circumferentially extending
row of airfoils forming an airfoil stage; a stationary turbine
component positioned radially outward from a tip of at least one
airfoil; an adjustable clearance control system including an
axially adjustable ring segment releasably coupled to the
stationary turbine component, wherein the axially adjustable ring
segment is adjustable axially; and wherein the axially adjustable
ring segment is adjustable axially when positioned radially outward
from a single row of turbine blades, thereby enabling the axially
adjustable ring segment to be moved axially to change a gap size
between the tip of the airfoil and the axially adjustable ring
segment independently of other airfoil stages
2. The airfoil system of claim 1, wherein the axially adjustable
ring segment includes a radially inward segment contact surface
having at least a portion aligned with the tip of the airfoil.
3. The airfoil system of claim 1, wherein the axially adjustable
ring segment includes a honeycomb seal land coupled to the axially
adjustable ring segment
4. The airfoil system of claim 3, wherein a radially inward contact
surface of the honeycomb seal land is positioned nonparallel and
nonorthogonal relative to a direction of movement of the axial
adjustable ring segment.
5. The airfoil system of claim 1, further comprising at least one
rib extending radially outward from the tip of the turbine blade
and terminating before contacting a radially inward contact surface
of the axially adjustable ring segment
6. The airfoil system of claim 1, wherein the plurality of airfoils
forming the circumferentially extending row of airfoils forming the
airfoil stage forms an airfoil stage closest to an exhaust
diffuser.
7. The airfoil system of claim 1, wherein the axially adjustable
ring segment of the adjustable clearance control system extends at
least partially circumferentially about the airfoil stage.
8. An airfoil system for use in a gas turbine engine, comprising: a
rotor assembly having a plurality of airfoils extending radially
therefrom and aligned axially to form a circumferentially extending
row of airfoils forming an airfoil stage; a stationary turbine
component positioned radially outward from a tip of at least one
airfoil; an adjustable clearance control system including an
axially adjustable ring segment releasably coupled to the
stationary turbine component, wherein the axially adjustable ring
segment is adjustable axially; wherein the axially adjustable ring
segment is adjustable axially when positioned radially outward from
a single row of turbine blades, thereby enabling the axially
adjustable ring segment to be moved axially to change a gap size
between the tip of the airfoil and the axially adjustable ring
segment independently of other airfoil stages; and a radially
inward segment contact surface of the axially adjustable ring
segment is positioned nonparallel and nonorthogonal relative to a
direction of movement of the axial adjustable ring segment
9. The airfoil system of claim 8, wherein the axially adjustable
ring segment includes a radially inward contact surface having at
least a portion aligned with the tip of the airfoil
10. The airfoil system of claim 8, wherein the axially adjustable
ring segment includes a honeycomb seal land coupled to the axially
adjustable ring segment
11. The airfoil system of claim 8, further comprising at least one
rib extending radially outward from the tip of the turbine blade
and terminating before contacting a radially inward contact surface
of the axially adjustable ring segment.
12. The airfoil system of claim 8, wherein the plurality of
airfoils forming the circumferentially extending row of airfoils
forming the airfoil stage forms an airfoil stage closest to an
exhaust diffuser.
13. The airfoil system of claim 8, wherein the axially adjustable
ring segment of the adjustable clearance control system extends at
least partially circumferentially about the airfoil stage
14. An airfoil system for use in a gas turbine engine, comprising:
a rotor assembly having a plurality of airfoils extending radially
therefrom and aligned axially to form a circumferentially extending
row of airfoils forming an airfoil stage; a stationary turbine
component positioned radially outward from a tip of at least one
airfoil; an adjustable clearance control system including an
axially adjustable ring segment releasably coupled to the
stationary turbine component, wherein the axially adjustable ring
segment is adjustable axially; wherein the axially adjustable ring
segment is adjustable axially when positioned radially outward from
a single row of turbine blades, thereby enabling the axially
adjustable ring segment to be moved axially to change a gap size
between the tip of the airfoil and the axially adjustable ring
segment independently of other airfoil stages and without changing
a tip clearance of other upstream airfoil stages; wherein a
radially inward segment contact surface of the axially adjustable
ring segment is positioned nonparallel and nonorthogonal relative
to a direction of movement of the axial adjustable ring segment;
and wherein the axially adjustable ring segment includes a
honeycomb seal land coupled to the axially adjustable ring
segment.
15. The airfoil system of claim 14, wherein the axially adjustable
ring segment includes a radially inward contact surface having at
least a portion aligned with the tip of the airfoil.
16. The airfoil system of claim 14, further comprising at least one
rib extending radially outward from the tip of the turbine blade
and terminating before contacting a radially inward contact surface
of the axially adjustable ring segment
17. The airfoil system of claim 14, wherein the plurality of
airfoils forming the circumferentially extending row of airfoils
forming the airfoil stage forms an airfoil stage closest to an
exhaust diffuser.
18. The airfoil system of claim 14, wherein the axially adjustable
ring segment of the adjustable clearance control system extends at
least partially circumferentially about the airfoil stage.
Description
FIELD OF THE INVENTION
[0001] This invention is directed generally to turbine engines, and
more particularly to systems for controlling gaps between airfoil
tips and radially adjacent components in turbine engines so as to
control the flow of fluids downstream based upon diffuser inlet
conditions.
BACKGROUND
[0002] Typically, gas turbine engines are formed from a compressor
positioned upstream from a turbine blade assembly. The compressor
and turbine are formed from a plurality of blade stages coupled to
discs that are capable of rotating about a longitudinal axis Each
blade stage is formed from a plurality of blades extending radially
about the circumference of the disc
[0003] The tips of the blades are located in close proximity to an
inner surface of the casing of the turbine engine There typically
exists a gap between the blade tips and the casing of the turbine
engine so that the blades may rotate without striking the casing.
Likewise, for nonshrouded vanes, there typically exists a gap
between the vane tips and an internal rotatable blade and disc
assembly so that the rotatable blade and disc assembly may rotate
without the vanes contacting the rotatable blade and disc assembly
During operation, gases pass the blades and vanes and compress to
high temperature and pressure These gases also heat the casing,
blades, vanes and discs causing each to expand due to thermal
expansion. After the turbine engine has been operating at full load
conditions for a period of time, the components reach a maximum
operating condition at which maximum thermal expansion occurs In
this state, it is desirable that the gap between the blade tips and
the casing of the turbine engine and the gap between the vanes and
rotatable blade and disc assembly be as small as possible to limit
leakage past the tips of the airfoils
[0004] However, reducing the gap cannot be accomplished by simply
positioning the components so that the gap is minimal under full
load conditions because the configuration of the components forming
the gap must account for warm restart conditions in which the
casing and the compressor vane carriers, having less mass than the
blade and disc assembly, cools faster than the blade and disc
assembly. During a warm restart, the discs expand due to
centrifugal forces and the clearances tighten before the casing
begins to heat up and expand. Therefore, unless the components have
been positioned so that a sufficient gap has been established
between the blades and the casing and between the vanes and the
rotatable blade and disc assembly under operating conditions, the
airfoils may strike the casing or the rotatable blade and disc
assembly because the diameter of components forming the casing have
not heated up and expanded yet. Collision between the blades and
the casing or compressor vanes and the rotatable blade and disc
assembly often causes severe airfoil tip rubs and may result in
damage. There exist systems that reducing gaps between every blade
tip and a casing simultaneously However, at certain conditions,
diffuser performance can be significantly improved by increasing
the last blade tip leakage jet Thus, a need exists for an improved
system for regulating the last blade tip leakage jet
SUMMARY OF THE INVENTION
[0005] An airfoil system for use in a gas turbine engine having an
adjustable clearance control system including an axially adjustable
ring segment releasably coupled to the stationary turbine component
whereby the axially adjustable ring segment may be controlled
independently of other airfoil stages is disclosed The adjustable
clearance control system may thus control the flow of hot gases
passing one particular airfoil stage while the flow passing other
airfoil stages within the component of the turbine engine remains
unchanged The adjustable clearance control system may control the
size of the gap between the axially adjustable ring segment and the
tip of an airfoil through axial movement of the axially adjustable
ring segment. The axially adjustable ring segment may include a
radially inward contact surface that is positioned nonparallel and
nonorthogonal relative to a direction of movement of the axial
adjustable ring segment to adjust the size of the gap.
[0006] The adjustable clearance control system may address the need
to open the clearance on the last blade tip without opening the
clearance on the other blades Normally, turbine engines are
operated with the tightest possible clearance on every blade.
However, at certain conditions, diffuser performance can be
significantly improved by increasing the last blade tip leakage jet
Even though opening the gap reduces the blade efficiency, that loss
can be more than offset by the improvement in the downstream
diffuser. One particular situation in which diffuser performance
can be significantly improved is at baseload, which is maximum
power standard day, if the turbine has a hub strong velocity
profile. This situation occurs when turbine engines are operated at
higher power levels, which can occur by increasing mass flow,
without changing the turbine design. This situation can also occur
in well designed turbines having a flat velocity profile when they
operate on cold days
[0007] The airfoil system may include a rotor assembly having a
plurality of airfoils extending radially therefrom and aligned
axially to form a circumferentially extending row of airfoils
forming an airfoil stage. The airfoil system may also include a
stationary turbine component positioned radially outward from a tip
of the airfoil The airfoil system may include an adjustable
clearance control system including an axially adjustable ring
segment releasably coupled to the stationary turbine component,
wherein the axially adjustable ring segment is adjustable axially
The axially adjustable ring segment may be adjustable axially when
positioned radially outward from a single row of turbine blades,
thereby enabling the axially adjustable ring segment to be moved
axially to change a gap size between the tip of the airfoil and the
axially adjustable ring segment independently of other airfoil
stages.
[0008] The axially adjustable ring segment may include a radially
inward contact surface having at least a portion aligned with the
tip of the airfoil. The axially adjustable ring segment may include
a honeycomb seal land coupled to the axially adjustable ring
segment A radially inward contact surface of the honeycomb seal
land may be positioned nonparallel and nonorthogonal relative to a
direction of movement of the axial adjustable ring segment. One or
more ribs may extend radially outward from the tip of the turbine
blade and may terminate before contacting a radially inward contact
surface of the axially adjustable ring segment or of the honeycomb
seal land In at least one embodiment, the rib may be formed from a
tip shroud. The plurality of airfoils forming the circumferentially
extending row of airfoils forming the airfoil stage may form an
airfoil stage closest to an exhaust diffuser. The axially
adjustable ring segment of the adjustable clearance control system
may extend at least partially circumferentially about the airfoil
stage.
[0009] During operation, the adjustable clearance control system
may enable the axially adjustable ring segment to be moved axially
such that the gap may be reduced or increased. For instance, during
startup of the gas turbine engine, the adjustable clearance control
system may move the axially adjustable ring segment axially such
that the gap may be increased to prevent tip rubbing and possible
damage. Once at steady state operating conditions, the adjustable
clearance control system may move the axially adjustable ring
segment to reduce the gap independently of other stages. During
shutdown of the gas turbine engine, the adjustable clearance
control system may move the axially adjustable ring segment to
increase the gap independently of other stages to prevent tip
rubbing and possible damage. In addition, increasing the clearance
of the last stage blade on cold days or on turbine engines that
have a hub strong exit velocity profile may improve engine
performance Under such conditions, the exhaust diffuser may be
overloaded on the outer diamater (OD) due to higher velocities on
the hub pulling flow away. Increasing the last stage blade tip gap
can inject high velocity air at the OD to keep the OD flow healthy
in the exhaust diffuser.
[0010] An advantage of the adjustable clearance control system is
that the system may control the size of a gap between airfoil tips
and a radially outward ring segment by adjusting the axially
adjustable ring segment axially independently of other airfoil
stages. Thus, the size of the gap between the airfoil tips and a
radially outward ring segment may be adjusted without adjusting or
interfering with gap sizes between airfoil tips and ring segments
of other airfoil stages within the same gas turbine engine
[0011] Another advantage of the adjustable clearance control system
is that the system may include a honeycomb seal land having a
radially inward contact surface configured to absorb contact from
airfoil tips without damaging the airfoils.
[0012] These and other embodiments are described in more detail
below
BRIEF DESCRIPTION OF THE DRAWINGS
[0013] The accompanying drawings, which are incorporated in and
form a part of the specification, illustrate embodiments of the
presently disclosed invention and, together with the description,
disclose the principles of the invention
[0014] FIG. 1 is a cross-sectional perspective view of a turbine
engine with an adjustable clearance control system
[0015] FIG. 2 is a cross-sectional detail view of the adjustable
clearance control system positioned in a gas turbine engine taken
at detail 2-2 in FIG. 1.
[0016] FIG. 3 is a perspective view of a honeycomb seal land
coupled to the axially adjustable ring segment.
[0017] FIG. 4 is a perspective view of a shrouded turbine blade
usable with the adjustable clearance control system
DETAILED DESCRIPTION OF THE INVENTION
[0018] As shown in FIGS. 1-4, an airfoil system 10 for use in a gas
turbine engine 12 having an adjustable clearance control system 14
including an axially adjustable ring segment 16 releasably coupled
to the stationary turbine component 36 whereby the axially
adjustable ring segment 16 may be controlled independently of other
airfoil stages 18 is disclosed The adjustable clearance control
system 14 may thus control the flow of hot gases passing one
particular airfoil stage 20 while the flow passing other airfoil
stages 18 within the component of the turbine engine 12 remains
unchanged. The adjustable clearance control system 14 may control
the size of the gap 22 between the axially adjustable ring segment
16 and a tip 24 of an airfoil 26 through axial movement of the
axially adjustable ring segment 16. The axially adjustable ring
segment 16 may include a radially inward segment contact surface 27
that is positioned nonparallel and nonorthogonal relative to a
direction of movement 30 of the axial adjustable ring segment
16.
[0019] The airfoil system 10 may include a rotor assembly 32 having
a plurality of airfoils 26 extending radially therefrom and aligned
axially to form a circumferentially extending row 34 of airfoils 26
forming an airfoil stage 20 The airfoils 26 may have any
appropriate shape or configuration. The airfoil system 10 may also
include a stationary turbine component 36 positioned radially
outward from the tip 24 of the airfoil 26. In at least one
embodiment, the stationary component 36 may be a turbine ring
segment. In other embodiments, the stationary component 36 may be
other components that remain stationary relative to the rotor
assembly 32.
[0020] The airfoil system 10 may also include an adjustable
clearance control system 14 including an axially adjustable ring
segment 16 releasably coupled to the stationary turbine component
36, whereby the axially adjustable ring segment 16 may be
adjustable axially. In particular, the axially adjustable ring
segment 16 may be adjustable axially when positioned radially
outward from a single row 34 of turbine blades 38 forming an
airfoil stage 20, thereby enabling the axially adjustable ring
segment 16 to be moved axially to change a gap size 22 between the
tip 24 of the airfoil 26 and the axially adjustable ring segment 16
independently of other airfoil stages 18. In at least one
embodiment, the plurality of airfoils 26 forming the
circumferentially extending row 34 of airfoils forming the airfoil
stage 20 forms an airfoil stage 20 closest to an exhaust diffuser
48. As such, the airfoil stage 20 may be the last turbine blade
stage in a turbine blade assembly in reference to moving in a
downstream direction. The exhaust diffuser 48 may be positioned aft
of the turbine. The axially adjustable ring segment 16 may be used
to control the flow of fluid through the gap 22 based on the
diffuser inlet conditions. The axially adjustable ring segment 16
of the adjustable clearance control system 14 may extend at least
partially circumferentially about the airfoil stage 20 In one
embodiment, the axially adjustable ring segment 16 may extend
circumferentially around the airfoil stage 20.
[0021] The axially adjustable ring segment 16 may include a
radially inward contact surface 28 having at least a portion
aligned with the tip 24 of the airfoil 26. In other embodiments,
the radially inward contact surface 28 may be misaligned with the
tip 24 of the airfoil 26. In at least one embodiment, the axially
adjustable ring segment 16 may include a honeycomb seal land 40
coupled to the axially adjustable ring segment 16 The honeycomb
seal land 40 may be formed from a plurality of hollow cavities 42
with walls 44 taking the shape of a honeycomb shape. The honeycomb
shaped walls 44 may have any appropriate shape. The honeycomb seal
land 40 may be formed from any appropriate material capable of
withstanding the environment within a hot gas path in a gas turbine
engine 12. A radially inward contact surface 28 of the honeycomb
seal land 40 may be positioned nonparallel and nonorthogonal
relative to the direction of movement 30 of the axial adjustable
ring segment 16
[0022] In at least one embodiment, the adjustable clearance control
system 14 may include one or more ribs 46 extending radially
outward from the tip 24 of the turbine blade 38 and terminating
before contacting a radially inward contact surface 28 of the
axially adjustable ring segment 16 In at least one embodiment, the
rib 46 may be formed from a tip shroud, as shown in FIG. 4. The rib
46 may have any appropriate configuration and shape. The rib 46 may
also may be formed from any appropriate material capable of
withstanding the environment within a hot gas path in a gas turbine
engine 12.
[0023] During operation, the adjustable clearance control system 14
may enable the axially adjustable ring segment 16 to be moved
axially such that the gap 22 may be reduced or increased. For
instance, the adjustable clearance control system 14 may move the
axially adjustable ring segment 16 to increase the gap 22
independently of other stages 18 to increase tip jet flow when
diffuser inlet conditions cause high OD loading. High OD loading
occurs where a hub strong velocity profile entrains or pulls flow
away from the OD and toward the hub. This causes more diffusion to
occur near the OD wall. Hence higher loading or diffusion which can
lead to flow separation off the OD wall.
[0024] Also, during startup of the gas turbine engine 12, the
adjustable clearance control system 14 may move the axially
adjustable ring segment 16 axially such that the gap 22 may be
increased to prevent tip rubbing and possible damage Once at steady
state operating conditions, the adjustable clearance control system
14 may move the axially adjustable ring segment 16 to reduce the
gap 22 independently of other stages 18 During shutdown of the gas
turbine engine 12, the adjustable clearance control system 14 may
move the axially adjustable ring segment 16 to increase the gap 22
independently of other stages 18 to prevent tip rubbing and
possible damage. In addition, increasing the clearance of the last
stage blade 20 on cold days or on turbine engines 12 that have a
hub strong exit velocity profile may improve engine performance
Under such conditions, the exhaust diffuser 48 may be overloaded on
the outer diamater (OD) due to higher velocities on the hub pulling
flow away Increasing the last stage blade 20 tip gap 22 can inject
high velocity air at the OD to keep the OD flow healthy in the
exhaust diffuser 48.
[0025] The adjustable clearance control system 14 may address the
need to open the clearance on the last blade tip 24 without opening
the clearance on the other blades. Normally, turbine engines 12 are
operated with the tightest possible clearance on every blade.
However, at certain conditions, diffuser performance can be
significantly improved by increasing the last blade tip leakage
jet. Even though opening the gap reduces the blade efficiency, that
loss can be more than offset by the improvement in the downstream
diffuser. One particular situation in which diffuser performance
can be significantly improved is at baseload, which is maximum
power standard day, if the turbine has a hub strong velocity
profile This situation occurs when turbine engines 12 are operated
at higher power levels, which can occur by increasing mass flow,
without changing the turbine design. This situation can also occur
in well designed turbines having a flat velocity profile when they
operate on cold days.
[0026] The foregoing is provided for purposes of illustrating,
explaining, and describing embodiments of this invention
Modifications and adaptations to these embodiments will be apparent
to those skilled in the art and may be made without departing from
the scope or spirit of this invention.
* * * * *