U.S. patent application number 12/044017 was filed with the patent office on 2009-09-10 for gas turbine engine including temperature control device and method using memory metal.
This patent application is currently assigned to SIEMENS POWER GENERATION, INC.. Invention is credited to David A. Little, Hubertus E. Paprotna.
Application Number | 20090226327 12/044017 |
Document ID | / |
Family ID | 41053785 |
Filed Date | 2009-09-10 |
United States Patent
Application |
20090226327 |
Kind Code |
A1 |
Little; David A. ; et
al. |
September 10, 2009 |
Gas Turbine Engine Including Temperature Control Device and Method
Using Memory Metal
Abstract
The gas turbine engine includes a compressor section, a
combustion section downstream from the compressor section and a
turbine section downstream from the combustion section. The turbine
section includes a rotor shaft, a plurality of turbine blades, and
a plurality of discs coupling corresponding ones of the plurality
of turbine blades to the rotor shaft. Each disc has a plurality of
disc cooling fluid passages therein associated with cooling the
turbine blades, and a respective thermal shape memory sleeve is in
at least some of the disc cooling fluid passages. The thermal shape
memory sleeve defines a sleeve throat opening that changes based
upon a temperature to adjust a flow of cooling fluid therethrough.
Part load efficiency may be increased by supplying only the needed
cooling fluid to areas of the blades.
Inventors: |
Little; David A.; (Chuluota,
FL) ; Paprotna; Hubertus E.; (Winter Springs,
FL) |
Correspondence
Address: |
SIEMENS CORPORATION;INTELLECTUAL PROPERTY DEPARTMENT
170 WOOD AVENUE SOUTH
ISELIN
NJ
08830
US
|
Assignee: |
SIEMENS POWER GENERATION,
INC.
Orlando
FL
|
Family ID: |
41053785 |
Appl. No.: |
12/044017 |
Filed: |
March 7, 2008 |
Current U.S.
Class: |
416/96A ;
416/229R |
Current CPC
Class: |
F01D 5/082 20130101;
F05D 2300/505 20130101 |
Class at
Publication: |
416/96.A ;
416/229.R |
International
Class: |
F01D 5/14 20060101
F01D005/14 |
Claims
1. A gas turbine engine comprising: a compressor section; a
combustion section downstream from said compressor section; a
turbine section downstream from said combustion section including a
rotor shaft, a plurality of turbine blades, and a plurality of
discs coupling corresponding ones of the plurality of turbine
blades to the rotor shaft, each disc having a plurality of disc
cooling fluid passages therein associated with cooling the turbine
blades, and a respective thermal shape memory sleeve in at least
some of the disc cooling fluid passages and defining a sleeve
throat opening that changes based upon a temperature to adjust a
flow of cooling fluid therethrough.
2. The gas turbine engine according to claim 1, wherein the thermal
shape memory sleeve comprises a thermal shape memory metal
sleeve.
3. The gas turbine engine according to claim 1, wherein the thermal
shape memory sleeve comprises a thermal shape memory polymer
sleeve.
4. The gas turbine engine according to claim 1, further comprising
a cooling system including a cooler to cool at least a portion of
compressed air from the compressor section and to provide cooling
fluid to the plurality of disc cooling fluid passages.
5. The gas turbine engine according to claim 1, wherein each of the
plurality of turbine blades has a plurality of blade cooling fluid
passages therein coupled in fluid communication with the disc
cooling fluid passages.
6. The gas turbine engine according to claim 1, wherein the sleeve
throat opening has a cylindrically shape.
7. The gas turbine engine according to claim 1, wherein the sleeve
throat opening of the thermal shape memory sleeve is larger when
the temperature is relatively high, and the sleeve throat opening
is smaller when the temperature is relatively low.
8. The gas turbine engine according to claim 1, wherein the sleeve
throat opening of the thermal shape memory sleeve has a shape
transition temperature in a range of 175-400.degree. C.
9. A gas turbine engine comprising: a combustion section; a turbine
section downstream from said combustion section including a rotor
shaft, a plurality of turbine blades, and a plurality of discs
coupling corresponding ones of the plurality of turbine blades to
the rotor shaft, each disc having a plurality of disc cooling fluid
passages therein associated with cooling the turbine blades, and a
respective thermal shape memory metal sleeve in each of the cooling
fluid passages and defining a cylindrical sleeve throat opening
that changes based upon a temperature to adjust a flow of cooling
fluid therethrough.
10. The gas turbine engine according to claim 9, further comprising
a cooling system including a cooler to provide cooling fluid to the
plurality of disc cooling fluid passages.
11. The gas turbine engine according to claim 9, wherein each of
the plurality of turbine blades has a plurality of blade cooling
fluid passages therein coupled in fluid communication with the disc
cooling fluid passages.
12. The gas turbine engine according to claim 9, wherein the
cylindrical sleeve throat opening of the shape memory sleeve is
larger when the temperature is relatively high, and the cylindrical
sleeve throat opening is smaller when the temperature is relatively
low.
13. A method of controlling blade tip clearance in a gas turbine
engine including a compressor section, a combustion section
downstream from the compressor section, and a turbine section
downstream from the combustion section and including a rotor shaft,
a plurality of turbine blades, and a plurality of discs coupling
corresponding ones of the plurality of turbine blades to the rotor
shaft, the method comprising: forming a plurality of disc cooling
fluid passages in each disc for cooling the turbine blades; and
positioning a respective thermal shape memory sleeve in at least
some of the cooling fluid passages to define a sleeve throat
opening that changes based upon a temperature to adjust a flow of
cooling fluid therethrough.
14. The method according to claim 13, wherein the thermal shape
memory sleeve comprises a thermal shape memory metal sleeve.
15. The method according to claim 13, wherein the thermal shape
memory sleeve comprises a thermal shape memory polymer sleeve.
16. The method according to claim 13, further comprising cooling at
least a portion of compressed air from the compressor section to
provide cooling fluid to the plurality of disc cooling fluid
passages.
17. The method according to claim 13, further comprising providing
each of the plurality of turbine blades with a plurality of blade
cooling fluid passages therein coupled in fluid communication with
the disc cooling fluid passages.
18. The method according to claim 13, wherein each the sleeve
throat opening has a cylindrical shape.
19. The method according to claim 13, wherein the sleeve throat
opening of the thermal shape memory sleeve is larger when the
temperature is relatively high, and the sleeve throat opening is
smaller when the temperature is relatively low.
20. The method according to claim 13, wherein the sleeve throat
opening of the thermal shape memory sleeve has a shape transition
temperature in a range of 175-400.degree. C.
Description
FIELD OF THE INVENTION
[0001] The present invention relates to the field of turbine
engines, and, more particularly, to cooling air flow in turbines
and related methods.
BACKGROUND OF THE INVENTION
[0002] A gas turbine typically includes a compressor section that
produces compressed air. Fuel is then mixed with and burned in a
portion of this compressed air in a combustion section having one
or more combustors, thereby producing a hot compressed gas. The hot
compressed gas is then expanded in a turbine section to produce
rotating shaft power. The turbine section typically uses a
plurality of alternating rows of stationary vanes and rotating
blades. Each of the rotating blades has an airfoil portion and a
root portion by which it is affixed to a rotor, e.g. via a disc.
The root portion includes a platform from which the airfoil portion
extends.
[0003] Blades and vanes in gas turbine engines may include cooling
air passages leading to an outer surface of the airfoil that
requires cooling. These cooling air passages are typically located
in specific locations on the airfoil where extremely high
temperatures exists during operation of the engine. Certain regions
of the surface require larger amounts of cooling air than other
areas that require less cooling air. When designing the size of the
cooling air passages, the designer typically sizes the passages to
be able to supply the amount of cooling air to cool the airfoil
surface under the worst case situation of highest possible heat
load. This design temperature is reached under normal operation of
the engine. Also, the heat load varies on surfaces of the airfoil,
so not every surface requires the same amount of cooling airflow.
Thus, the amount of cooling air passing through the passage and
onto the external surface of the airfoil is designed to adequately
cool that area of the airfoil. Thus, no cooling air flow is wasted
and overall engine performance and efficiency is optimized.
[0004] U.S. Pat. No. 6,408,610 issued to Caldwell et al. is
directed to "A METHOD OF ADJUSTING GAS TURBINE COMPONENT COOLING
AIR FLOW", and shows an airfoil that includes a plurality of
cooling holes having a thermal barrier coating applied at various
thicknesses in the holes to provide a desired hole diameter. Under
this method, the size of the cooling air passages can be designed
to provide a desired amount of cooling air flow onto the surface of
the airfoil, depending upon the air pressure within the blade and
around the opening of the cooling air passage, such that a desired
amount of cooling can occur. However, the sizes of the cooling
holes do not vary based upon the operating conditions of the engine
in the region of the specific cooling air passage. So, the size of
the cooling air passage may be smaller than needed, resulting in
less cooling air flow than required, or larger than needed,
resulting in more cooling air flow than required. Either way, the
engine performance or efficiency is reduced.
[0005] U.S. Pat. No. 6,416,279 issued to Weigand et al entitled "A
COOLED GAS TURBINE COMPONENT WITH ADJUSTABLE COOLING" in which the
cooling air passage includes different means to vary the amount of
cooling air flow during engine operation. In one method, a
restrictor having an opening of specific size is placed in the
cooling air passage to regulate the cooling air flow during engine
operation. In this method, the size of the restrictor cannot be
changed during engine operation. In another method, a control
system is used and includes a temperature sensor and a control
valve, where the control valve regulates an amount of cooling air
flow based upon a value from the temperature sensor.
[0006] U.S. Pat. No. 6,485,255 issued to Care et.al. entitled "A
COOLING AIR FLOW CONTROL DEVICE FOR A GAS TURBINE ENGINE" is
directed to a single shape memory metal valve disposed in a cooling
passage upstream of the many cooling air passages that open out
onto the outer surface of the airfoil. In the Care et al. device,
the valve varies the air flow depending upon temperature, but all
of the cooling air passages opening onto the airfoil surfaces are
controlled by this single valve. The passages exposed to the
hottest surface of the airfoil are regulated by the same valve and
supply airflow as the openings exposed to the coolest airfoil
surface so that cooling may be insufficient at some locations.
[0007] U.S. Pat. No. 7,241,107 issued to Spanks et al. entitled
"GAS TURBINE AIRFOIL WITH ADJUSTABLE COOLING AIR FLOW PASSAGES"
discloses an airfoil for a gas turbine engine, wherein the airfoil
includes a plurality of cooling air passages to supply cooling air
to an external surface of the airfoil. The cooled surface of the
airfoil has a critical temperature in which any cooled surface of
the airfoil should not exceed. The cooling air passages have a
coating applied within the passages, and the coating is made of a
material that has an oxidizing property such that the material
oxidizes away and opens the passage to more flow when exposed to a
temperature above the critical temperature. When the airfoil
surface is not properly cooled by a flow passing through the
passage, the material oxidizes away until the size of the passage
increases to allow for the proper amount of cooling air to flow to
cool the airfoil. Each passage is located in a different part of
the airfoil that requires more or less cooling flow, and each
passage will oxidize until the size of the passage is large enough
to allow for the proper amount of cooling flow.
SUMMARY OF THE INVENTION
[0008] In view of the foregoing background, it is therefore an
object of the present invention to provide a gas turbine engine
with efficient controlled temperature regulation of the turbine
section blades.
[0009] This and other objects, features, and advantages in
accordance with the present invention are provided by a gas turbine
engine including a compressor section, a combustion section
downstream from the compressor section and a turbine section
downstream from the combustion section and including a rotor shaft,
a plurality of turbine blades, and a plurality of discs coupling
corresponding ones of the plurality of turbine blades to the rotor
shaft. Each disc has a plurality of disc cooling fluid passages
therein associated with cooling the turbine blades, and a
respective thermal shape memory sleeve is in at least some of the
disc cooling fluid passages. The thermal shape memory sleeve
defines a sleeve throat opening that changes based upon a
temperature to adjust a flow of cooling fluid therethrough.
[0010] Thus, the approach addresses blade tip clearance control
that may be needed when the turbine inlet temperature is maintained
at a high level during part load operation of a gas turbine engine,
which may be done, for example, to reduce CO emissions. To that
end, features may include ensuring a minimum blade tip clearance
under such conditions.
[0011] Conventionally, during part load conditions when the rotor
cooling air temperature is decreased to maintain the integrity of
the blade tips, the blades are overcooled. The present approach may
more efficiently cool the blades at the cooler rotor cooling air
temperature (discs set tip clearances) by reducing the throat area
of the memory metal sleeves to reduce the cooling flows to those
required at the cooler rotor air temperature and turbine inlet
temperature. Accordingly, part load efficiency may be improved by
supplying only needed cooling air to all areas of the blades.
[0012] The thermal shape memory sleeve may be a thermal shape
memory metal sleeve or a thermal shape memory polymer sleeve, for
example. A cooling system may be included and have a cooler to cool
at least a portion of compressed air from the compressor section
and to provide cooling fluid to the plurality of disc cooling fluid
passages. Each of the plurality of turbine blades may have a
plurality of blade cooling fluid passages therein coupled in fluid
communication with the disc cooling fluid passages.
[0013] The sleeve throat opening may have a cylindrical shape. The
sleeve throat opening of the thermal shape memory sleeve may be
larger when the temperature is relatively high, and the sleeve
throat opening may be smaller when the temperature is relatively
low. The sleeve throat opening of the thermal shape memory sleeve
may have a shape transition temperature in a range of
175-400.degree. C.
[0014] A method aspect is directed to controlling blade tip
clearance in a gas turbine engine including a compressor section, a
combustion section downstream from the compressor section, and a
turbine section downstream from the combustion section and
including a rotor shaft, a plurality of turbine blades, and a
plurality of discs coupling corresponding ones of the plurality of
turbine blades to the rotor shaft. The method includes forming a
plurality of disc cooling fluid passages in each disc for cooling
the turbine blades, and positioning a respective thermal shape
memory sleeve in at least some of the cooling fluid passages to
define a sleeve throat opening that changes based upon a
temperature to adjust a flow of cooling fluid therethrough.
BRIEF DESCRIPTION OF THE DRAWINGS
[0015] FIG. 1 is partly schematic cross-sectional view of an
example of a portion of a turbine engine in accordance with the
present invention.
[0016] FIG. 2 is a more detailed cross-sectional view of a rotor
disc of the turbine engine of FIG. 1.
[0017] FIG. 3 is an enlarged detailed cross-sectional view of a
portion of the rotor disc of FIG. 2 within the dashed circle B.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
[0018] The present invention will now be described more fully
hereinafter with reference to the accompanying drawings, in which
preferred embodiments of the invention are shown. This invention
may, however, be embodied in many different forms and should not be
construed as limited to the embodiments set forth herein. Rather,
these embodiments are provided so that this disclosure will be
thorough and complete, and will fully convey the scope of the
invention to those skilled in the art. Like numbers refer to like
elements throughout.
[0019] Referring initially to FIG. 1, a gas turbine engine 10 in
accordance with features of the present invention will now be
described. The turbine 10 includes a turbine section 11, a
compressor section 12 and a combustion section 15. Turbine engines,
such as single shaft industrial gas turbines, are designed to
operate at a constant design turbine inlet temperature under any
ambient air temperature (i.e., the compressor inlet temperature).
This design turbine inlet temperature allows the engine to produce
maximum possible power, known as base load. Any reduction from the
maximum possible base load power is referred to as part load
operation. In other words, part load entails all engine operation
from 0% to 99.9% of base load power.
[0020] Part load operation may result in the production of high
levels of carbon monoxide (CO) during combustion. One known method
for reducing part load CO emissions is to bring the combustor exit
temperature or the turbine inlet temperature near that of the base
load design temperature. Also, some of the compressor exit air 14
from the combustion section 15 may be used to cool the stationary
support structure 16 of the turbine near the first row of blades
20a. The stationary support structure 16 can include the outer
casing, blade rings, and ring segments. In addition, some
compressed air may be piped directly out of the compressor section
12 through piping 19a (additional pipes not shown). This compressor
bleed air is used to cool the stationary support structure 16 near
the second, third and fourth rows of blades 20b, 20c, 20d and is
supplied through piping 19b, 19c, 19d.
[0021] The support structure 16 may contract or shrink in radius
when exposed to the cooler compressor exit and bleed air. But, at
the same time, the temperature of the hot gas leaving the combustor
18 and flowing over the turbine blades 20 (illustratively 20a, 20b,
20c, 20d) is kept at a high level, causing a constant radially
outward thermal expansion of the blades 20.
[0022] The expansion of the blades 20 along with the shrinkage of
the support structure 16 reduces the clearance C between the tips
21 of the blades 20 and the surrounding support structure 16,
commonly referred to as the blade tip clearance C. While the
clearance C is shown between the fourth row of blades 20d and the
adjacent support structure 16, similar clearances C exist between
the first, second and third rows of blades 20a, 20b, 20c and the
stationary support structure 16. A minimal blade tip clearance C
should be maintained so that the blades 20 do not rub against the
support structure 16.
[0023] In this gas turbine engine 10, compressor exit air 14 from
the combustion section 15 can be used to cool at least the turbine
rotor 50, discs 52, and blades 20. The compressor exit air may
routed out of the engine, passed through a cooling fluid loop 22,
including a cooler 24, and is ultimately redelivered to the engine
at a substantially constant design cooling air return temperature.
The design temperature is held substantially constant so that the
discs 52 and blades 20 temperatures are held substantially
constant, e.g. to maintain the life of the discs 52 and blades
20.
[0024] The design cooling return temperature can be specific to a
particular engine design. For example, in the Siemens W501G engine,
the design cooling air return temperature is about 350.degree.
Fahrenheit for loads below 90%. When load reaches 90%, the
temperature of the rotor cooling air can be increased to about
405.degree. Fahrenheit to about 480.degree. Fahrenheit. The rotor
cooling temperature is increased at loads of 90% and above as an
active way to reduce tip clearances C at full load to maximize
engine power and efficiency. When rotor cooling air return
temperature increases, the discs 52 and blades 20 expand radially
outward, causing the clearance C between the tips 21 of the blades
20 and the nearby stationary support structure 16 to decrease. The
smaller clearance C means less losses and, thus, more power
extraction for the same fuel input, thereby increasing efficiency.
Further, the rotor cooling air return temperature can be increased
above 90% load because, by the time the engine reaches that level,
most of the stationary components 16 of the engine have thermally
grown to their final shapes. Thus, distortion and ovalization,
which can cause blade tip rubbing, are minimized.
[0025] The Siemens Westinghouse W501G turbine engine is only one
example of a design cooling air return temperature. In other
engines, the design cooling air return temperature can range from
about 350.degree. Fahrenheit to about 750.degree. Fahrenheit.
[0026] Whatever the range, the design cooling air return
temperature can be supplied at a substantially constant temperature
so that the rotor 50 and the discs 52 are always at substantially
constant metal temperatures. However, the temperature of the
compressor exit air 14 often exceeds the design cooling air return
temperature. Therefore, to provide cooling air at the design
temperature, a portion of compressor exit 14 air may be bled from
the Combustion section 15 and cooled to the appropriate temperature
in the external cooling loop 22, which is configured to reduce the
cooling air to the design cooling air return temperature. The
external cooling loop 22 can also include heat exchanger devices as
well as valves for controlling the quantity of air passing through
or bypassing the heat exchanger devices so as to achieve the design
cooling air return temperature. Once treated, the cooled air can be
returned to the engine to cool at least the rotor 50 and discs 52
at the substantially constant design cooling air return
temperature.
[0027] As discussed above, features of the turbine engine 10 relate
to ensuring adequate blade tip clearance C under part load
conditions. Such features may include reducing the cooling air
return temperature to below the design temperature level to
maintain a minimum acceptable clearance C between a blade tip 21
and surrounding stationary support structure 16. For example, the
minimum acceptable clearance C may be about 1 millimeter or about
0.040 inches. Because the temperature of the cooling return air may
be lower than the design temperature, the discs 52 and blades 20
will tend to shrink when exposed to the cooling return air. In
spite of the expansion of the turbine blades 20 due to the passing
high temperature gases as well as the shrinkage of the stationary
support structure 16 due to cooler compressor exit air 14 and
compressor bleed air temperatures, as described earlier, an
adequate blade tip clearance C may be maintained with the cooler
temperature of the cooling air return causing the discs 52 and
blades 20 to shrink, thus widening the gap between the tip 21 of
the blade 20 and the nearby support structure 16 at part load.
[0028] The turbine section 11, downstream from the combustion
section 15, includes the rotor shaft 50, the plurality of turbine
blades 20, and a plurality of discs 52 coupling corresponding ones
of the plurality of turbine blades to the rotor shaft. Referring
additionally to FIGS. 2 and 3, each disc 52 has a plurality of disc
cooling fluid passages 54 therein associated with cooling the
turbine blades 20, and a respective thermal shape memory sleeve 56
is in at least some of the disc cooling fluid passages. The thermal
shape memory sleeve 56 defines a sleeve throat opening A that
changes based upon a temperature to adjust a flow of cooling fluid
60 therethrough.
[0029] Tip clearance control is obtained by operating with rotor
cooling air return temperature low at lower loads (transient
conditions, i.e. loading and unloading the turbine engine) to keep
the blade tips from rubbing, then by increasing rotor cooling air
temperature near base load. This increase tightens tip clearances
(once the stationary components have stabilized) primarily because
of the increase in radii of the discs. The increase in blade
lengths is a secondary contributor to the reduction in tip
clearance.
[0030] Blading is designed at base load conditions where both
turbine inlet temperature and rotor cooling air are maximized.
These worst case conditions will occur continuously across the
ambient temperature range when the engine is on base load control
(100% load). Cooling flow consumption is regulated, e.g. optimized
or minimized, everywhere on each airfoil at base load conditions.
However, in conventional systems, during part load conditions when
the rotor cooling air temperature has been decreased to maintain
the integrity of the blade tips, the blades may be overcooled.
[0031] The present approach, including the use of thermal shape
memory sleeves 56 in disc cooling fluid passages 54, may more
efficiently cool the blades at the cooler rotor cooling air
temperature (discs set tip clearances) by reducing the throat area
of the thermal shape memory sleeves 56 to reduce the cooling flows
to those needed at the cooler rotor air temperature and turbine
inlet temperature. Part load efficiency may be made more efficient
by supplying only needed cooling air to all areas of the blades
20.
[0032] The thermal shape memory sleeve 56 may be a thermal shape
memory metal sleeve or a thermal shape memory polymer sleeve, for
example. A shape memory alloy (SMA, also known as a smart alloy or
memory metal) is an alloy that "remembers" its geometry. After a
sample of SMA has been deformed from its original crystallographic
configuration, it regains its original geometry by itself during
heating (one-way effect) or, at higher ambient temperatures, simply
during unloading (pseudo-elasticity or superelasticity). These
extraordinary properties are due to a temperature-dependent
martensitic phase transformation from a low-symmetry to a highly
symmetric crystallographic structure. Those crystal structures are
known as martensite (at lower temperatures) and austenite (at
higher temperatures).
[0033] Examples of types of SMAs include
copper-zinc-aluminum-nickel, copper-aluminum-nickel, and
nickel-titanium (NiTi) alloys. Other SMAs may include Ag--Cd,
Au--Cd, Cu--Al--Ni, Cu--Sn, Cu--Zn, Cu--ZnS--i, Fe--Pt, Mn--Cu,
Fe--Mn--Si, Pt alloys, Co--Ni--Al, Co--Ni--Ga, Ni--Fe--Ga and
Ti--Pd, for example.
[0034] As discussed above, a cooling loop or system 22 may be
included and have a cooler 24 to cool at least a portion of
compressed air 14 from the compressor section 12 and to provide
cooling fluid 60 to the plurality of disc cooling fluid passages
54. Each of the plurality of turbine blades 20 may have a plurality
of blade cooling fluid passages 58 (FIG. 2) therein coupled in
fluid communication with the disc cooling fluid passages 54, as
would be appreciated by those skilled in the art.
[0035] The sleeve throat opening A may have a cylindrically shape.
The sleeve throat opening A of the thermal shape memory sleeve may
be larger when the temperature is relatively high, as indicated by
the dashed line in FIG. 3, and the sleeve throat opening A may be
smaller when the temperature is relatively low. The sleeve throat
opening A of the thermal shape memory sleeve 56 may have a shape
transition temperature in a range of 175-400.degree. C., for
example. (0036] A method aspect is directed to controlling blade
tip clearance C in a gas turbine engine 10 including a compressor
section 12, a combustion section 15 downstream from the compressor
section, and a turbine section 11 downstream from the combustion
section and including a rotor shaft 50, a plurality of turbine
blades 20, and a plurality of discs 52 coupling corresponding ones
of the plurality of turbine blades to the rotor shaft. The method
includes forming a plurality of disc cooling fluid passages 54 in
each disc 52 for cooling the turbine blades 20, and positioning a
respective thermal shape memory sleeve 56 in at least some of the
cooling fluid passages to define a sleeve throat opening A that
changes based upon a temperature to adjust a flow of cooling fluid
60 therethrough.
[0036] Thus, features of the approach address blade tip clearance C
control that may be needed when the turbine inlet temperature is
maintained at a high level during part load operation of a gas
turbine engine 10, which may be done, for example, to reduce CO
emissions. To that end, features of the approach may include
ensuring a minimum blade tip clearance C under such conditions. The
present approach may more efficiently cool the blades at the cooler
rotor cooling air temperature (discs set tip clearances) by
reducing the throat area of the thermal shape memory sleeves 56 to
reduce the cooling flows to those required at the cooler rotor air
temperature and turbine inlet temperature. Accordingly, part load
efficiency may be improved by supplying only needed cooling air to
all areas of the blades.
[0037] Many modifications and other embodiments of the invention
will come to the mind of one skilled in the art having the benefit
of the teachings presented in the foregoing descriptions and the
associated drawings. Therefore, it is understood that the invention
is not to be limited to the specific embodiments disclosed, and
that modifications and embodiments are intended to be included
within the scope of the appended claims.
* * * * *