U.S. patent application number 13/604729 was filed with the patent office on 2013-04-04 for gas turbine.
This patent application is currently assigned to Hitachi, Ltd.. The applicant listed for this patent is Hidetoshi Kuroki, Hayato Maekawa, Hidetaro Murata, Atsushi SANO, Hironori Tsukidate. Invention is credited to Hidetoshi Kuroki, Hayato Maekawa, Hidetaro Murata, Atsushi SANO, Hironori Tsukidate.
Application Number | 20130084162 13/604729 |
Document ID | / |
Family ID | 46750237 |
Filed Date | 2013-04-04 |
United States Patent
Application |
20130084162 |
Kind Code |
A1 |
SANO; Atsushi ; et
al. |
April 4, 2013 |
Gas Turbine
Abstract
A gas turbine is provided that can suppress an increase in the
supplied amount of cooling air in clearance adjustment through
casing cooling. The gas turbine 101 includes a turbine casing 19
enclosing a turbine shaft 105, the turbine casing 19 including a
cooling air header 21 and a cooling passage 24; and an exhaust
diffuser 113 connected to the exhaust side of the casing 19, the
exhaust diffuser 113 including an exhaust diffuser cooling passage
26. A plate 22 formed with a plurality of impingement cooling holes
28 is installed inside the cooling air header, and a route is
formed which allows cooling air introduced from the impingement
cooling holes to flow from the casing cooling passage to the
diffuser cooling passage.
Inventors: |
SANO; Atsushi; (Hitachinaka,
JP) ; Kuroki; Hidetoshi; (Hitachi, JP) ;
Tsukidate; Hironori; (Hitachi, JP) ; Murata;
Hidetaro; (Hitachi, JP) ; Maekawa; Hayato;
(Hitachi, JP) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
SANO; Atsushi
Kuroki; Hidetoshi
Tsukidate; Hironori
Murata; Hidetaro
Maekawa; Hayato |
Hitachinaka
Hitachi
Hitachi
Hitachi
Hitachi |
|
JP
JP
JP
JP
JP |
|
|
Assignee: |
Hitachi, Ltd.
Tokyo
JP
|
Family ID: |
46750237 |
Appl. No.: |
13/604729 |
Filed: |
September 6, 2012 |
Current U.S.
Class: |
415/116 |
Current CPC
Class: |
F05D 2260/201 20130101;
F01D 11/24 20130101 |
Class at
Publication: |
415/116 |
International
Class: |
F01D 25/14 20060101
F01D025/14 |
Foreign Application Data
Date |
Code |
Application Number |
Sep 29, 2011 |
JP |
2011-213681 |
Claims
1. A gas turbine comprising: a casing enclosing a turbine shaft,
the casing including a cooling air header and a casing cooling
passage; and an exhaust diffuser connected to an exhaust side of
the casing, the exhaust diffuser including an exhaust diffuser
cooling passage; wherein a plate formed with a plurality of
impingement cooling holes is installed inside the cooling air
header, and a route is formed which allows cooling air introduced
from the impingement cooling holes to flow from the casing cooling
passage to the diffuser cooling passage.
2. The gas turbine according to claim 1, wherein the plate is
disposed to impingement-cool a casing located at a position
corresponding to a front-stage side of turbine stages composed of a
plurality of stages.
3. The gas turbine according to claim 1, wherein the casing has a
cover isolating the cooling air header from outside space.
4. The gas turbine according to claim 3, wherein the plate is
attached inside the cover and the cover is configured to be
attachable to and detachable from the casing integrally with the
plate.
5. The gas turbine according to claim 1 wherein the casing includes
an upper-half casing and a lower-half casing separated from each
other, the upper-half casing and the lower-half casing being joined
to each other via respective flanges thereof, a plurality of the
plates are installed along a circumferential direction of the
casing, a plurality of systems are provided which each supply
cooling air to each of spaces defined by the plates, and a flow
rate control device for regulating a flow rate of cooling air is
mounted in the system connected to a space defined by a plate
located on the flange side among the plurality of plates.
6. The gas turbine according to claim 1 wherein the casing includes
an upper-half casing and a lower-half casing separated from each
other, the upper-half casing and the lower-half casing being joined
to each other via respective flanges thereof, a plurality of the
plates are installed along a circumferential direction of the
casing, a plurality of systems are provided which each supply
cooling air to each of spaces defined by the plates, an orifice is
attached to each of the systems, and respective diameters of the
orifices are set based on, at circumferential positions of the gas
turbine, a clearance value between a rotating body and a stationary
body in a stationary state of the gas turbine, and a relationship
between the size of a diameter of the orifice and an amount of
deformation of the casing.
Description
BACKGROUND OF THE INVENTION
[0001] 1. Field of the Invention
[0002] The present invention relates to a gas turbine in which are
adjusted clearances between a casing enclosing a turbine shaft,
turbine rotor blades, etc., and the turbine rotor blades.
[0003] 2. Description of the Related Art
[0004] Gas turbines are configured such that a rotor (a rotating
body) is enclosed inside a casing (a stationary body). Turbine
rotor blades are installed on the outer circumferential portion of
the rotor. A clearance exists between the tip (the outermost
circumferential side) of the turbine rotor blade and a shroud
mounted on the inner circumference of the casing. When high
temperature and high pressure mainstream gas passes through the
clearance, a leakage loss occurs, which results in performance
degradation. Thus, it is desirable that the clearance between the
turbine rotor blade and the shroud be small in terms of improvement
in turbine performance.
[0005] On the other hand, too a small clearance between the tip of
the rotor blade and the shroud may cause the tip of the rotor blade
and the shroud to come into contact with each other and they may be
broken. The clearance is varied during the operation of the turbine
due to the thermal expansion and centrifugal expansion of the
rotor, the casing and the like. The clearance is determined at the
time of assembly of the casing (at the time of start-up) in order
to prevent breakage attributable to the contact under the entire
operating conditions.
[0006] In general, this minimum clearance appears in the process of
the start-up in industrial gas turbines. This is because the casing
is harder to be heated up than the rotor due to a difference in
heat capacity therebetween. If the clearance is minimized at times
other than during steady operation, the clearance has to be
designed so that contact may not occur at times other than during
the steady operation. The clearance during rated operation is
larger than that in the middle of start-up. Thus, the turbine is
operated while having an undesirable excessively large
clearance.
[0007] To avoid the undesirable excessively large clearance, some
gas turbines shown in e.g. JP-2008-196490-A have manifolds
installed on the outer circumference of the casing to cool the
outer circumference of the casing by use of air flow. Thus, the
thermal expansion of the casing is suppressed to adjust the
clearance.
SUMMARY OF THE INVENTION
[0008] High-temperature components of a gas turbine are subjected
to temperature control by supplying thereto air extracted from a
compressor or cooling air from a separate-placement blower in view
of high-temperature strength, thermal deformation and material
costs. The high-temperature components to be cooled include a
combustor, turbine blades and an exhaust diffuser.
[0009] Also casing cooling for improving the gas turbine
performance needs the supply of cooling air, for which a blower is
generally used. Since the temperature of the casing reaches as high
as several hundred degrees centigrade, it is possible to use the
compressor extraction air having temperature lower than such casing
temperature.
[0010] Power is needed to supply the compressor extraction air or
the cooling air from the blower or the like. If a casing cooing
system is simply added, the consumption of cooling air is increased
and also the power for supplying the cooling air is increased.
Thus, an improvement in performance resulting from clearance
adjustment is partially offset by the increased power.
[0011] It is an object of the present invention to provide such a
gas turbine that an increase in the amount of cooling air to supply
is suppressed upon clearance adjustment through casing cooling.
[0012] According to an aspect of the present invention, there is
provided a gas turbine including a casing enclosing a turbine
shaft, the casing including a cooling air header and a casing
cooling passage; and an exhaust diffuser connected to an exhaust
side of the casing, the exhaust diffuser including an exhaust
diffuser cooling passage. A plate formed with a plurality of
impingement cooling holes is installed inside the cooling air
header, and a route is formed which allows cooling air introduced
from the impingement cooling holes to flow from the casing cooling
passage to the diffuser cooling passage.
[0013] The present invention provides a gas turbine in which a
turbine casing can be cooled by use of a slightly increased amount
of cooling air.
BRIEF DESCRIPTION OF THE DRAWINGS
[0014] FIG. 1 is a partial cross-sectional view of a gas turbine
according to an embodiment of the present invention.
[0015] FIG. 2 is a conceptual diagram showing a gas turbine
embodying the present invention.
[0016] FIG. 3 is a partial cross-sectional view of the gas turbine
according to the embodiment of the present invention.
[0017] FIG. 4 is a characteristic diagram of a rotor blade tip
clearance of a conventional gas turbine.
[0018] FIG. 5 is a characteristic diagram of a rotor blade tip
clearance of the gas turbine according to the embodiment of the
present invention.
[0019] FIG. 6 is an enlarged view of an impingement cooling plate
shown in FIG. 1.
[0020] FIG. 7 is a conceptual view showing a thermal deformation
state of a casing.
[0021] FIG. 8 is a diagram of a gas turbine cooling system
according to an embodiment of the present invention.
[0022] FIG. 9 is a conceptual view showing a state where the center
of the casing is not coincident with the center of a turbine
shaft.
[0023] FIG. 10 is a diagram of a gas turbine cooling system
according to another embodiment of the present invention.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
[0024] In the invention described in JP-2008-196490-A, the
impingement manifolds are installed on the outer surface of the
casing so as to impingement-cool the casing. Cooling air used for
the impingement cooling is led from a blower, impingement-cools the
casing, and then is discharged to the atmosphere. Thus, the total
used amount of the cooling air for the gas turbine is increased by
the amount of cooling air used for cooling the casing.
[0025] High-temperature components including an exhaust diffuser in
a gas turbine are usually cooled by air extracted from a compressor
or air from a separate-placement blower. Power is needed to supply
cooling air by use of a compressor or a blower. If a casing is
cooled for clearance adjustment and an amount of cooling air is
increased, also power used to supply cooling air is increased.
Therefore, an improvement in performance resulting from clearance
adjustment is partially offset by the increased power. Thus, if the
addition of a casing cooling system is assumed, it is desired that
the casing can be cooled by the less increased amount of cooling
air.
[0026] Additionally, it is desired that a clearance between a
turbine rotor blade and a shroud be small as much as possible.
However, if the clearance is too small, breakage may be likely to
occur when the rotor blade and the shroud come into contact with
each other. For this reason, for example, a combination of a
honeycomb seal and a shroud fin is used on a rear stage side to
permit the contact. This absorbs the influence of manufacturing
tolerance and an influence of the deformation of the casing,
thereby keeping the clearance small. A turbine front stage side
where the temperature of mainstream gas is high cannot use the
honeycomb seal because of a heat resistance problem. Therefore, a
margin is provided at the clearance located at the tip of the rotor
blade to avoid the contact due to the influence of the
manufacturing tolerance or of the deformation of the casing.
[0027] As described above, the front stage side of the gas turbine
is likely to increase the clearance according to the provision of
the margin compared with the rear stage side. Therefore, it is
desired that a clearance adjustment amount of the front stage side
can be more enlarged.
[0028] Air extracted from a compressor is led as cooling air for
high-temperature components toward a turbine side via extraction
pipes installed on the outside of the gas turbine. An increase in
the number of the extraction pipes leads to an increased cost.
Therefore, the number of the extraction stages is limited to
several stages such as, for example, the intermediate stages and
rear stages of the turbine. The number of the extraction stages is
generally smaller than the number of the turbine stages to be
cooled. The use of excessive high-pressure air leads to an
increased loss. However, the number of the extraction stages is
limited; therefore, a portion exists to which cooling air is
supplied at a slightly excessive pressure. This excessive pressure
is regulated to an appropriate pressure; therefore, an orifice or
the like causes a pressure loss. However, it is desired to avoid
pressure regulation performed by such an orifice because of a
pressure waste.
[0029] The present invention will be described using embodiments
hereinafter. The present invention provides a gas turbine in which
a turbine casing can be cooled by the slightly increased flow of
cooling air and preferable clearance control is executable. In
addition, the present invention provides a gas turbine that allows
a reduction in the distortion of an exhaust diffuser. First, the
overall system configuration of the gas turbine will be described
with reference to FIG. 2.
[0030] FIG. 2 is a configurational diagram of an overall system of
a gas turbine embodying the present invention.
[0031] The gas turbine 101 mainly includes a compressor 102, a
combustor 103 and a turbine 104. The compressor 102 compresses
ambient air 111 to generate compressed air 106 and supplies the
compressed air 106 thus generated to the combustor 103. The
combustor 103 mixes fuel with the compressed air 106 generated by
the compressor 102 for combustion to generate combustion gas 107
and discharges it to the turbine 104.
[0032] The turbine 104 uses the combustion gas 107 increased in the
energy of the compressed air and discharged from the combustor 103
to allow a turbine shaft 105 to generate rotational force. The
rotational force of the turbine shaft 105 drives equipment 109
(driven machines such as a generator, a pump, and a screw)
connected to the gas turbine 101. The energy of the combustion gas
107 is recovered by the turbine 104 and then the combustion gas 107
is discharged as exhaust gas 112 from the turbine 104 via the
exhaust diffuser 113.
[0033] Air extracted from the compressor 102 or air from a blower
(not shown) is supplied as cooling air 110 to the turbine 104 or
the exhaust diffuser 113 not via the combustor 103.
[0034] FIG. 3 is a partial cross-sectional view of the gas turbine.
There are shown a first-stage stator blade 1, a first-stage rotor
blade 2, a second-stage stator blade 3, a second-stage rotor blade
4, a third-stage stator blade 5, a third-stage rotor blade 6, a
fourth-stage stator blade 7 and a fourth-stage rotor blade 8.
Reference numeral 9 denotes a flow direction of the combustion gas
107 in the turbine.
[0035] The first-stage rotor blade 2 is connected to the outer
circumference of a first-stage wheel 10. In addition, the
first-stage wheel 10, a second-stage wheel 11 to which a
second-stage rotor blade 4 is connected, a compressor rotor 20,
which is a constituent element of the compressor 102, and a spacer
14 are stacked by means of stacking bolts. In this way, a
high-pressure side turbine shaft 105 is configured. The third-stage
rotor blade 6 is connected to the outer circumference of a
third-stage wheel 12. The third-stage wheel 12, a four-stage wheel
13 to which the fourth-stage rotor blade 8 is connected, a rotor
connected to the equipment 109 such as a generator, and a spacer 14
are stacked by means of stacking bolts. In this way, a low-pressure
side turbine shaft 105 is configured. The turbine shaft 105
recovers the energy of the combustion gas 107 discharged from the
combustor 103 by use of the first-stage rotor blade 2, the
second-stage rotor blade 4, the third-stage rotor blade 6 and the
fourth-stage rotor blade 8. In addition, the turbine shaft 105
drives the compressor 102 and the equipment 109 connected to an end
portion of the turbine shaft.
[0036] The turbine shaft 105 is enclosed by a turbine casing 19.
The first-stage stator blade 1, the second-stage stator blade 3,
the third-stage stator blade 5, the fourth-stage stator blade 7, a
first-stage shroud 15, a second-stage shroud 16, a third-stage
shroud 17 and a fourth-stage shroud 18 are connected to the inner
circumferential side of the turbine casing 19. Further, diaphragms
27 are connected to the inner circumferential side of the
second-stage stator blade 3 and of the forth-stage stator blade
7.
[0037] Clearances are provided between the first-stage rotor blade
2 and the first-stage shroud 15, between the second-stage rotor
blade 4 and the second-stage shroud 16, between the third-stage
rotor blade 6 and the third-stage shroud 17, between the
fourth-stage rotor blade 8 and the fourth-stage shroud 18, and
between the spacers 14 and the corresponding diaphragms 27. The
clearances serve as an interface between a stationary body and a
rotating body.
[0038] The clearances are each varied depending on the operating
conditions of the gas turbine. FIG. 4 shows a variation trend of a
clearance in a conventional gas turbine. Immediately after
start-up, the turbine shaft 105 is first increased in rotation rate
so that it is radially expanded by centrifugal force to reduce the
clearance. Thereafter, the mainstream gas is increased in
temperature so that the turbine shaft 105, the shrouds 15, 16, 17,
18, and the turbine casing 19 are thermally expanded. The turbine
shaft 105 is expanded radially outwardly and the shrouds 15, 16,
17, 18 are expanded radially inwardly. Thus, the clearance is
reduced. The turbine casing 19 is expanded radially outwardly to
enlarge the clearance. In general, the turbine shaft 105 and the
shrouds 15, 16, 17, 18 are likely to increase in temperature
compared with the turbine casing 19. Therefore, the clearance is
minimized before the turbine is thermally stabilized, specifically,
approximately at the time of reaching a rated load. Thus, the
clearance during steady operation is greater than the minimum
clearance.
Embodiment
[0039] A casing cooling structure is described with reference to
FIG. 1. FIG. 1 is an enlarged view of the turbine casing 19. A
cooling air header 21 is installed on a front-stage side outer
circumferential portion of the turbine casing 19 so as to form an
annular space. Impingement cooling plates 22 having a division
structure are annularly installed inside the cooling air header 21.
Further, the cooling air header 21 is isolated from space outside
the cooling air header 21 by cooling air header covers 23 to form
the annular space. The impingement cooling plates 22 and the
cooling air header covers 23 are plurally installed along the
circumferential direction of the casing 19. A cooling air pipe is
connected to each of the cooling air header covers 23. Cooling
passages 24 are connected to an end face of the cooling air header
21. The cooling passages 24 extend inside the turbine casing 19
toward an axially rear stage side. The cooling passages 24 each
have a generally circular cross-section and are intermittently
arranged in a circumferential direction.
[0040] Cooling air 110 used to cool the exhaust diffuser 113 is
generally led from the cooling air pipe to the cooling air header
21. The cooling air 110 is jetted as jet flows from impingement
holes 28 formed in the impingement cooling plate 22 installed in
the annular cooling air header 21, and impingement-cools the
turbine casing 19. Thereafter, the cooling air 110 flows in the
cooling passage 24 toward the rear stage side in the axial
direction of the turbine shaft. The cooling passages 24 are
connected to respective diffuser cooling passages 26 via
corresponding connection holes 25. The cooling air flowing inside
the cooling passages 24 is supplied to the exhaust diffuser cooling
passages 26 to cool the exhaust diffuser 113.
[0041] FIG. 6 is an enlarged view of the impingement cooling plate
22 shown in FIG. 1. The annular impingement cooling plate 22 is
formed with a plurality of impingement holes 28. The impingement
holes 28 are formed at least in a surface opposed to the outer
circumferential surface of the casing. Cooling air 110 fed from the
cooling air header cover 23 is jetted from the plurality of
impingement holes 28. The impingement cooling air 110a having been
jetted from the impingement holes 28 impinges the outer surface of
the casing opposed to the impingement cooling plate 22. This
impingement jet cools the casing from the outer circumferential
side thereof.
[0042] FIG. 5 shows clearance characteristics of the gas turbine
according to the present embodiment. The execution of casing
cooling can reduce the deformation amount of the turbine casing 19
which expands radially outwardly. Consequently, a difference
between the minimum clearance during the process from the start-up
to the steady state and the clearance in the steady state is
reduced compared with the case where the casing is not cooled.
Thus, the clearance in the steady state can be kept small compared
with the conventional clearance. In this case, the minimum
clearance can be made nearly equal to the conventional clearance;
therefore, it is possible to improve performance without impairing
the reliability of the gas turbine.
[0043] It is difficult to use, on the front-stage side of the gas
turbine, a seal structure capable of following clearance
variations. This is because of the following reasons. To apply a
labyrinth seal to the front-stage side of the gas turbine, a shroud
is needed to be formed at a blade tip. However, the formation of
the shroud increases the weight of a blade end face, which
excessively increases the stress of the blade. Further, it is
difficult to use a seal structure permitting contact with a
honeycomb seal or the like because of a problem with heat
resistance. To suppress a leakage loss, it is necessary to keep the
clearance small. For breakage prevention, however, design with a
margin has to be done to some extent. By contrast, since the
mainstream gas on the rear-stage side has low temperature, a
honeycomb seal capable of permitting such contact can be applied to
the rear-stage side. Thus, design with a small margin can be done,
that is, the clearance can be designed to be small in size. As
described above, since the clearance on the front-stage side tends
to increase excessively, it is desirable that the clearance
adjustment amount on the front-stage side can be increased, that
is, the casing cooling effect on the front-stage side can be
enhanced.
[0044] The present embodiment has no large limitations, in the
axial direction, on the installation of the cooling air header 21.
The impingement cooling plate 22 is attached to the inside of the
cooling air header cover 23. Therefore, the cooling air header
cover 23 can be attached to and detached from the turbine casing 19
integrally with the impingement cooling plate 22. With the
configuration as above, it becomes easy to dispose the impingement
cooling plate 22 with respect to the cooling air header 21. Thus,
it is possible to keep small the front-stage side clearance that
would otherwise have to be increased during non-cooling of the
casing.
[0045] The configuration of the present embodiment is such that the
cooling passages 24 and the exhaust diffuser cooling passages 26
are connected to each other via the corresponding communication
holes 25. The cooling air 110 having cooled the casing is led via
the communication holes 25 to the exhaust diffuser cooling passages
26 to cool the exhaust diffuser 113. A conventional gas turbine is
such that the exhaust diffuser 113 and the casing 19 are cooled by
different air. However, when cooling air for the casing is reused
as cooling air for the exhaust diffuser in the present embodiment,
it is possible to suppress an additional increase in the amount of
cooling air resulting from the application of casing cooling.
[0046] Further, since the cooling air 110 flows in the cooling
passages 24, the rear side of the casing is cooled by convection
cooling, which makes it possible to reduce the clearance on the
rear-stage side of the turbine.
[0047] Another embodiment of the present invention is next
described with reference to FIGS. 7 and 8. FIG. 7 is a conceptual
diagram showing a state where thermal deformation occurs in a
casing. FIG. 8 is a diagram of a gas turbine cooling system. As
shown in FIG. 7, a casing 19 includes an upper-half casing 19a and
a lower-half casing 19b separated from each other, which are joined
to each other via respective flanges 35 thereof. During the
start-up of a plant, if thermal expansion occurs in the casing 19
having the flanges 35 as described above, the upper-half casing 19a
and the lower-half casing 19b each have a larger amount of thermal
expansion on the top side than that on the flange side. More
specifically, the top side portion is thermally expanded large in a
horizontal direction, whereas the flange side portion is thermally
expanded small in a vertical direction. This is because the flanges
35 formed at the division surface of the casing 19 exist, so that
the flange side portion has larger thermal capacity than the top
side portion. Consequently, as shown by a solid line in FIG. 7,
non-uniform thermal expansion (deformation) occurs in the overall
casing, that is, the flange side portion is displaced large
leftward and rightward outwardly.
[0048] Therefore, the present embodiment is configured such that a
flow rate of cooling air supplied to the top side of the casing
which has relatively large thermal expansion is made greater than
that supplied to the flange side which has relatively small thermal
expansion. A description is given of the configuration of the
present embodiment that achieves control for uniform clearance with
reference to FIG. 8.
[0049] A plurality of impingement cooling plates 22 are installed
inside cooling air headers of the upper-half casing 19a and the
lower-half casing 19b along the circumferential direction of the
casing 19. FIG. 8 shows an example in which eight impingement
cooling plates 22 are installed. For convenience sake, impingement
cooling plates disposed on the top side (on the vertical side) of
the upper-half casing 19a and the lower-half casing 19b are
referred to as the top side impingement cooling plates 22a. In
addition, impingement cooling plates disposed on the flange 35 side
are referred to as the flange side impingement cooling plates 22b.
A plurality of cooling air supply systems 38 are connected via
cooling air header covers 23 (not shown for convenience sake in
FIG. 8) to spaces each defined by each impingement cooling plate 22
(the spaces each defined by the impingement cooling plate 22 and
the cooling air header cover 23 shown in FIG. 6). The cooling air
supply system 38 supplies a cooling air (a cooling medium) for
impingement cooling. The cooling air supply system 38 includes a
common system 38a and a plurality of systems 38b, 38c bifurcated
from the common system 38a. The system 38b supplies cooling air to
a space defined by the top side impingement cooling plate 22a. The
system 38c supplies cooling air to a space defined by the flange
side impingement cooling plate 22b. An orifice 30 is installed in
the system 38c, of the bifurcate systems 38b, 38c, which is
connected to the space defined by the flange side impingement
cooling plate 22b. The orifice 30 serves as a flow control device
which regulates the flow rate of cooling air.
[0050] With the present embodiment described above, the flow rate
of cooling air flowing from the common system 38a to the system 38c
toward the flange side impingement cooling plate 22b, is regulated
by the orifice 30 so as not to exceed a predetermined flow rate. As
a result, the circumferential distributions in thermal expansion on
the top side and flange side of the casing are made uniform. This
makes it possible to uniformly reduce the clearances at the tips of
the turbine blades on the front-stage side of the gas turbine.
[0051] Another embodiment of the present invention is described
with reference to FIGS. 9 and 10. FIG. 9 is a conceptual view
showing a state where the center of a casing is not coincident with
the center of a turbine shaft. FIG. 10 is a diagram of a gas
turbine cooling system according to the present embodiment. As
shown in FIG. 9, the center of a casing 19 is not completely
coincident with the center of a turbine shaft 105 due to
manufacturing tolerance and a temporal change of the casing.
Therefore, the size of a clearance between a turbine rotor blade
and a shroud has circumferential deviation. If the casing is to
uniformly be cooled over the whole circumference thereof, the
thermal expansion of the casing will be reduced uniformly in the
whole circumference thereof. Thus, the non-uniformity of the
clearance cannot be eliminated.
[0052] The present embodiment is adapted to eliminate the
non-uniformity of the clearance mentioned above by installing a
device for regulating a circumferential cooling amount for the
casing and controlling radial and circumferential deformations of
the casing.
[0053] As shown in FIG. 10, impingement cooling plates 22 installed
in the casing 19 are sectioned in a circumferential direction. A
plurality of cooling air supply systems 38 are connected via
cooling air header covers 23 (not shown for convenience sake in
FIG. 10) to spaces each defined by each impingement cooling plate
22. The cooling air supply system 38 includes a common system 38a
and a plurality of systems 38d branched from the common system 38a.
The each system 38d supplies cooling air to each space defined by
each impingement cooling plate 22. An orifice 30 as a flow control
device for regulating a flow rate of cooling air is installed in
the each system 38d. A description is below given of an
orifice-diameter setting method.
[0054] After a gas turbine is assembled, it is confirmed that
setting clearances fall within tolerance, wherein a clearance
between the tip of a turbine rotor blade (a rotating body) and a
shroud (a stationary body) in a stationary state is
circumferentially measured at plural points. A deviation 8 between
the rotor center and the casing center is obtained from this
clearance measurement record. How much the clearance is to be
reduced in which direction during the operation of the gas turbine
is estimated to eliminate the non-uniformity of the clearance.
Thus, a clearance reduction amount to be targeted is
determined.
[0055] A relationship between the size of an orifice diameter and a
casing deformation amount at each position is previously evaluated
based on analysis using a finite element method and/or clearance
measurement results obtained by a real machine test. If a casing
deformation amount encountered when each orifice diameter is
independently changed is found, a casing deformation amount
encountered when a plurality of orifice diameters are
simultaneously changed can be estimated by synthesizing the
deformation amounts.
[0056] A target clearance reduction amount is determined based on
the clearance measurement record. An orifice diameter and
arrangement appropriate for achievement of the target clearance
reduction amount are determined based on the relationship between
the orifice diameter and the casing deformation amount. When the
gas turbine is assembled, several different types of orifices are
previously prepared. After clearance measurement, the orifice
diameter is determined and an original orifice is replaced with an
appropriate orifice in a short time.
[0057] Further, clearances are measured in the stationary state of
the turbine every disassembly and reassembly for periodic
inspections. Thus, also the temporal deformation of the casing can
be coped with when the orifice diameter is set again on the basis
of the clearance measurement record.
* * * * *