Rotary kinetic fluid motors or pumps – With passage in blade – vane – shaft or rotary distributor...
Reexamination Certificate
2002-07-08
2004-08-31
Look, Edward K. (Department: 3745)
Rotary kinetic fluid motors or pumps
With passage in blade, vane, shaft or rotary distributor...
C415S175000
Reexamination Certificate
active
06783323
ABSTRACT:
FIELD OF THE INVENTION
The present invention relates to a cooling structure of a gas turbine stationary blade mainly used for a two or more-staged stationary blade.
BACKGROUND OF THE INVENTION
A stationary blade of a gas turbine used for generating electric power provides a flow passage for combustion gas, of which temperature reaches about 1300° C. Therefore, in order to prevent melt or damage of the stationary blade by combustion gas, various cooling structures are provided to the gas turbine stationary blade. As a technique concerning such a cooling structure, there is a widely-known technique in which a passage for cooling air is provided in the gas turbine stationary blade, cooling air is sent into this passage, thereby cooling the gas turbine stationary blade from inside (refer to Japanese Patent Application Laid-open No. 11-132005 for example).
The gas turbine stationary blade has a dividable structure capable of assembling and disassembling the gas turbine stationary blade in consideration of easy maintenance after installation thereof.
FIG. 14
is a perspective view showing a segment
1
that is a unit constituent element of a two-staged stationary blade of a gas turbine. Each unit constituting this segment
1
comprises a substantially parallelogram inner shroud
2
, one columnar stationary blade section
3
whose one end is fixed to the inner shroud
2
, and a substantially parallelogram outer shroud
4
installed substantially in parallel to the inner shroud
2
and fixed to the other end of the stationary blade section
3
. The segment
1
comprises a pair of the units welded and connected to each other side-by-side. The gas turbine stationary blade comprises a plurality of segments
1
connected to one other side-by-side through detachable connection members (not illustrated) such as bolts such that the gas turbine stationary blade is formed into an annular structure as a whole. The gas turbine stationary blade is fixed and installed in a gas turbine casing (not illustrated) with a cantilever structure by means of legs
5
provided on an outer peripheral side face of the outer shroud
4
.
A bolt joint section
7
of the segment
1
keeps a specific distance so as to absorb expansion of the gas turbine stationary blade when the gas turbine is driven. This distance is set such that the distance is made zero by expansion of the gas turbine stationary blade when the gas turbine is driven. However, due to tolerance during the actual producing procedure, a gap
7
a
ranging from about 0.5 mm to 1 mm is produced in the bolt joint section
7
.
FIG. 15
is an enlarged perspective view around the inner shroud
2
shown in FIG.
14
.
FIG. 16
is a plan sectional view of the inner shroud
2
shown in FIG.
15
.
FIG. 17
is a side sectional view of the inner shroud
2
taken along the line I—I in FIG.
16
.
FIG. 18
is a side sectional view of the inner shroud
2
taken along the line II—II in FIG.
16
. In
FIG. 15
to
FIG. 18
, the gas turbine stationary blade has a stationary blade section front edge passage
9
and a stationary blade section rear edge passage
10
isolated from each other by a rib
8
which are provided inside the stationary blade section
3
. The stationary blade section front edge passage
9
is in communication with an open chamber
11
provided in the inner shroud
2
. The stationary blade section rear edge passage
10
passes through the inner shroud
2
, and is in communication with a cavity
12
formed in a bottom face section of the inner shroud
2
. The open chamber
11
and the cavity
12
are isolated from each other by a bottom plate
13
installed on the bottom face section of the inner shroud
2
. A member
14
, shown in
FIG. 17
, in the stationary blade section
3
is an impingement tube
14
comprising a metal member inserted into the stationary blade section front edge passage
9
and the stationary blade section rear edge passage
10
so as to subject the stationary blade section
3
to impingement cooling.
In the inner shroud
2
, a front edge
15
is located in upstream portion in the flow passage for combustion gas
6
. A front edge flow passage
16
is provided along the front edge
15
. The front edge flow passage
16
and the open chamber
11
are in communication with each other through an intermediate flow passage
17
provided therebetween. A regulating plate
18
is laid on a floor section of the front edge flow passage
16
to narrow a cross sectional area of the flow passage. A plurality of turbulators
20
are provided on the regulating plate
18
and a ceiling section of the front edge flow passage
16
to agitate the cooling air
19
.
From an outlet orifice of the front edge flow passage
16
, a central flow passage
21
having a cross sectional area smaller than that of the front edge flow passage
16
is pulled out. The central flow passage
21
comes out from a rear edge
23
of the inner shroud
2
that is downstream of the flow passage of combustion gas along the welded joints
22
of the inner shrouds
2
. From a position near an inlet orifice of the front edge flow passage
16
also, a side edge flow passage
24
having a cross sectional area smaller than that of the front edge flow passage
16
is pulled out. The side edge flow passage
24
comes out from the rear edge
23
along a side edge
25
of the inner shroud
2
(refer to FIG.
16
and FIG.
18
). The cooling structure is provided for each pair of units constituting the segment
1
, and a pair of left and right cooling structures are provided to constitute a cooling structure of the inner shroud
2
.
At the time of actuation of the gas turbine, when the inner shroud
2
is to be cooled, cooling air
19
is sent to the impingement tube
14
in the stationary blade section
3
from the outer shroud
4
. The cooling air
19
subjects the stationary blade section
3
to impingement cooling, a portion of the cooling air
19
flows into the open chamber
11
in the inner shroud
2
through the front edge passage
9
of the stationary blade section, and a portion of the cooling air
19
penetrates the inner shroud
2
through the stationary blade section rear edge passage
10
and is supplied to the cavity
12
(refer to FIG.
17
). The cooling air
19
which has flowed into the open chamber
11
flows into the front edge flow passage
16
through the intermediate flow passage
17
to convection-cool the front edge of the inner shroud
2
. A portion of the air flows into the side edge flow passage
24
from an inlet orifice of the front edge flow passage
16
, convection-cools the side edge
25
of the inner shroud
2
, and is discharged from the rear edge
23
. Remaining cooling air
19
flows into the central flow passage
21
from the outlet orifice of the front edge flow passage
16
, convection-cools welded joints
22
of the inner shrouds, and is discharged from the rear edge
23
.
The regulating plate
18
is provided to prevent reduction in flow speed of the cooling air
19
which passes through the front edge flow passage
16
by narrowing the cross sectional area of the passage, and to enhance the cooling efficiency of the front edge
15
. The turbulator
20
agitates the cooling air
19
in the front edge flow passage
16
, and enhances cooling efficiency of the front edge
15
. The central flow passage
21
and the side edge flow passage
24
have cross sectional areas smaller than those of the front edge flow passage
16
. Therefore, flow speed of the cooling air
19
passing through the flow passages
21
and
24
is faster than that in the front edge flow passage
16
. Thus, the structure in which the flow passage is narrowed enhances the cooling efficiency near the welded joints
22
of the inner shrouds
2
and near the side edge
25
.
The cooling air
19
supplied to the cavity
12
is used as sealing air for sealing a gap (not illustrated) between the gas turbine stationary blade and a gas turbine rotor blade. A portion of the sealing air is blown out from a bottom surface section of the front edge
15
, to film-cool the inner shroud
2
Kuwabara Masamitsu
Shiozaki Shigehiro
Tomita Yasuoki
Edgar Richard A.
Look Edward K.
Mitsubishi Heavy Industries Ltd.
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