Rotary kinetic fluid motors or pumps – With passage in blade – vane – shaft or rotary distributor...
Reexamination Certificate
2002-07-25
2004-07-13
Look, Edward K. (Department: 3745)
Rotary kinetic fluid motors or pumps
With passage in blade, vane, shaft or rotary distributor...
C416S09600A
Reexamination Certificate
active
06761529
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to a cooling system of a stationary blade of a gas turbine, in particular, a cooling system of a stationary blade having superior cooling efficiency, and to a gas turbine.
2. Description of Related Art
A gas turbine used for a generator and the like is shown in FIG.
4
.
Compressor
1
, combustor
2
, and turbine
3
are shown in
FIG. 4
, and rotor
4
extends from compressor
1
to turbine
3
in the axial direction.
Inner housing
6
, and cylinders
7
and
8
provided at the compressor
1
side enclose the outside of compressor
1
. Furthermore, cylindrical shell
9
forming chamber
14
, outside shell
10
of turbine
3
, and inside shell
11
are provided in the gas turbine.
Inside of cylinder
8
which is provided in compressor
1
, stationary blades
12
are disposed in the circumferential direction at equal intervals. Moving blades
13
, which are disposed around rotor
4
at equal intervals, are disposed between stationary blades
12
.
Combustor
15
is disposed in chamber
14
which is enclosed by cylindrical shell
9
. Fuel supplied from fuel feeding pipe
35
is injected from fuel injection nozzle
34
into combustor
15
to burn.
A high temperature combustion gas generated in combustor
15
is introduced into turbine
3
while passing through duct
16
.
In turbine
3
, two-stage type stationary blades
17
, which are disposed in the circumferential direction at equal intervals on inside shell
11
, and moving blades
18
, which are disposed in the circumferential direction at equal intervals on rotor
4
, are alternately provided in the axial direction. The high temperature combustion gas is fed into turbine
3
and is discharged as an expanded gas, and further, the high temperature combustion gas rotates rotor
4
on which moving blades
18
are fixed.
Manifolds
21
and
22
are provided in compressor
1
and turbine
3
respectively. Manifolds
21
and
22
are connected with each other by air piping
32
, and cooling air is supplied from the compressor
1
side to the turbine
3
side via air piping
32
.
A portion of cooling air from compressor
1
is supplied from a rotor disc to moving blades
18
in order to cool moving blades
18
. As shown in
FIG. 4
, a portion of cooling air from manifold
21
of compressor
1
passes through air piping
32
and is introduced into manifold
22
of turbine
3
to cool stationary blades
17
, and simultaneously, the cooling air is supplied as sealing air.
Next, a structure of stationary blades
17
will be explained below.
In
FIG. 5
, inner shroud
26
and outer shroud
27
are provided at the inside and the outside of blade
25
respectively.
Inside of blade
25
, leading edge path
42
and trailing edge path
44
are formed by rib
40
. Cylindrical insert parts
46
and
47
, in which plural cooling air holes
70
,
71
,
72
, and
73
are formed at the peripheral surfaces and bottom surfaces, are inserted from the outer shroud
27
side into the leading edge path
42
and trailing edge path
44
.
Blade
25
is equipped with pin fin cooling part
29
comprising a flow path having plural pins
62
at the trailing edge side.
When cooling air is supplied from manifold
22
into insert parts
46
and
47
, the cooling air is ejected from cooling air holes
70
,
71
,
72
, and
73
, and hits the inner walls of leading edge path
42
and trailing edge path
44
to carry out so-called impingement cooling. Furthermore, the cooling air flows through pin fin cooling part
29
comprising flow paths formed between plural pins
62
at the trailing edge side of blade
25
to carry out pin fin cooling.
On inner shroud
26
, forward flange
81
and rearward flange
82
are formed at the leading edge side and the trailing edge side, and are connected to seal supporting part
66
, which supports seal
33
for sealing arm
48
of rotor
4
and seal supporting part
66
. Furthermore, cavity
45
is formed between seal supporting part
66
and inner shroud
26
. The cooling air ejected from cooling air holes
70
,
71
,
72
, and
73
of insert parts
46
and
47
is supplied into cavity
45
.
Flow path
85
is formed at the forward side of seal supporting part
66
. Air is injected from cavity
45
while passing through flow path
85
toward the front stage moving blade
18
and toward the rear stage moving blade while passing through spaces formed in seal
33
, and the inside is maintained at a pressure higher than that of a path of high temperature combustion gas in order to prevent high temperature combustion gas from penetrating to the inside.
As shown in
FIGS. 6 and 7
, leading edge flow path
88
equipped with plural needle fins
89
is formed at the leading edge side of inner shroud
26
. Leading edge flow path
88
is connected to cavity
45
via flow path
90
. Rails
96
are formed along the leading edge toward the trailing edge at both sides of inner shroud
26
. In each rail
96
, flow path
93
is formed in which one end of each rail
96
is connected to leading edge flow path
88
and the other end of each rail
96
opens at the trailing edge of inner shroud
26
.
On the bottom surface of inner shroud
26
, collision plates
84
having plural small holes
101
are provided at an interval from the bottom surface. By providing these collision plates
84
, chamber
78
is formed at the bottom surface side of inner shroud
26
.
Furthermore, at the trailing edge side of inner shroud
26
, plural flow paths
92
are formed so as to be connected to the trailing edge of inner shroud
26
and chamber
78
.
Cooling air flowing into cavity
45
is injected into leading edge flow path
88
of inner shroud
26
via flow path
90
, passes through the space between needle fins
89
to cool the leading edge side of inner shroud
26
, and subsequently passes through side flow path
93
to be ejected from the trailing edge of inner shroud
26
.
Moreover, cooling air flowing into cavity
45
flows into chamber
78
from small holes
101
and passes through flow path
92
to be ejected from the trailing edge of inner shroud
26
. When cooling air flows into chamber
78
from small holes
101
of collision plate
84
, cooling air hits the bottom surface of inner shroud
26
, carrying out impingement cooling. Due to impingement cooling, cooling air passes through plural flow paths
92
to cool the trailing edge side of inner shroud
26
.
As shown in
FIG. 8
, collision plates
102
having plural small holes
100
are provided at the upper surface of outer shroud
27
at an interval from the upper surface. By providing these collision plates
102
, chamber
104
(not shown) is formed at the upper surface side of outer shroud
27
.
Leading edge flow path
105
is formed in outer shroud
27
, and side flow path
106
, which opens at the trailing edge of outer shroud
27
, is formed at both sides thereof. Leading edge flow path
105
is connected to one chamber
104
.
Furthermore, at the trailing edge side of outer shroud
27
, plural flow paths
107
are formed so as to be connected to the trailing edge of outer shroud
27
and chamber
104
.
Cooling air flowing into manifold
22
flows into chamber
104
from small holes
100
of collision plate
102
and passes through trailing edge flow path
107
to be ejected from the trailing edge of outer shroud
27
. When cooling air flows into chamber
104
from small holes
100
of collision plate
102
, cooling air hits the upper surface of outer shroud
27
, carrying out impingement cooling.
Furthermore, cooling air flowing into chamber
104
flows into leading edge flow path
105
and passes through leading edge flow path
105
and side flow paths
106
to cool the leading edge and both sides of outer shroud
27
. Subsequently, cooling air is ejected from the trailing edge of outer shroud
27
.
As described above, in stationary blades of this type of gas turbine, the blade metal temperature is maintained at an allowable temperature or less using various cooling techniques, such as impingement cooling, and pin fin co
Hada Satoshi
Kuwabara Masamitsu
Masada Junichiro
Soechting Friedrich
Tomita Yasuoki
Edgar Richard
Look Edward K.
Mitshubishi Heavy Industries, Ltd.
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