Rotary kinetic fluid motors or pumps – Working fluid passage or distributing means associated with... – Plural distributing means immediately upstream of runner
Reexamination Certificate
2000-10-12
2002-07-16
Verdier, Christopher (Department: 3745)
Rotary kinetic fluid motors or pumps
Working fluid passage or distributing means associated with...
Plural distributing means immediately upstream of runner
C415S115000, C415S208200, C415S210100, C416S09600A, C416S09700R, C416S19300A
Reexamination Certificate
active
06419447
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to a turbine blade of a gas turbine or the like and a gas turbine equipment using this turbine blade.
2. Description of the Prior Art
FIG. 5
is a schematic explanatory view of a structure of a turbine portion and a cooling air system for cooling this turbine portion in a gas turbine equipment in the prior art.
The turbine portion comprises a rotational portion of a rotor
1
and a turbine moving blade
2
and a stationary portion
5
of a casing
3
, a turbine stationary blade
4
, various supporting members and the like.
In the turbine portion, a high temperature high pressure combustion gas supplied from a combustor
6
is converted into a high velocity flow by the turbine stationary blade
4
to rotate the turbine moving blade
2
for generation of power.
Construction members of the rotational portion and the stationary portion which are adjacent to the combustion gas need to be cooled so that their temperature due to heat input from the combustion gas may not exceed their respective allowable temperature and, for cooling of the rotational portion having the rotor
1
and the turbine moving blade
2
, it is usual that cooling medium
7
is supplied as shown by arrows in FIG.
5
.
The cooling medium
7
is often a bleed air or discharge air taken from a compressor (not shown) or sometimes the bleed air or discharge air once supplied into a cooler (not shown) and cooled to an appropriate temperature.
Further, as the cooling medium to cool the mentioned portions, there is recently a case where steam from an outside system is applied in place of the bleed air or discharge air from the compressor, but herebelow description will be made based on the cooling air system which is generally employed as a typical example.
While the cooling medium
7
flowing in the rotational portion takes a route to flow through an interior of the rotor
1
to enter an interior of the turbine moving blade
2
for cooling thereof and then to join into a combustion gas path, in the case of using steam as the cooling medium as mentioned above, the cooling medium which has been heat-exchanged by cooling the turbine moving blade
2
and the like is recovered so that thermal energy thereof may be made use of in an outside system and thermal efficiency of the plant may be enhanced.
In the gas turbine equipment having the mentioned basic structure, description will be made concretely on the prior art turbine portion thereof with reference to
FIGS. 6
to
10
.
FIG. 6
is a longitudinal cross sectional view showing a main structure of a prior art turbine moving blade,
FIG. 7
is a perspective view showing a main structure of a prior art turbine stationary blade,
FIG. 8
is an enlarged view of a part of the turbine stationary blade of
FIG. 7
,
FIG. 9
is a qualitative explanatory view showing a metal temperature behavior due to thickness difference between thickness of a turbine moving blade trailing edge portion and that of a platform in the prior art, and
FIG. 10
is likewise a qualitative explanatory view showing a metal temperature behavior due to thickness difference between thickness of a turbine stationary blade trailing edge portion and that of a shroud in the prior art.
In a leading edge portion of the turbine moving blade
2
which is exposed to an especially high temperature combustion gas, in order to stand a high thermal load, it is usual to provide a cooling passage
8
through which the cooling medium
7
is supplied for effecting a convection cooling in the turbine moving blade
2
.
Cooling passage in the moving blade is often constructed to repeat several turnings so as to form a serpentine passage on design demand, wherein the passage turns at a turning portion
11
provided in the vicinity of a tip portion
9
of the turbine moving blade
2
and a joint portion
10
of the turbine moving blade
2
.
Thus, the cooling medium
7
flows through the cooling passages to cool the interior of the turbine moving blade
2
. However, in case the turbine moving blade
2
is one which receives higher thermal load, there is provided a film cooling hole
12
in a blade surface of the turbine moving blade
2
and a portion of the cooling medium
7
is blown therethrough onto the blade surface on the combustion gas path side so that the blade surface may be covered by a low temperature air curtain and thereby a film cooling for reducing the thermal load from the blade surface as well can be effected.
On the other hand, a trailing edge portion
14
of the turbine moving blade
2
is usually designed to be relatively thin in order to reduce an aerodynamic loss of the combustion gas and, for this purpose, if the turbine moving blade
2
is to be cooled, a pin fin cooling or a slot cooling by way of many slots is employed for cooling the interior of the blade, or the film cooling by way of blowing air from a ventral side surface of the blade through the film cooling hole is effected.
In case of the turbine stationary blade
16
, in order to form a gas flow path, structure of the blade is made such that an inner end of a blade profile portion
17
is inserted into an inner shroud
18
and an outer end of the blade profile portion
17
is inserted into an outer shroud
19
, and while this set of one inner shroud
18
and one outer shroud
19
is usually provided for each of the turbine stationary blades
16
, there is also such a case where the set of one inner shroud
18
and one outer shroud
19
is provided so as to cover a plurality of the turbine stationary blades
16
.
The turbine stationary blade
16
is usually formed by precision casting and is then worked by machining, wherein the inner shroud
18
, the outer shroud
19
and the blade profile portion
17
are generally formed integrally by casting.
As mentioned above, the platform
15
supporting the turbine moving blade
2
forms a part of the gas flow path in an axial flow turbine and is made relatively thicker as compared with the trailing edge portion
14
of the blade so as to stand centrifugal force or the like.
For this reason, in operation of the gas turbine including start and stop, load change or the like, there may arise an excessively large temperature difference between the platform
15
and the blade trailing edge portion
14
, by which thermal stress is liable to occur at a transition time or in a steady operation time so that there is a risk to cause cracks and if the cracks occur, there is a problem to damage a reliability of the turbine moving blade.
Also, in the turbine stationary blade
16
, in order to reduce an aerodynamic loss, a trailing edge portion
20
of the blade is designed as thin as possible and, on the other hand, the inner shroud
18
and the outer shroud
19
are usually designed relatively thicker for holding the strength. Thus, like the turbine moving blade
2
, there is a problem that cracks are considered to occur by the thermal stress following a start and stop of the gas turbine or the like, which results in damaging the reliability.
The mentioned relation between the moving blade trailing edge portion and the platform is shown in
FIG. 9
qualitatively as a metal temperature behavior which is caused by a thickness difference between thickness of the moving blade trailing edge portion and that of the platform. Likewise, the mentioned relation between the stationary blade trailing edge portion and the shroud is shown in
FIG. 10
qualitatively as a metal temperature behavior which is caused by a thickness difference between thickness of the stationary blade trailing edge portion and that of the shroud.
In
FIGS. 9 and 10
, the vertical axis means a gas turbine rotational speed and metal temperature and the horizontal axis means a lapse of time. When the gas turbine is stopped, gas turbine rotational speed C
1
, C
2
is reduced. In the area of C
1
and C
2
, the blade trailing edge portion which is of a smaller thermal capacity is cooled quicker and moving blade trailing edge portion metal temperature B
1
and stationary blade
Matsuura Masaaki
Suenaga Kiyoshi
Watanabe Koji
Mitsubishi Heavy Industries Ltd.
Wenderoth , Lind & Ponack, L.L.P.
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