Rotary kinetic fluid motors or pumps – With passage in blade – vane – shaft or rotary distributor...
Reexamination Certificate
2002-10-16
2004-07-06
Look, Edward K. (Department: 3745)
Rotary kinetic fluid motors or pumps
With passage in blade, vane, shaft or rotary distributor...
C416S09700R
Reexamination Certificate
active
06758651
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to gas turbines where cooling of the stationary blades inside the turbine can be effectively performed when combustion gas from the combustor is introduced into the combustion gas flow path inside the turbine.
2. Description of the Related Art
A gas turbine is constructed by providing a compressor, a combustor and a turbine (not shown in the figure). According to this gas turbine, compressed air compressed by the compressor is supplied to the combustor, and combined with separately supplied fuel and combusted. The combustion gas generated by this combustion is then supplied to the turbine, and produces a rotational driving force on the turbine.
FIG. 4
shows an example of the inner structure of the turbine. As shown in this figure, inside the turbine, multiple moving blades
1
annularly arranged on the rotor side (not shown in the figure), and multiple stationary blades
2
provided on the stationary side around the rotor periphery, are arranged alternately in the rotation axis direction of the rotor (left to right in the figure), and a combustion gas flow path
3
is formed passing through these. Accordingly, combustion gas which has been introduced into the combustion gas flow path
3
from the combustor turns the moving blades
1
, applying a rotation force to the rotor. This rotation force turns a generator (not shown in the figure) connected to the rotor, to thus generate electricity.
The moving blades
1
and the stationary blades
2
are arranged alternately in:the rotation axis direction to form a multi-stage structure. Incidentally,
FIG. 4
only illustrates the part up to the first and second stages, counting from upstream where the combustion gas flows in, but in reality the multi-stage structure continues further to the third stage, fourth stage and so on. Also, reference symbol
4
shown in this figure denotes a tail pipe of the combustor, which is connected to the upstream portion of the first stage.
In this turbine, in order to cool the components of the second stage blades and the like, which are heated due to introducing the combustion gas to the inside, the components of the second stage blades and the like must be cooled, and for example a structure is generally adopted which bleeds and extracts part of the compressed air compressed by the compressor and uses this for cooling the parts of moving blade
1
and stationary blade
2
and the like.
As an example of this type of cooling structure, a cooling structure for the outside shroud
2
a
of the stationary blades
2
is shown in FIG.
5
. This figure is an enlarged crosssection of the part corresponding to part A of FIG.
4
.
As shown in this figure, in the outside shroud
2
a
, a plurality of cooling air flow paths
2
a
1
is piercingly provided around the peripheral direction along the upstream edge thereof, enabling film cooling where the inner surface
2
a
2
of outside shroud
2
a
is covered by cooling air c.
All of the cooling air flow paths
2
a
1
are arranged so as to flow the cooling air c from upstream to downstream (that is, from left to right in the figure) matching the flow direction of the combustion gas. In this manner, the cooling air c which is discharged from the upstream side edge, covers the inner surface
2
a
2
, and hence the heat from the combustion gas f towards the outside shroud
2
a
is decreased.
However, in the conventional gas turbine described above, there is a problem in that, at the upstream portion of inner surface
2
a
2
corresponding to the stationary blades
2
, film cooling cannot be effectively demonstrated, so that the wall temperature of the outside shroud
2
is locally increased.
That is to say, the combustion gas f which has reached the leading edge
2
b
1
of the blade porition
2
b
of the stationary blade
2
, separates into a flow towards the driving face side of the blade portion
2
b
, and a flow flowing along the suction surface, being the reverse face side. However, at the portion near the outside shroud
2
a
, as shown in
FIG. 5
, a U-shaped reverse flow is produced (since the shape of this flow resembles a horseshoe, this is called a horseshoe vortex). This horseshoe vortex is produced in the opposite direction to the flow of cooling air c discharged from the cooling air flow path
2
a
1
, and thus disturbs and obstructs the flow of cooling air c, reducing the cooling function. Therefore, the wall temperature of the outside shroud
2
is locally raised compared to at other places.
This increase in wall temperature attributable to the horseshoe vortex is not limited to the outside shroud
2
a
, and there is the likelihood of a similar occurrence also at the inside shroud
2
c
shown in FIG.
4
.
SUMMARY OF THE INVENTION
The present invention takes the above situation into consideration, with the object of providing a gas turbine where cooling failure attributable to the occurrence of a horseshoe vortex produced in the vicinity of the stationary blades of the turbine, can be prevented.
The present invention adopts the following means for solving the above problem.
That is to say, the present invention provides a gas turbine comprising moving blades provided on a rotor side which rotate together with the rotor, and stationary blades provided on a stationary side which cover the periphery of the moving blades and form a combustion gas flow path in the interior, and which are arranged alternately with the moving blades in the rotation axis direction of the rotor, and where the stationary blades have a blade portion arranged inside the combustion gas flow path, an outside shroud provided on an outer peripheral end side of the blade portion, and an inside shroud provided on an inner peripheral end side of the blade portion, wherein in one or both of the outside shroud and the inside shroud, corresponding to a leading edge of the blade portion, there is provided a first cooling air flow path which blows out cooling air into the combustion gas flow path, from downstream to upstream in the flow direction of the combustion gas.
According to the gas turbine, the horseshoe vortex which is generated corresponding to the leading edge of the blade portion, flows in the opposite direction to the direction of flow of the combustion gas, but because cooling air which is discharged from the first cooling air flow path also flows from downstream to upstream in the flow direction of the combustion gas, there is no direct confrontation of the flow direction of the cooling air with the horseshoe vortex as in conventional method. As a result, the supply of cooling air to the combustion gas flow path is easier than for the conventional method.
In the gas turbine, it is preferable that in one or both of the outside shroud and the inside shroud, there is provided a second cooling air flow path which blows out cooling air into the combustion gas flow path, at a position between a connection point with the leading edge and the first cooling air flow path.
According to the gas turbine, by means of the cooling air discharged from the second cooling air flow path, the portion between the connection point with the leading edge, and the first cooling air path can be cooled by convection cooling.
In the gas turbines, it is preferable that the first cooling air flow path is provided in the outside shroud of a first stage stationary blade, being the stationary blade arranged in the most upstream position in the flow direction of the combustion gas.
According to the gas turbine, since cooling failures attributable to the occurrence of a horseshoe vortex, is particularly likely to be a problem at the outside shroud of the first stage stationary blades, by applying the present invention to this portion, the effect of the present invention can be particularly effectively demonstrated.
REFERENCES:
patent: 2685429 (1954-01-01), Auyer
patent: 4732531 (1988-03-01), Minoda et al.
patent: 5584651 (1996-12-01), Pietraszkiewicz et al.
patent: 6139257 (2000-10-01), Proctor et al.
patent: 7-305603 (1995-11-01), Non
Shiozaki Shigehiro
Soechting Friedrich
Tomita Yasuoki
Look Edward K.
McAleenan J. M.
Mitsubishi Heavy Industries Ltd.
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