Note: Descriptions are shown in the official language in which they were submitted.
CA 02287577 1999-10-22
. GAS TURBINE COOLED MOVING BLADE
BACKGROUND OF THE INVENTION:
Field of the Invention:
The present invention relates generally to a cooled
moving blade of gas turbine and more particularly to a cooled
moving blade made in a structure to improve flow of cooling air
to enhance cooling efficiency as well as to improve flow of
sealing air to enhance sealing performance and cooling
performance.
Description of the Prior Art:
Fig. 3 is a cross sectional view of a representative
first stage moving blade of gas turbine in the prior art. In
Fig. 3, numeral 20 designates a moving blade, numeral 21
designates a blade root portion and numeral 22 designates a
platform. In the blade root portion 21, there are provided
cooling passages 23, 24, 25, 26 independently of each other.
The cooling passage 23 is a passage provided on a leading edge
side of the blade to communicate with a cooling passage 23a
provided in the blade, and while cooling air 40 entering the
cooling passage 23 from rotor side flows through the cooling
passage 23a, it cools a leading edge portion of the blade and
.",
at the same time flows out of cooling holes 29 to effect a shower
head film cooling on and around the leading edge portion.
Cooling air 41 entering the cooling passage 24 passes through
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a cooling passage 24a in the blade to turn at a tip portion of
the blade to flow through a cooling passage 24b and then turns
again at a base portion of the blade to flow through a cooling
passage 24c and flows out of the blade tip portion. At this
time, the cooling air 41 flows out of a blade surface through
cooling holes to effect a film cooling of the blade surface,
as described later with respect to Fig. 4.
Cooling air 42 entering the cooling passage 25 and
that 43 entering the cooling passage 26 join together to flow
through a cooling passage 25a and turn at the blade tip portion
to flow through a cooling passage 25b and then turn again at
the blade base portion to flow into a cooling passage 25c. While
the cooling air 42, 43 so flows through the cooling passage 25c,
a part thereof flows out of the blade surface through cooling
holes to effect a film cooling, as described later in Fig. 4,
and remainder thereof flows out of a trailing edge 27 of the
blade through between cooling fins 28 to effect a pin fin
cooling.
Fig. 4 is a cross sectional view taken on line B-B
of Fig. 3. As shown there, a portion of the cooling air 40 in
the cooling passage 23a on the blade leading edge side flows
out of the blade through cooling holes 29 to effect a shower
head film cooling for cooling of the blade leading edge portion.
Also, a portion of the cooling air 41 flowing through the
cooling passage 24c flows out obliquely through cooling holes
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30 to effect.a film cooling of the blade surface. Likewise,
a portion of the cooling air 42, 43 flowing through the cooling
passage 25c flows out of the blade surface obliquely through
cooling holes 31 to effect a film cooling of the blade trailing
edge portion. It is to be noted that although the cooling holes
29, 30, 31 only are illustrated in the figure, there are
provided actually a multiplicity of cooling holes in the blade
other than those mentioned here.
Fig. 5 is an explanatory view showing flow of sealing
air in a gas turbine in the prior art. In Fig. 5, a stationary
blade 50 is arranged in a rear stage of the moving blade 20.
Numeral 51 designates an inner shroud and numeral 52 designates
an outer shroud. Numeral 53 designates a cavity, which is
formed on an inner side of an end portion of the inner shroud
51 and numeral 54 designates a seal holding ring, which holds
a labyrinth seal 58. The labyrinth seal 58 together with a rotor
disc 80 forms a seal portion. Numeral 55 designates a hole,
which is bored in the seal holding ring 54 and through which
sealing air flows out, as described later. Numeral 56
designates a space, which is formed by and between the mutually
adjacent moving blade 20 and stationary blade 50, numeral 57
designates a honeycomb seal, which is provided at a front end
portion of the inner shroud 51 of the stationary blade 50 and
numeral 59 designates a space, which is formed by and between
the stationary blade 50 and a rear stage moving blade adjacent
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thereto. -
In the seal structure mentioned above, sealing air
70 entering a sealing air tube 71 provided in the stationary
blade 50 flows into the cavity 53 formed on the inner side of
the inner shroud 51 to elevate pressure therein higher than that
in a combustion gas path outside thereof and flows out through
the hole 55 of the seal holding ring 54 to flow into the space
56 and then passes through a gap of a seal portion formed by
a rear end portion 22b of the platform 22 of the moving blade
20 and the honeycomb seal 57 of the front end portion 57a of
the inner shroud 51 of the stationary blade 50 to flow out into
the combustion gas path as sealing air 70a. On the other hand,
a portion of the sealing air which has flown out of the hole
55 of the seal holding ring 54 passes through a gap of a seal
portion formed by the labyrinth seal 58 and the rotor disc 80
to flow into the space 59 and then passes through a seal portion
formed by a rear end portion 57b of the inner shroud 51 and a
platform front end portion of the rear stage moving blade to
flow out into the combustion gas path as sealing air 70b, like
in the case of the front stage.
The same seal structure is applied between the moving
blade 20 and a front stage stationary blade thereof, that is,
sealing air passes through a seal portion formed by a front end
portion 22a of the platform 22 of the moving blade 20 and an
inner shroud rear end portion of the front stage stationary
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blade and flows out into the combustion gas path as sealing air
70c. Using such seal structures as mentioned above, the spaces
on the inner side of the inner shroud 51 of the stationary blade
50 and the platform 22 of the moving blade 20 are held in a higher
pressure than in the combustion gas path so that a high
temperature combustion gas may be prevented from flowing into
these spaces on the inner side.
As mentioned above, in the f first stage moving blade
of gas turbine, cooling air is led into the blade for cooling
thereof and while the cooling air flows through cooling
passages in the blade, it flows out of the blade surface through
cooling holes to effect a shower head film cooling of the blade
leading edge portion and a film cooling of the blade ventral
and dorsal side portions. In the recent gas turbines wherein
combustion gas temperature is being elevated higher, combustor
outlet temperature of approximately 1150°C has been realized
and moreover a plant comprising that of as high as 1300°C or
more is being developed. As the first stage moving blade is
a portion that is highly exposed to the high temperature
combustion gas, cooling of the blade needs to be done most
efficiently and further improvement of the cooling structure
is desired.
Also, as to the sealing by air accompanying with the
higher temperature of the combustion gas, it is desired for
further enhancement of the sealing performance that a
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sufficient sealing pressure is ensured so that the combustion
gas may not come into the inner side of the platform and of the
inner shroud as well as a more efficient use of the sealing air
is attained.
SUMMARY OF THE INVENTION:
It is an object of the present invention, therefore,
to provide a gas turbine cooled moving blade, especially a first
stage moving blade, which comprises improved cooling air
passages in which cooling air flows efficiently as well as
comprises a seal structure in which sealing air flows also
efficiently, so that both of the cooling performance and the
sealing performance may be enhanced.
In order to achieve said object, the present
invention provides the following means of (1) and (2).
(1) A gas turbine cooled moving blade comprising a
cooling air passage provided in the moving blade for leading
therethrough cooling air supplied from below a platform for
cooling of the moving blade and a seal portion formed between
an inner shroud end each of a front stage stationary blade and
a rear stage stationary blade both adjacent to the moving blade
and each end of said platform and constructed such that sealing
air having passed through said seal portion flows out into a
combustion gas path, characterized in that said cooling air
passage comprises a first passage constructed such that the
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cooling air flows on a leading edge side of the moving blade
to flow out of a tip portion of the moving blade and a second
passage constructed such that a portion of the cooling air
supplied to a trailing edge side of the moving blade flows out
of a multiplicity of slots provided in a trailing edge of the
moving blade and a remainder thereof flows in a serpentine form
toward the leading edge side to flow out of the tip portion of
the moving blade.
(2) A gas turbine cooled moving blade as mentioned
in ( 1 ) above, characterized in that said seal portion comprises
a seal structure formed by a knife edge provided on each said
end of the platform and said inner shroud end each of the front
stage stationary blade and the rear stage stationary, blade and
the sealing air having passed through said seal portion flows
in a serpentine form to flow out into the combustion gas path
obliquely upwardly so as to go along a flow direction of
combustion gas therein.
In the invention of ( 1 ) above, the cooling air passing
through the first passage cools the blade leading edge portion
which is exposed to the combustion gas of the highest
temperature to be under the severe thermal influence and then
flows out of the blade tip portion as it is, thereby the blade
leading edge portion can be cooled effectively by the cold
cooling air. In the second passage of the cooling air, the blade
trailing edge portion is cooled first by the cold air, thereby
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the blade hub. portion which receives the thermal influence to
deteriorate the fatigue strength is cooled effectively so as
to prevent deterioration of the fatigue strength. Then, the
cooling air partially flows out of the slotted portion provided
in the blade trailing edge for cooling therearound and the
remaining portion flows through the passage formed in the
serpentine shape toward the blade leading edge side for cooling
of the blade main portions and flows out of the blade tip portion.
By the structure so constructed, the blade leading edge portion
is cooled effectively by the first passage of the cooling air
and the blade trailing edge portion and main portions are cooled
effectively by the second passage, respectively, and the
cooling efficiency can be enhanced. . ,
In the invention of (2) above, in addition to that
of ( 1 ) above, the seal portion is constructed by the knife edge
provided on each end of the platform and the inner shroud end
each of the front stage and rear stage stationary blades
adjacent to the moving blade and the sealing air carries out
a sealing with the effect of the knife edge, thereby the sealing
pressure which is elevated enough can be obtained at this
portion. Further, the sealing air after having passed through
the seal portion flows in the serpentine form with less quantity
of air leakage and with increased resistance of the flow passage
by the serpentine shape, thereby the sealing performance can
be enhanced. Also, the sealing air flows out into the
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combustion ga.s path obliquely upwardly so as to go along the
flow direction of combustion gas therein, thereby the sealing
air flows out to strike the end portion of the platform of the
moving blade and the inner shroud end of the stationary blade
and the effect to cool these portions can be obtained as well.
BRIEF DESCRIPTION OF THE DRAWINGS:
Fig. 1 is a cross sectional view of a gas turbine
cooled moving blade of an embodiment according to the present
invention.
Fig. 2 is a cross sectional view taken on line A-A
of Fig. 1.
Fig. 3 is a cross sectional view of a gas turbine
cooled moving blade in the prior art.
Fig. 4 is a cross sectional view taken on line B-B
of Fig. 3.
Fig. 5 is an explanatory view showing flow of sealing
air in a gas turbine in the prior art.
DESCRIPTION OF THE PREFERRED EMBODIMENTS:
Herebelow, embodiments according to the present
invention will be described concretely with reference to
figures. Fig. 1 is a cross sectional view of a gas turbine
cooled moving blade, especially as an example of a first stage
moving blade, of an embodiment according to the present
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invention and Fig. 2 is a cross sectional view taken on line
A-A of Fig. 1.
In both of Figs. 1 and 2, numeral 1 designates a moving
blade and numeral 2 designates a platform thereof. Numerals
3, 4, 5 designate cooling passages, respectively, which are
provided in a blade root portion 16 and into which cooling air
supplied from rotor side is led. The cooling passage 3
communicates with a cooling passage 3a provided in the blade
and the cooling passages 4, 5 join together in the blade root
portion 16 to form a cooling passage 6. That is, as compared
with the cooling passages 23 to 26 in the prior art shown in
Fig. 3, the cooling passages 3 to 5 shown in Fig. 1 are made
in a less number of pieces and also the flow passage of the
cooling passages 4, 5 is throttled by a rib so as to adjust flow
rate of air entering there.
The cooling passage 6 is provided on a trailing edge
side of the blade to communicate sequentially with cooling
passages 6a, 6b, 6c, 6d, 6e provided in the blade so as to form
a serpentine cooling passage. Numeral 7 designates a trailing
edge of the blade and a multiplicity of slots 15 are provided
there substantially along a turbine axial direction. Numeral
8 designates a multiplicity of oblique turbulators provided on
inner walls of the cooling passages 6a to 6e. Numerals 9a, 9b
designate a multiplicity of turbulators provided on an inner
wall of the cooling passage 3a on a leading edge side of the
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blade.
Numeral 10 designates an upper surface sloping
portion of a front end portion 2a of the platform 2, which slopes
upward so as to go along a gas flow direction in a gas path and
forms a passage 17 between itself and an inner shroud rear end
portion of a front stage stationary blade adjacent thereto.
Numeral 11 designates a lower surface sloping portion of a rear
end portion 2c of the platform 2, which slopes upward so as to
go along the gas flow direction in the gas path and forms a
passage 18 between itself and an inner shroud front end portion
of a rear stage stationary blade adjacent thereto. Numerals
12a, 12b, 12c designate seal pins, respectively, which are
provided for sealing a portion between the platform 2 of the
moving blade 1 and a platform of a moving blade provided
adjacently to the moving blade 1 in a turbine circumferential
direction.
Numerals 13, 14 designate knife edges, respectively.
The knife edge 13 is provided on a front end lower portion 2b
of the platform 2 adjacently to a seal portion of the front stage
stationary blade so as to seal an inner side of that seal portion.
The knife edge 14 is likewise provided on a rear end lower
portion 2d of the platform 2 adjacently to a seal portion of
the rear stage stationary blade so as to seal an inner side of
that seal portion. Numerals 100, 101, 102 designate cooling
air, respectively, which is supplied from the rotor side to
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enter the cooling passages 3, 4, 5, respectively. The moving
blade 1 is applied to its surface by a TBC (thermal barrier
coating ) , such as a ceramics coating, so as to stand a thermal
influence of a high temperature combustion gas in the gas path.
In the cooled moving blade constructed as mentioned
above, the cooling air 100 entering the cooling passage 3 flows
into the cooling passage 3a linearly and, while becoming
turbulent by the turbulators 9a, 9b so as to enhance a heat
transfer rate, cools a leading edge portion of the blade, which
is exposed to a high temperature to be in the thermally severest
state, and then flows out of a tip portion of the blade to flow
into the combustion gas path.
The cooling air 101, 102 entering the cooling
passages 4, 5, respectively, join together in the cooling
passage 6 to flow toward the blade tip portion through the
cooling passage 6a, and, while becoming turbulent by the
turbulators 8 so as to enhance the heat transfer rate, cools
a trailing edge portion of the blade. At the same time, a
portion of the cooling air flowing through the cooling passage
6a flows out of the multiplicity of slots 15 provided in the
trailing edge 7 substantially in the turbine axial direction
for cooling therearound.
Remaining portion of the cooling air flowing through
the cooling passage 6a turns at the blade tip portion to enter
the cooling passage 6b to flow toward the platform 2 for cooling
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that portion of the blade and then turns below the platform 2
to enter the cooling passage 6c to flow therethrough and further
through the cooling passages 6d and 6e, turning at the blade
tip portion and below the platform 2, respectively, and flows
outside of the blade tip portion. While the cooling air 101,
102 so flows through the serpentine cooling passage comprising
the cooling passages 6a, 6b, 6c, 6d, 6e, it becomes turbulent
by the turbulators 8 provided obliquely for enhancement of the
cooling effect and cools main portions of the blade and then
flows out of the blade tip portion.
As mentioned above, as the cooling air 101, 102 is
led into the trailing edge side of the moving blade 1 so that
a cold air may flow into a hub portion of the blade trailing
edge portion or especially so that a fatigue strength against
heat of the trailing edge hub portion may be enhanced and
further as the cold air enters the cooling passage 6a first,
slotted portion of the blade trailing edge 7 is cooled
efficiently, the oblique turbulators 8 in the cooling passage
6a can be made less in the number of pieces than in other cooling
passages and the cooling passage 6a itself can be made larger
than other cooling passages . Also, in the moving blade 1 having
the cooling system so constructed, the cooling holes 29, 30,
31 (Fig. 4 ) in the prior art become unnecessary and the shower
head film cooling there becomes also unnecessary.
Sealing air 103 on a front end side of the platform
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2, which is supplied through the adjacent front stage
stationary blade as described with respect to Fig. 5, flows
through between the knife edge 13 and the seal portion of the
front stage stationary blade to flow in a serpentine form and
then to flow out obliquely upwardly through the passage 17
formed by the upper surface sloping portion 10, hence by the
flow of the serpentine form, sealing performance is enhanced
resulting in reducing leaking loss of air. Also, as the sealing
air having so flown out flows obliquely upwardly along the
direction of the combustion gas flow, it strikes the front end
portion 2a of the platform 2 with the effect to cool that portion
as well.
Further, sealing air 104 on a rear end side of the
platform 2, which is likewise supplied through the adjacent
rear stage stationary blade, flows through between the knife
edge 14 and the seal portion of the rear stage stationary blade
to flow in the serpentine form and then to flow out through the
passage 18 obliquely upwardly along the direction of the
combustion gas flow, thus sealing performance there is likewise
enhanced and by the sealing air so flowing outside, there is
obtained the effect to cool the inner shroud front end portion
of the adjacent rear stage stationary blade.
According to the present embodiment described as
above, the cooling air 100 flows linearly through the cooling
passage 3a on the blade leading edge side, becoming turbulent
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by the turbulators 9a, 9b, to cool the blade leading edge
portion only. The cooling air 101, 102 enters on the blade
trailing edge side to flow through the serpentine cooling
passage comprising the cooling passages 6a, 6b, 6c, 6d, 6e,
becoming turbulent by the turbulators 8 provided obliquely
therein, to cool those portions of the blade, hence the entire
blade can be cooled with the enhanced cooling effect.
Moreover, the sealing air 103, 104 flows through the
serpentine routes formed by the knife edges 13, 14 and the upper
surface and lower surface sloping portions 10, 11 to flow out
obliquely upwardly through the passages 17, 18, thereby the
sealing performance is enhanced and also the front end portion
2a of the platform 2 of the moving blade 1 and the inner shroud
front end portion of the adjacent rear stage stationary blade
can be cooled as well.
It is understood that the invention is not limited
to the particular construction and arrangement herein
illustrated and described but embraces such modified forms
thereof as come within the scope of the appended claims.
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