Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.
CA 02506206 2002-01-11
28964-53G
BLADE STRUCTURE IN A GAS TURBINE
This application is a divisional application of
Canadian Patent Application No. 2,367,711 filed on January 11,
2002.
FIELD OF THE INVENTION
This invention relates to a blade structure in a
gas turbine. More particularly, this invention relates to a
blade structure of a gas turbine with improved turbine
efficiency by restricting pressure loss to a minimum level.
BACKGROUND OF THE INVENTION
A gas turbine will be explained with reference to
Fig. 16. In general, a gas turbine is equipped with a
plurality of stages of stationary blades 2 and 3 arrayed in
a circle on a casing (a blade circle or a vehicle chamber)
1, and a plurality of moving blades 5 arrayed in a circle on
a rotor (a hub of a base) 4. Fig. 16 shows the moving blade
5 at a certain stage, the stationary blade 2 at the same
stage (the inlet side of combustion gas 6) as this moving
blade 5, and the stationary blade 3 at the next stage (the
outlet side of the combustion gas 6) of this moving blade 5.
When pressure loss is large in the gas turbine,
turbine efficiency is lowered. Therefore, it is important
to improve the turbine efficiency by minimizing the pressure
loss.
However, as shown in Fig. 16, there is a case where
the moving blade 5 at a certain stage is what is called a
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free-standing moving blade that has a clearance 8 between
a tip 7 of this moving blade 5 and the casing 1. In the
case of this free-standing moving blade 5, there is the
following problem.
Namely, as shown in Fig. 17, a main flow (shown by
a solid-line arrow mark in Fig. 17) of combustion gas 6 flows
to the next-stage stationary blade 3 side by passing through
between the moving blade 5 and the moving blade 5. In the
mean time, in the clearance 8 between the chip 7 of the moving
l0 blade 5 and the casing 1, there is generated a leakage flow
9 (shown by a broken-line arrow mark in Fig. 17) that is
separate from the main flow of the combustion gas 6.
A mechanism of generating the leakage flow 9 is that
as the pressure at a belly surface 10 side of the moving
blade 5 is higher than the pressure at a rear surface 11
side of the moving blade 5, the leakage flow 9 is generated
from the belly surface 10 side to the rear surface 11 side
based on a difference between these pressures.
As shown in Fig. 17, the leakage flow 9 flows at an
incidence angle ic to the rear surface 13 side at a front
edge 12 of the tip of the stationary blade 3 at the next
stage. This leakage flow 9 becomes a flow perpendicular
to the main flow of the combustion gas 6 that flows to the
belly surface 14 side of the stationary blade 3.
Therefore, a vortex flow 15 (shown by a solid-line
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spiral arrowmark in Fig. 17) is generated at the belly surface
14 side of the front edge 12 of the tip of the stationary
blade 3. When this vortex flow 15 is generated, pressure
loss occurs. The main flow of the combustion gas 6 may
deviate from the belly surface 14 side of the stationary
blade 3. In Fig. 17, a ref erence symbol (3c denotes an entrance
metal angle at the tip portion of the stationary blade 3.
Similarly, a reference symbol Oc denotes an inlet
included angle at the tip portion of the stationary blade
3. Similarly, a reference number 22 denotes a camber line
for connecting between the front edge 12 of the tip portion
of the stationary blade 3 and a rear edge 23 of the tip
portion.
The incidence angle ic of the leakage flow 9 and the
pressure loss have a relative relationship as shown by a
solid-line curve in Fig. 18. The solid-line curve in Fig.
18 shows a case of the inlet included angle Oc at the tip
portion of the stationary blade 3 shown in Fig. 17.
In this case, the inlet included angle Oc at the
tip portion of the stationary blade 3 has been set such
that the pressure loss becomes minimum (refer to a point
P1 in Fig. 18). However, as described above, the leakage
flow 9 is generated, and the pressure loss also becomes large
when the incidence angle ic of this leakage flow 9 is large
(refer to a point P2 in Fig. 18) . When this pressure loss
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is large, the turbine efficiency is lowered by that amount.
Further, as shown in Fig. 16, seal-air 16 (shown by
a two-dot chained line arrow mark in Fig. 16) flows from
the rotor 4 side at the upstream of the moving blade 5 at
a certain stage. When this seal-air 16 is flowing, there
is the following problem.
Namely, the seal-air 16 simply flows out straight in
a direction of the height (a radial direction of the turbine)
of the moving blade 5 without being squeezed by a nozzle
or the like. On the other hand, the moving blade 5 is rotating
in a direction of an outline arrow mark together with the
rotor 4. Therefore, from the relative relationship between
the flow-out of the seal-air 16 and the rotation of the moving
blade 5, the seal-air 16 flows at the incidence angle is
to the rear-surface side 11 at the front edge 17 of the hub
portion of the moving blade 5, as shown in Fig. 17.
As explained above, when the incidence angle is of
the seal-air 1,6 becomes large at the front edge 17 of the
hub portion of the moving blade 5 as well, the pressure loss
becomes large and the turbine efficiency is lowered by that
amount as shown in Fig. 17 and Fig. 18, in a similar manner
to that at the front edge 12 of the tip portion of the
stationary blade 3.
This problem of the hub portion of the moving blade
5 also applies to a shrouded moving blade in addition to
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the above-described free-standing moving blade. In Fig.
17, a reference symbol Ps denotes an entrance metal angle
at the hub portion of the moving blade 5. Similarly, a
reference symbol As denotes an inlet included angle
at the hub portion of the moving blade S. Similarly, a
reference number 24 denotes a camber line for connecting
between the front edge 17 of the hub portion of the moving
blade 5 and a rear edge 25 of the hub portion.
Further, when the moving blade 5 at a certain stage
is a free-standing*moving blade, there is the following
problem.
Namely, as shown in Fig. 17, the leakage flow 9 is
generated from the belly surface 10 side of the moving blade
5 to the rear surface 11 side, at the clearance 8 between
the tip 7 of the free-standing moving blade 5 and the casing
1.
Then, as shown in Fig. 19B, a design Mach number
distribution shown by a solid-line curve becomes an actual
Mach number distribution as shown by a broken-line curve.
As a result, on the rear surface 11 of the tip portion 18
of the moving blade 5, deceleration from an intermediate
portion to a rear edge 19 is larger in actual Mach distribution
G2 than in design Mach distribution G1.
When the deceleration is large, as shown in Fig. 19A,
a boundary layer (a portion provided with shaded lines) 2,0
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at a portion from the intermediat,e portion to the rear edge
19 swells on the rear surface 11 of the tip portion 18 of
the moving blade 5. As a result, the pressure loss becomes
large, and the turbine efficiency is lowered by that amount.
A reference number 21 in Fig. 19 denotes a front edge of
the tip portion 18 of the moving blade 5.
SUMMARY OF THE INVENTION
'It is an object of this invention to provide a blade
structure in a gas turbine capable of improving the turbine
efficiency by minimizing the pressure loss.
In the blade structure in a gas turbine according to
one aspect of this invention, an inlet included angle
at a tip portion of the stationary blade that is the
stationary blade at the rear stage of the moving blade having
the tip clearance is larger than a front-edge including
angle at other portions than the tip portion of the
stationary blade.
According to the above-mentioned aspect, a curve of
a relative relationship between the incidence angle and the
pressure loss becomes mi.ldbymaking the inlet included
angle large. It is possible to reduce the pressure loss
by that amount, and therefore, it becomes possible to improve
the turbine efficiency.
In the blade structure in a gas turbine according to
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another aspect of this invention, an entrance metal angle
at a tip portion of the stationary blade that is the
stationary blade at the rear stage of the moving blade having
the tip clearance is made smaller than an entrance metal
angle at other portions than the tip portion of the
stationary blade.
According to the above-mentioned aspect, it is
possible to make the incidence angle small by making the
entrance metal angle small. it is possible to reduce the
pressure loss by that amount, and therefore, it becomes
possible to improve the turbine efficiency.
In the blade structure in a gas turbine according to
still another aspect of this invention, a front-edge
including angle at a tip portion of the stationary blade
that is the stationary blade at the rear stage of the moving
blade having the tip clearance is made larger than
an inlet included angle at other portions than the tip
portion of the stationary blade, and also an entrance metal
angle at a tip portionof the stationaryblade is made smaller
than an entrance metal angle at other portions than the tip
portion of the stationary blade.
According to the above-mentioned aspect, a curve of
a relative relationship between the incidence angle and the
pressure loss becomes nmildbymaking the inlet included
angle large. It is possible to reduce the pressure loss
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by that amount, and therefore, it becomes possible to improve
the turbine efficiency. Moreover, it is possible to make
the incidence angle small by making the entrance metal angle
small. Also, it is possible to reduce the pressure loss
by that amount, and therefore, it becomes possible to improve
the turbine efficiency. Moreover, it is possible to make
the pressure loss much smaller based on a synergy effect
of the work that a curve of a relative relationship between
the incidence angle and the pressure loss becomes mild and
the work that the incidence angle can be made small.
In the blade structure in a gas turbine according to
still another aspect of this invention, an inlet included
angle at a hub portion of the stationary blade is made
larger than an inlet included angle at other portions than
the hub portion of the moving blade.
According to the above-mentioned aspect, a curve of
a relative relationship between the incidence angle and the
pressure loss becomes mildbymaking the inlet included
angle large. It is possible to reduce the pressure loss
by that amount, and therefore, it becomes possible to improve
the turbine efficiency.
In the blade structure in a gas turbine according to
still another aspect of this invention, an entrance metal
angle at a hub portion of the stationary blade is made smaller
than an entrance metal angle at other portions than the hub
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portion of the moving blade.
According to the above-mentioned aspect, it is
possible to make the incidence angle small by making the
entrance metal angle small. It is possible to reduce the
pressure loss by that amount, and therefore, it becomes
possible to improve the turbine efficiency.
In the blade structure in a gas turbine according to
still another aspect of this invention, an inlet
included angle at a hub portion of the stationary blade
is made larger than an inlet included angle at other
portions than the hub portion of the moving blade, and also
an entrance metal angle at a hub portion of the stationary
blade is made smaller than an entrance metal angle at other
portions than the hub portion of the moving blade.
According to the above-mentioned aspect, a curve of
a relative relationship between the incidence angle and the
pressure loss becomes mild by making the inlet included
angle large. It is possible to reduce the pressure loss
by that amount, and therefore, it becomes possible to improve
the turbine efficiency. Moreover, it is possible to make
the incidence angle small by making the entrance metal angle
small. It is possible to reduce the pressure loss by that
amount, and therefore, it becomes possible to improve the
turbine efficiency. Furthermore, it is possible to make
the pressure loss much smaller based on a synergy effect
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28964-53G
of the work that a curve of a relative relationship between
the incidence angle and the pressure loss becomes mild and
the work that the incidence angle can be made small.
In the blade structure in a gas turbine according
to still another aspect of this invention, a chord length at
a tip portion of the moving blade having the tip clearance
is made larger than a minimum chord length at other portions
than the tip portion of the moving blade.
According to the above-mentioned aspect, it is
possible to make small the deceleration from the
intermediate portion to the rear edge on the rear surface of
the tip portion of the moving blade by making the chord
length of the moving blade large. Then, it is possible to
minimize the swelling of the boundary layer. As a result,
it is possible to make the pressure loss small, and it
becomes possible to improve the turbine efficiency by that
amount.
According to one aspect of the present invention,
there is provided a blade structure in a gas turbine,
comprising: stationary blades arrayed in a circle on a
casing; moving blades arrayed in a circle on a rotor,
wherein a clearance is provided between tips of the moving
blades and the casing, wherein a chord length at a tip
portion of the moving blade is other than a minimum chord
length of the moving blade and a tip portion of the
stationary blade is provided with an escape section for
avoiding an interference with the tip portion of the moving
blade, wherein the escape section is a clearance to be
arranged in such a manner that an entrance metal angle at
the tip portion of the stationary blade is a minimum
entrance metal angle of the stationary blade, or that the
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28964-53G
entrance metal angle at the tip portion of the stationary
blade is directed toward a rear surface side of the
stationary blade.
Other objects and features of this invention will
become apparent from the following description with
reference to the accompanying drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
Fig. 1 is an explanatory diagram of a cross
section of a tip portion of a stationary blade showing a
first embodiment of a blade structure in a gas turbine
according to this invention.
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Fig. 2 is an explanatory diagram of a cross section
of a tip portion of a stationary blade showing a second
embodiment of a blade structure in a gas turbine according
to this invention.
Fig. 3 is an explanatory diagram of a cross section
of a tip portion of a stationary blade showing a third
embodiment of a blade structure in a gas turbine according
to this invention.
Fig. 4 is a perspective view of the stationary blade
of the same.
Fig. 5 is an explanatory diagram of a cross section
of a hub portion of a moving blade showing a fourth embodiment
of a blade structure in a gas turbine according to this
invention.
Fig. 6 is an explanatory diagram of a cross section
of a hub portion of a moving blade showing a fifth embodiment
of a blade structure in a gas turbine according to this
invention.
Fig. 7 is an explanatory diagram of a cross section
of a hub portion of a moving blade showing a sixth embodiment
of a blade structure in a gas turbine according to this
invention.
Fig. 8 is a perspective view of the moving blade of
the same.
Fig. 9 is an explanatory diagram of a cross section
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of a stacking shape of a moving blade showing a seventh
embodiment of a blade structure in a gas turbine according
to this invention.
Fig. 10 is a diagram of Fig. 9 viewed from a direction
of X.
Fig. 11 is a diagram of Fig. 9 viewed from a direction
of XI.
Fig. 12A is an explanatory diagram of a cross section
of a hub portion of a moving blade showing a chord length,
Fig. 12B is an explanatory diagram of a Mach number
distribution according the moving blade shown in Fig. 12A.
Fig. 13 is an explanatory diagram showing a
modification of the seventh embodiment of a blade structure
in a gas turbine according to this invention.
Fig. 14A is an explanatory diagram of a cross section
of a moving blade and a stationary blade showing a
conventional blade structure, and Fig. 14B is an explanatory
diagram of a cross section of a moving blade and a stationary
blade showing a modification of the seventh embodiment of
a blade structure in a gas turbine according to this
invention.
Fig. 15A is an explanatory diagram of a cooling moving
blade showing a modification of the seventh embodiment of
a blade structure in a gas turbine according to this invention,
and Fig. 15B is an explanatory diagram of a moving blade
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having a taper according to the same.
Fig. 16 is an explanatory diagram of a moving blade
and a stationary blade showing a conventional blade
structure.
Fig. 17 is an explanatory diagram of a cross section
of a moving blade and a stationary blade showing a
conventional blade structure.
Fig. 18 is an explanatory diagram showing a relative
relationship between an incidence angle and a pressure loss.
Fig. 19A is an explanatory diagram of a cross section
of a hub portion of a moving blade showing a conventional
blade structure, and Fig. 19B is an explanatory diagram of
a Mach number distribution according to the moving blade
shown in Fig. 19A.
DETAILED DESCRIPTIONS
Embodiments of a blade structure in a gas turbine
relating to this invention will be explained below with
reference to the accompanying drawings. Itshould be noted
that the blade structure in the gas turbine is not limited
to these embodiments.
Fig. 1 is an explanatory diagram showing a first
embodiment of a blade structure in a gas turbine relating
to this invention. In the drawing, reference numbers that
are the same as those in Fig. 16 to Fig. 19 show the identical
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portions.
A blade structure in a first embodiment relates to
a stationary blade 3 at the rear stage of a moving blade
having a tip clearance. An inlet included angle Ac1 at a
front edge of a tip portion (a cross section of a chip) of
the stationary blade 3 is made larger than an inlet
included angle of portions (a cross section of a hub portion
to a mean portion) other than the tip portion of this
stationary blade 3. For example, this is m,ade larger than
about 5 .
According to the blade structure of this first
embodiment, the inlet included angle 9cl is taken large
at the tip portion of the stationary blade 3 at the rear
stage of the moving blade having the tip clearance. With
this arrangement, a curve of a relative relationship between
the incidence angle and the pressure loss becomes mild as
shown by a broken-line curve in Fig. 18. As a result, it
is possible to make the pressure loss small as shown by a
point P3 in Fig. 18. Therefore, it becomes possible to
iznprove the turbine efficiency.
Fig. 2 is an explanatory diagram showing a second
embodiment of a blade structure in a gas turbine relating
to this invention. In the drawing, reference numbers that
are the same as those in Fig. 1 and Fig. 16 to Fig. 19 show
the identical portions.
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A blade structure in a second embodiment relates to
a stationary blade 3 at the rear stage of a moving blade
having a tip clearance. An entrance metal angle (3c1 of
a tip portion (a cross section of a chip) of this stationary
blade 3 is made smaller than an entrance metal angle of
portions (a cross section of a hub portion to a mean portion)
other than the tip portion of this stationary blade 3. In
other words, the entrance metal angle Ocl of the cross section
of the tip portion of the stationary blade 3 is directed
toward a rear surface 13 side by about 10 , for example,
as compared with the entrance metal angle of the cross section
of the hub portion to the mean portion.
According to the blade structure of this second
embodiment, the entrance metal angle Ocl is taken small at
the tip portion of the stationary blade 3 at the rear stage
of the moving blade having the tip clearance. Wi.th this
arrangement, it is possible to make an incidence angle icl
small as shown by a point P4 in Fig. 18. As a result, it
is possible to make the pressure loss small. Therefore,
it becomes possible to improve the turbine efficiency.
Fig. 3 and Fig. 4 are explanatory diagrams showing
a third embodiment of a blade structure in a gas turbine
relating to this invention. In the drawings, re f-erence
numbers that are the same as those in Fig. 1, Fig. 2 and
Fig. 16 to Fig. 19 show the identical portions.
CA 02506206 2002-01-11
A blade structure in a third embodiment relates to
a stationary blade 3 at the rear stage of a moving blade
having a tip clearance. An inlet included angle Acl
at a front edge of a tip portion (a cross section of a chip)
of the stationary blade 3 is made larger than an inlet
included angle of portions (a cross section of a hub portion
to a mean portion) other than the tip portion of this
stationary blade 3. For example, this is made larger than
about 5 .
Further, an entrance metal angle Pclof a tip portion
,
(a cross section of a tip) lof this stationary blade 3 is
made smaller than an entrance metal angle of portions (a
cross section of a hub portion to a mean portion) other than
the tip portion of this stationary blade 3. In other words,
the entrance metal angle (3c1 of the cross section of the
tip portion of the stationary blade 3 is directed toward
a rear surface 13 side by about 10 , for example, as compared
with the entrance metal angle of the cross section of the
hub portion to the mean portion.
According to the blade structure of this third
embodiment, the front-edge including angle 8ci is taken large
at the tip portion of the stationary blade 3 at the rear
stage of the moving blade having the chip clearance. With
this arrangement, a curve of a relative relationship between
the incidence angle and the pressure loss becomes mild as
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CA 02506206 2002-01-11
shown by the broken-line curve in Fig. 18. As a result,
it is possible to make the pressure loss small as shown by
the point P3 in Fig. 18. Therefore, it becomes possible
to improve the turbine efficiency.
Further, according to the blade structure of this third
embodiment, the entrance metal angle Ocl is taken small at
the tip portion of the stationary blade 3 at the rear stage
of the moving blade having the tip clearance. With this
arrangement, it is possible to make an incidence angle icl
small as shown by the point P4 in Fig. 18. As a result,
it is possible to make the pressure loss small. Therefore,
it becomes possible to improve the turbine efficiency.
Particularly, according to the blade structure of this
third embodiment, it is possible to make the pressure loss
much smaller, based on a synergy effect of the work that
a curve of a relative relationship between the incidence
angle and the pressure loss becomes mild as shown by the
broken-line curve in Fig. 18 and the work that the incidence
angle icl can be made small as shown by a point P5 in Fig.
18. As a result, it becomes possible to improve the turbine
efficiency. I
Fig. 5 is an explanatory diagram showing a first
embodiment of a blade structure in a gas turbine relating
to this invention. In the drawing, reference numbers that
are the same as those in Fig. 1 to Fig. 4 and Fig. 16 to
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CA 02506206 2002-01-11
Fig. 19 show the identical portions.
A blade structure in a fourth embodiment re_lates to
a moving blade 5 like a free-standing movinq blade and a
shrouded moving blade. An inlet included angle 6si
at a hub portion (a cross section of a hub portion) of
this moving blade 5 is made larger than an inlet included
angle of portions (a cross section of a tip portion to a
mean portion) other than the hub portion of this moving blade
5. For example, this is made larger than about 5 .
According to the blade structure of this fourth
embodiment, the f ront-edge including angle 8siistaken large
at the hub portion of this moving blade 5. With this
arrangement, a curve of a relative relationship between the
incidence angle and the pressure loss becomes mild as shown
by the broken-line curve in Fig. 18. As a result, it is
possible to make the pressure loss small as shown by the
point P3 in Fig. 18. Therefore, it becomes possible to
improve the turbine efficiency.
Fig. 6 is an explanatory diagram showing a fifth
embodiment of a blade structure in a gas turbine relating
to this invention. In the drawing, reference numbers that
are the same as those in Fig. 1 to Fig. 5 and Fig. 16 to
Fig. 19 show the identical portions.
A blade structure in a fifth embodiment relates to
a moving blade 5 like a free-standing moving blade and a
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shrouded moving blade. An entrance metal angle 5s1 of a
hub portion (a cross section of a hub portion) of this moving
blade 5 is made smaller than an entrance metal angle of
portions (a cross section of a tip portion to a mean portion)
other than the hub portion of this moving blade 5. In other
words, the entrance metal angle (3s1 of the cross=section
of the hub portion of the moving blade 5 is directed toward
a rear surface 11 side by about 100, for example, as compared
with the entrance metal angle of the cross section of the
tip portion to the mean portion.
According to the blade structure of this fifth
embodiment, the entranct metal angle Osl is taken small at
the hub portion of the moving blade 5. With this arrangement,
it is possible to make an incidence angle isl small as shown
by the point P4 in Fig. 18. As a result, it is possible
to make the pressure loss small. Therefore, it becomes
possible to improve the turbine efficiency.
Fig. 7 and Fig. 8 are explanatory diagrams showing
a sixth embodiment of a blade structure in a gas turbine
relating to this invention. In the drawings, reference
numbers that are the same as those in Fig. 1 to Fig. 6 and
Fig. 16 to Fig. 19 show the identical portions.
A blade structure in a sixth embodiment relates to
a moving blade 5 like a free-standing movinq blade and a
shrouded moving blade. An inlet included angle 6s1
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CA 02506206 2002-01-11
at a hub portion (a cross section of a hub portion) of this
moving blade 5 is made larger than an inlet included
angle of portions (a cross section of a tip portion to a
mean portion) other than the hub portion of this moving blade
5. For example, this is made larger than about 5 .
Further, an entrance metal angle Osl of a hub portion
(a cross section of a hub portion) of this moving blade 5
is made smaller than an entrance metal angle of portions
(a cross section of a chip portion to a meark portion) other
than the hub portion of this moving blade 5. in other words,
the entrance metal angle Osl of the cross section of the
hub portion of the moving blade 5 is directed toward a rear
surface 11 side by about 10 , for example, as compared with
the entrance metal angle of the cross section of the tip
portion to the mean portion.
According to the blade structure of this sixth
embodiment, the front-edge including angle 6si is taken large
at the hub portion of thi.s moving blade 5. With this
arrangement, a curve of a relative relationship between the
incidence angle and the pressure loss becomes mild as shown
by the broken-line curve in Fig. 18. As a result, it is
possible to make the pressure loss small as shown by the
point P3 in Fig. 18. Therefore, it becomes possible to
improve the turbine efficiency.
Further, according to the blade structure of this sixth
CA 02506206 2002-01-11
embodiment, the entrance metal angle Os1 is taken small at
the hub portion of the moving blade 5. With this arrangement,
it is possible to make an incidence angle isl small as shown
by the point P4 in Fig. 18. As a result, it is possible
to make the pressure loss small. Therefore, it becomes
possible to improve the turbine efficiency.
Particularly, according to the blade structure of this
sixth embodiment, it is possible to make the pressure loss
much smaller, based on a synergy effect of the work that
a curve of a relative relationship between the incidence
angle and the pressure loss becomes mild as shown by the
broken-line curve in Fig. 18 and the work that the incidence
angle isl can be made small as shown by the point P5 in Fig.
18. As a result, it becomes possible to improve the turbine
efficiency.
Fig. 9 and Fig. 12 are explanatory diagrams showing
a seventh embodiment of a blade structure in a gas turbine
relating to this invention. In the drawings, reference
numbers that are the same as those in Fig. 1 to Fig. 8 and
Fig. 16 to Fig. 19 show the identical portions.
A blade structure in a seventh embodiment relates to
a moving blade 5 like a free-standing moving blade and a
shrouded moving blade. A chord length 26 at a tip portion
18 (a cross section of the tip portion 18) of this moving
blade 5 is made larger than a minimum chord length at other
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portions (a cross section of a hub portion to a mean section)
than the tip portion of the moving blade 5. In other words,
the chord length 26 of the cross section of the tip portion
18 is made equal to or larger than the chord length of the
mean cross section (a ratio of pitch to chord is set larger
than a conventional ratio).
Fig. 9 is an explanatory diagram of a cross section
showing a stacking shape of the moving blade 5. in Fig.
9 to Fig. 11, a stacking shape shown by a reference number
50 and a solid line show a tip. A stacking shape shown
by a reference number 51 and a one-dot chained line show
a tip at a position of about 75% of the height from a hub.
Further, a stacking shape shown by a reference number 52
and a two-dot chained line show a mean. Further, a stacking
shape shown by a reference number 53 and a three-dot chained
line show a tip at a position of about 25% of the height
from the hub. Last, a stacking shape shown by a reference
number 54 and a broken line show the hub.
According to the blade structure of this sixth
embodiment, it is possible to make small the deceleration
from an intermediate portion to a rear edge 19 on a rear
surface 11 of a tip portion 18 of a moving blade 5, as shown
by G4 in Fig. 12B, by making large a chord length 26 of the
tip portion 18 of the moving blade 5.
Namely, in Mach number distributions in Fig. 12B and
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Fig. 19B, an area of a portion encircled by a solid-line
curve (an area of a portion provided with shaded lines, and
a pr=essure difference) S is constant. In this case, when
the chord length 26 of the tip portion 18 of the moving
blade 5 is made large, the area S of the Mach number
distribution changes from a vertically-long shape shown in
Fig. 19B to a laterally-long shape shown in Fig. 12B. As
a result, the deceleration changes frorn G2 shown in Fig.
19B to small G4 shown in Fig. 12B. Consequently, it is
possible to restrict the swelling of the boundary layer.
Therefore, it is possible to make the pressure loss small,
and it becomes possible to improve the turbine efficiency
by that amount.
Fig. 13 to Fig. 15 show modifications of a blade
structure in a gas turbine relating to this invention. In
these drawings, reference numbers that are the same as those
in Fig. 1 to Fig. 12 and Fig. 16 to Fig. 19 show the identical
portions.
First, amodification shown in Fig. 13 is amodification
of the seventh embodiment. Tip portions of stationary
blades 2 and 3 are provided with escape sections 27 for
avoiding an interference with a tip portion 18 of a moving
blade 5.
According to this seventh embodiment, there is no room
for mutual interference between the tip portion 18 of the
23
CA 02506206 2002-01-11
moving blade 5 and the tip portions of the stationary blades
2 and 3 adjacent to each other, even when the chord length
26 of the tip portion 18 of the moving blade 5 is made large.
A two-dot chained line in Fig. 13 shows a conventional blade
structure.
Next, amodification shown in Fig. 148 is amodification
of the seventh embodiment. As an escape section of the tip
portion of the stationary blade 3, the entrance metal angle
(3c1 of the chip portion of the stationary blade 3 is made
smaller than the entrance metal angle of portions (the hub
portion to the mean portion) other than the tip portion
of the stationary blade 3. In other words, as shown in Fig.
2, Fig. 3 and Fig. 4, the entrance metal angle (3c1 of the
tip portion of the stationary blade 3 is directed toward
the rear surface 13 side of the stationary blade 3. It is
also possible to have a similar structure for the stationary
blade 2 at the same stage as that of the moving blade S.
According to the modification shown in this Fig. 14B,
as the entrance metal angle Oc1 of the tip portion of the
stationary blade 3 is directed toward the rear surface 13
side of the stationary blade 3, it is possible to have a
width Wi in an axial direction of the stationary blade 3
smaller than a width W2 of a conventional moving blade shown
in Fig. 14A. As a result, even when a width W3 in the axial
direction of the moving blade 5 is made larger than a
24
CA 02506206 2002-01-11
conventional width W4 by increasing the chord length 26 of
the tip portion 18 of the moving blade 5, a width W5 from
the moving blade 5 to the stationary blade 3 makes little
change from a conventional width W6. Therefore, there is
no room for mutual interference between the tip portion
18 of the moving blade 5 and the tip portion of the stationary
blade 3 adjacent to each other, even when the chord length
26 of the tip portion 18 of the moving blade 5 is made large.
Further, according to the modification shown in this
Fig. 14B, as the entrance metal angle Ocl of the tip portion
of the stationary blade 3 is smaller than the entrance metal
angle of the hub portion to the mean portion other than the
tip portion of the stationary blade 3, it becomes possible
to make the incidence angle icl small as shown by the point
P4 in Fig. 18. As it is possible to make the pressure loss
smaller by that amount, it becomes possible to improve the
turbine efficiency.
Then, the blade structure relating to this invention
can also be applied to a cooling moving blade 29 having a
hollow portion 28 at the tip portion 18, as shown in Fig.
15A. Further, it is also possible to apply the blade
structure relating to this invention to a moving blade 31
of which tip portion 18 has a taper 30 along the taper of
the casing 1, as shown in Fig. 15B.
As is clear from the above, according to the blade
CA 02506206 2002-01-11
structure in a gas turbine relatinq to one aspect af this
invention, an inlet included angle is taken large, at
a tip portion of a stationary blade at a rear stage of a
moving blade having a tip clearance. Therefore, a curv2
of a relative relationship between the incidence angle and
the pressure loss becomes mild. As it is possible to reduce
the pressure loss by that amount, it becomes possible to
improve the turbine efficiency.
According to the blade structure in a gas turbine
relating to another aspect of this invention, it is possible
to make an incidence angle small by making an entrance metal
angle small, at a tip portion of a stationary blade at a
rear stage of a moving blade having a clearance. As it is
possible to reduce the pressure loss by that amount, it
becomes possible to improve the turbine efficiency.
According to the blade structure in a gas turbine
relating to still another aspect of this invention,
an inlet included angle is taken large at a tip portion
of a stationary blade, at a rear stage of a moving blade
having a tip clearance. Therefore, a curve of a relative
relationship between an incidence angle and a pressure loss
becomes mild. As it is possible to reduce the pressure loss
by that amount, it becomes possible to improve the turbine
efficiency.
According to the blade structure in a gas turbine
26
CA 02506206 2002-01-11
relating to still another aspect of this invention, it is
possible to make an incidence angle small by making an
entrance metal angle small, at a chip portion of a stationary
blade at a rear stage of a moving blade having a clearance.
As it is possible to reduce the pressure loss by that amount,
it becomes possible to improve the turbine efficiency.
According to the blade structure'in a gas turbine
relating to still another aspect of this invention, it is
possible to make the pressure loss much smaller based on
a synergy effect of the work that a curve of a relative
relationship between an incidence angle and a pressure loss
becomes mild and the work that the incidence angle can be
made small. As a result, it becomes possible to improve
the turbine efficiency.
According to the blade structure in a gas turbine
relating to still another aspect of this invention, a curve
of a relative relationship between an incidence angle and
a pressure loss becomes mild by making a front-edge including
angle large at a hub portion of a moving blade. As it is
possible to reduce the pressure loss by that amount, it
becomes possible to improve the turbine efficiency.
According to the blade structure in a gas turbine
relating to still another aspect of this-invention, it is
possible to make an incidence angle small by making an
entrance metal angle small at a hub portion of a moving blade.
27
CA 02506206 2002-01-11
As it is possible to reduce the pressure loss by that amount,
it becomes possible to improve the turbine efficiency.
According to the blade structure in a gas turbine
relating to still another aspect of this invention, a curve
of a relative relationship between an incidence angle and
a pressure loss becomes mild by making an inlet included
angle large at a hub portion of a moving blade. As it is
possible to reduce the pressure loss by that amount, it
becomes possible to improve the turbine efficiency.
According to the blade structure in a gas turbine
relating to still another aspect of this invention, it is
possible to make an incidence angle small by making an
entrance metal angle small at a hub portion of a moving blade.
As it is possible to reduce the pressure loSs by that amount,
it becomes possible to improve the turbine efficiency.
According to the blade structure in a gas turbine
relating to still another aspect of this invention, it is
possible to make the pressure loss much smaller based on
a synergy effect of the work that a curve of a relative
relationship between an incidence angle and a pressure loss
becomes mild and the work that the incidence angle can be
made small. As a result, it becomes possible to improve
the turbine efficiency.
According to the blade structure in a gas turbine
relating to still another aspect of this invention, it is
28
CA 02506206 2002-01-11
possible to make small the deceleration froman intermediate
portion to a rear edge on a rear surface of a tip portion
of a moving blade by making a chord length of the moving
blade large. Then, it is possible to minimize the swelling
of the boundary layer. As a result, it is possible to make
the pressure loss small, and it becomes possible to improve
the turbine efficiency by that amount.
Furthermore, a tip portion of a stationary blade is
provided with an escape section for avoiding an interference
with a tip portion of a moving blade . As a result, there
is no room for mutual interference between a tip portion
of the moving blade and tip portions of stationary blades
adjacent to each other, even when a chord length of the tip
portion of the moving blade is made large.
Moreover, as an entrance metal angle at a tip portion
of a stationary blade is directed toward the rear surface
side of the stationary blade, there is no room for mutual
interference between a tip portion of a moving blade and
tip portions of stationary blades adjacent to each other,
even when the chord length of the tip portion of the moving
blade is made large.
Furthermore, as an entrance metal angle at a tip
portion of a stationary blade is smaller than an entrance
metal angle at other portions than the tip portion of the
stationary blade, it is possible to make an incidence angle
29
CA 02506206 2002-01-11
small. As it is possible to reduce the pressure loss by
that amount, it becomes possible to improve the turbine
efficiency.
Although the invention hasbee.n described with respect
to a specific embodiment for a complete and clear disclosure,
the appended claims are not to be thus limited but are to
be construed as embodying all modifications and alternative
constructions that may occur to one skilled in the art which
fairly fall within the basic teaching herein set forth.
