Note: Descriptions are shown in the official language in which they were submitted.
CA 02549313 2002-01-09
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BLADE OF A GAS TURBINE
RELATED APPLICATION
This application is a divisional of Canadian
Patent Application Serial No. 2,366,969 filed January 9,
2002.
FIELD OF THE INVENTION
The present invention relates to a blade, of a gas
turbine, having a wide turning angle and suitable to a heavy
duty and high load gas turbine.
BACKGROUND OF THE INVENTION
General blades of a gas turbine will be explained
by referring to Fig. 7 to Fig. 12. A gas turbine generally
comprises plural stages of stationary blades disposed
annularly in a casing (blade ring or chamber), and plural
stages of moving blades 1 disposed annularly in a rotor (hub
or base). Two adjacent moving blades 1 are shown in Fig. 7.
The moving blade 1 is composed, as shown in
Fig. 7, of a front edge 2, a rear edge 3, and a belly (or a
belly side) 4 and a back (or a back side) 5 linking the
front edge 2 and rear edge 3. Combustion gases G1, G2, as
shown in Fig. 7, flow in a passage 6 between the belly 4 and
back 5 of two adjacent moving blades 1 at an influent angle
a1 (G1), and turn and flow out at an effluent angle a2 (G2).
By the flow of combustion gases G1, G2, the rotor rotates in
a direction of blank arrow U through the moving blades 1.
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The width of the passage 6 ("passage width") of the
moving blades 1 in which the combustion gases G1, G2 flow
gradually decreases from the front edge 2 to the rear edge
3 as indicated by solid line curve in Fig. 8. At the rear
end 3, the width is minimum, that is, throat 0. Thus, by
narrowing the passage width between the moving blades 1,
along the direction of flow of the combustion gases G1 and
G2, the combustion gases G1 and G2 are expanded and
accelerated, and the turbine efficiency is enhanced.
Recently, in the field of gas turbine, the mainstream
is the gas turbine of high load with the pressure ratio of
or more and the turbine inlet gas temperature of 1400
degree centigrade or more.
As the gas turbine of high load, the following two
15 types are known. One is a high load gas turbine in which
there are a large number, for example, from four to five,
of blades. The other is a high load gas turbine in which
the work of each blade of each stage is increased without
increasing the number of stages of blades, for example,
20 remaining at four stages. Of these two high load gas turbines,
the latter high load gas turbine is superior in the aspect
of the cost performance.
To increase the work AH of each blade in each stage,
it is required to increase the blade turning angle Aa as
shown in Fig. 9 and Fig. 10, and equations (1) and (2).
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OH = U x AVO ... (1)
AVO = V01 + v02 ...(2)
In equations (1) and (2), only the peripheral speed
component VO is defined in the absolute system, and the other
peripheral speed components are defined in the relative
system.
More specifically, symbol U denotes the peripheral
speed of moving blade 1. The peripheral speed U of moving
blade 1 is almost constant, being determined by the distance
from the center of rotation of the rotor and the tip of the
moving blade 1, and the rotating speed of the rotor and moving
blade 1. Accordingly, to increase the work OH of each blade
in each stage, it is first required to increase the difference
AVO between the peripheral speed components near the inlet
of the combustion gas Gl and outlet of the combustion gas
G2.
To increase the difference OVBbetween the peripheral
speed components, it is required to increase the peripheral
speed component V01 near the inlet of the combustion gas
Gl, and the peripheral speed component V02 near the outlet
of the combustion gas G2.
When the peripheral speed component V01 near the inlet
of the combustion gas G1 is increased, the influent angle
al becomes larger. When the peripheral speed component V02
near the outlet of the combustion gas G2 is increased, the
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effluent angle a2 becomes larger. When the influent angle
a1 and effluent angle a2 become larger, the turning angle
Da becomes larger (see Fig. 10). As a result, when the
turning angle Da is increased, the work AH of each blade
in each stage becomes larger.
Accordingly, as shown in Fig. 11 and Fig. 12, by setting
the influent angle a3 and effluent angle a4 larger than the
influent angle a1 and effluent angle a2 shown in Fig. 7,
it may be considered to increase the turning angle Dal larger
than the turning angle Aa shown in Fig. 10.
However, the following problems occurs when only the
influent angle a3 and effluent angle a4 are set larger. That
is, the passage width becomes the passage width as indicated
by single dot chain line curve shown in Fig. 8.
As a result, as shown in Fig. 8, a maximum width 7
occurs at a position behind the front edge 2, and a minimum
width 8 occurs at a position ahead of the rear edge 3, that
is, a width smaller than throat 0 is formed. Therefore,
as indicated by single dot chain line curve, a deceleration
passage (diffuser passage) is formed from the front edge
2 to the maximum width 7, and from the minimum width 8 to
the rear edge 3. Accordingly, the flow of the combustion
gases G1, G2 is decelerated, and the turbine efficiency loss
increases.
Thus, if only the blade turning angle is increased,
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the gas turbine with such blades is not suited to the heavy
duty and high load. The problem is the same in the
stationary blades as well as in the moving blades 1.
SUMMARY OF THE INVENTION
It is an object of the invention to present a
blade, of a gas turbine, having a wide turning angle and
suitable to a heavy duty and high load gas turbine.
The blade, according to the present invention, has
such a shape that the diameters of circles inscribing the
belly and back sides at different positions of adjacent
blades decreases as one goes from the front edge to the rear
edge. Since the blade has such a shape, even if the
influent angle and effluent angle of gases are increased, a
deceleration passage is not formed in the passage between
the adjacent moving blades.
According to one aspect, there is provided a
blade, of a gas turbine, having a turning angle of at least
115 , said blade having a belly side, a back side, a front
edge, and a rear edge, wherein diameter of circles
inscribing the belly side and the back side of adjacent
blades decrease gradually from the front edge to the rear
edge; and wherein a ratio of blade maximum wall thickness
and blade chordal length is 0.15 or more, and a wedge angle
of the rear edge is 10 degrees or less.
According to another aspect, there is provided a
blade, of a gas turbine, having a turning angle of at least
115 , said blade having a belly side, a back side, a front
edge, and a rear edge, wherein diameter of circles
inscribing the belly side and the back side of adjacent
blades decrease gradually from the front edge to the rear
edge; and wherein said blade is a cooling blade comprising a
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cooling passage which is near the rear edge, and the ratio
of the distance from the cooling passage to the rear edge
and the wall thickness of rear edge of the blade is 2 or
less.
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 influent
angle, effluent angle, throat, rear edge wall thickness, and
distance from cooling passage to rear edge in the hub of
moving blades in a first embodiment of blade according to
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the present invention;
Fig. 2 is an explanatory diagram of showing a passage
of which diameter of inscribed circle of belly and back of
adjacent blades gradually decreases from front edge to rear
edge of the sante;
,
Fig. 3 is an explanatory diagram showing wall thickness,
maximum wall thickness, blade chordal length, wedge angle,
camber line, influent angle, and effluent angle of the same;
Fig. 4A is a graph showing characteristic of Tmax/C,
Fig. 4B is a graph showing characteristic of WA, and Fig.
4C is a graph showing characteristic of d/O;
Fig. 5 is a graph showing the relation of turbine
efficiency and turning angle in the blade of Gas turbines
of the invention and the conventional blade of Gas turbines;
Fig. 6 is a graph showing the relation between the
turbine.efficiency loss and wedge angle;
Fig. 7 is an explanatory diagram of influent angle,
effluent angle, and throat in the hub of moving blades showing
the conventional turbine blades;
Fig. 8 is a graph showing an inappropriate passage
width;
Fig. 9 is an explanatory diagram showing direction
of influent side combustion gas and direction of effluent
side combustion gas;
Fig. 10 is an explanatory diagram showing the turning
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angle;
Fig. 11 is an explanatory diagram of a case with
an increased turning angle;
Fig. 12 is an explanatory diagram showing an
increased turning angle; and
Fig. 13 is a graph showing an ideal passage width.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
Embodiment of the blade of the gas turbine
according to this invention will be explained by referring
to Fig. 1 to Fig. 6 and Fig. 13. It must be noted, however,
that the invention is not limited to this embodiment alone.
In the drawings, same parts as in Fig. 7 to Fig. 12 are
identified with same reference numerals.
The blade of the embodiment, that is, the moving
blade 10 is large in the influent angle a3 and effluent
angle a4, and also large in the turning angle Aa1. For
example, the effluent angle a4 is about 60 to 70 degrees,
and the turning angle Lal is about 115 to 150 degrees.
Since the moving blade 10 has wider turning angle Lal (than
the conventional one), this blade is ideal and suited for
the heavy duty and high load gas turbine.
In the moving blade 10, as shown in Fig. 2,
diameters R1, R2, R3, and R4 of inscribed circles 91, 92,
93, and 94 of the belly 4 and back 5 of adjacent moving
blades 10 are designed to be smaller from the front edge 2
to the rear
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edge 3.
That is, the passage 6 is formed in the relation of
diameter R1 of solid line inscribed circle 91 (circle
inscribing at front edge 2) > diameter R2 of single-dot chain
line inscribed circle 92 > diameter R3 of double-dot chain
line inscribed circle 93 > diameter R4 (throat 0) of broken
line inscribed circle 94 (circle inscribing at rear edge
3).
The moving blades 10 of the embodiment are thus composed,
and if the influent angle a3 and effluent angle a4 are
increased, deceleration passage is not formed in the passage
6 between adjacent moving blades 10. Therefore, the moving
blades 10 of the embodiment present moving blades ideal for
a gas turbine of large turning angle Aal, heavy.work, and
high load.
A comparison of the efficiency of the conventional
blades (moving blades 1) and the moving blades 10 of the
embodiment will be undertaken by referring to Fig. 5. That
is, in case of the conventional blade, as indicted in the
shaded area enclosed by solid line curve in Fig. 5, when
the turning angle Aal is more than about 115 degrees, the
turbine efficiency drops suddenly. On the other hand, in
the moving blades 10 of the embodiment, as indicated by broken
line in Fig. 5, even if the turning angle Dal is more than
about 115 degrees, a high turbine efficiency is maintained.
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Fig. 3 is an explanatory diagram showing a specific
configuration of the moving blade 10. In this blade, the
turning angle Aal is about 1.15 to 150 degrees. The ratio
Tmax/C of maximum wall thickness Tmax of moving blade 10
and blade chordal length C is about 0.15 or more. The wedge
angle WA of the rear edge of the moving blade 10 is abou't
degrees or less.
The manufacturing process (design process) of the
moving blade 10 is explained by referring to Fig. 3. First,
10 the influent anqle a3 and effluent angle a4 are determined.
Along the turning angle Aal determined from the influent
angle a3 and effluent angle a4, a camber line 9 is determined.
Then the wedge angle WA of the rear edge is determined. The
wall thickness T and Tmax of the moving blade 10 are determined.
As a result, the moving blade 10 can be manufactured.
The ratio Tmax/C of maximum wall thickness Tmax of
moving blade 10 and blade chordal length C is about 0.15
or more in an area at the arrow direction side from straight
line L in the characteristic condition shown in the graph
in Fig. 4A. The wedge angle WA of the rear edge of the moving
blade 10 is about 10 degrees or less in an area at the arrow
direction side from straight line L in the characteristic
condition shown in the graph in Fig. 4B.
When these two characteristic conditionsaresatisfied,
the passage 6 indicated by solid line in Fig. 13 (as shown
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in Fig. 2, the passage 6 gradually decreased in diameters
R1, R2, R3, and R4 of inscribed circles 91, 92, 93, and 94
of the belly 4 and back 5 of adjacent moving blades 10 from
the front edge 2 to the rear edge 3) is determined
geometrically. That is, supposing the ratio Tmax/C of
maximum wall thickness Tmax of moving blade 10 and blade
chordal length C to be about 0.15 or more, the portion of
the maximum width 7 side indicated by single-dot chain line
in Fig. 8 is corrected so as to be along the solid line
curve of Fig. 13 as indicated by arrow. Supposing the wedge
angle WA of the rear edge of the moving blade 10 to be about
10 degrees or less, the portion of the minimum width 8 side
indicated by single-dot chain line in Fig. 8 is corrected so
as to be along the solid line curve of Fig. 13 as indicated
by arrow. Thus, the design of the moving blade 10 is easy.
Further, as shown in Fig. 6, if the wedge angle WA
of the rear edge of the moving blade 10 is more than about
10 degrees, the loss of turbine efficiency is significant,
but if it is smaller than about 10 degrees, the loss of
turbine efficiency is decreased,. In Fig. 6, the broken line
shows the moving blade 10 with the effluent angle a4 of 60
degrees, and the solid line shows the moving blade 10 with
the effluent angle a4 of 70 degrees.
The moving blade 10 includes a cooling moving
blade of which cooling passage 11 is near the rear edge 3 as
shown
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in Fig. 1. At the rear edge 3 of the cooling moving blade
10, there is an ejection port 12 for ejecting the cooling
ai r(a ). One or a plurality of ej ection ports 12 are provided
from the hub side to the tip side of the rear edge 3 of the
cooling moving blade 10.
The cooling moving blade 10 may be composed as shown
in Fig. 1. That is, the ratio d/O of the wall thickness
(d) of the rear edge 3 of the moving blade 10 and the throat
0 between the adjacent moving blades 10 is about 0.15 or
less.
The ratio d/O of the wall thickness (d) of the rear
edge 3 of the 'moving blade 10 and the throat 0 between the
adjacent moving blades 10 is about 0. 15 or less in an area
at the arrow direction side from the straight line L in the
characteristic condition shown in the graph in Fig. 4C.
When the characteristic condition is satisfied, even
in the case of the cooling moving blade 10 of which cooling
passage 11 is near the rear edge 3, the passage 6 indicated
by solid line in Fig. 13 (as shown in Fig. 2, the passage
6 gradually decreased in diameters R1, R2, R3, and R4 of
inscribed circles 91, 92, 93, and 94 of the belly 4 and back
5 of adjacent moving blades 10 from the front edge 2 to the
rear edge 3) is determined geometrically. Thus, the design
of the cooling moving blade 10 of which cooling passage 11
is near the rear edge 3 is easy.
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Further, in the cooZingmovingblade 10 of which cooling
passage 11 is near the rear edge 3, as shown in Fig. 1, the
ratio L1/d of the distance L1 from the cooling passage 11
to the rear edge 3 (regardless of presence or absence of
rear edge blow-out; however, the length of ejection port
12 in the presence of rear edge blow-out) and the blade rear-
edge wall thickness (d) is 2 or less.
When the characteristic condition is satisfied, same
as in case of the blade (moving blade 10),
even in the case of the cooling moving
blade 10 of which cooling passage 11 is near the rear edge
3, the passage 6 indicated by solid line _in Fig. 13 (as shown
in Fig. 2, the passage 6 gradually decreased in diameters
R1, R2, R3, and R4 of inscribed circles 91, 92, 93, and 94
of the belly 4 and back 5 of adjacent moving blades 10 from
the front edge 2 to the rear edge 3) is determined
geometrically. Thus, the design of the cooling moving blade
10 of which cooling passage 11 is near the rear edge 3 is
easy.
An explanation if given above about the moving blades .
However, this invention is applicable to stationary blades.
By applying the invention in the moving blades and stationary
blades, the flow of the combustion gases G1, G2 is smooth,
and the turbine efficiency is further enhanced.
The conditions in the embodiment (the turning angle
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Dal of about 115 to 150 degrees, the ratio Tmax/C of maximum
wall thickness Tmax and blade chordal length C of about 0. 15
or more, the wedge angle WA of the rear edge of about 10
degrees or less, the effluent angle a4 of 60 to 70 degrees,
the ratio d/O of wall thickness (d) of rear edge 3 and throat
0 of about 0.15 or less, and the ratio Ll/d of the distance
L1 from the cooling passage 11 to rear edge 3 and rear edge
wall thickness (d) of blade of 2 or less) may be satisfied
at least in the hub portion of the moving blades 10.
As explained above, according to the blade of this
invention, since the diameter of an inscribed circle of belly
side and back side of adjacent blades decreases gradually
from the front edge to the rear edge, if the influent angle
and effluent angle are set larger, deceleration passage is
not formed in the passage between adjacent blades.
Therefore, blade suited to a gas turbine of large turning
angle, heavy work, and high load can be presented.
Moreover, the turning angle is 115 degrees or more,
the ratio of blade maximum wall thickness and blade chordal
length is 0. 15 or more, and the wedge angle of the rear edge
is 10 degrees or less. As a result, the passage in which
the diameter of an inscribed circle of belly side and back
side of adjacent blades decreases gradually from the front
edge to the rear edge is determined geometrically.
Therefore, blade can be designed by an optimum design.
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Furthermore, in the case of the cooling blade of which
cooling passage is near the rear edge, the ratio of wall
thickness of rear edge and throat between adjacent blades
is 0. 15 or less. As a result, even in the case of the cooling
blade of which cooling passage is near the rear edge, the
passage in which the diameter of an inscribed circle of belly
side and back side of adjacent blades decreases gradually
from the front edge to the rear edge is determined
geometrically. Therefore, it is easy to design the cooling
blade of which cooling passage is near the rear edge.
Moreover, in the case of the cooling blade of which
cooling passage is near the rear edge, the ratio of the
distance from the cooling passage to the rear edge and the
wall thickness of rear edge of the blade is 2 or less.
As a result, even in the case of
the cooling blade of which cooling
passage is near the rear edge, the passage in which the
diameter of an inscribed circle of belly side and back side
of adjacent blades decreases gradually from the front edge
to the rear edge is determined geometrically. Therefore,
it is easy to design the cooling blade of which cooling passage
is near the rear edge.
Although the invention has been 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
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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.
