Information
-
Patent Grant
-
6733238
-
Patent Number
6,733,238
-
Date Filed
Friday, February 22, 200222 years ago
-
Date Issued
Tuesday, May 11, 200420 years ago
-
Inventors
-
Original Assignees
-
Examiners
- Look; Edward K.
- White; Dwayne J.
Agents
- Oblon, Spivak, McClelland, Maier & Neustadt P.C.
-
CPC
-
US Classifications
Field of Search
US
- 415 148
- 415 150
- 415 207
- 415 2081
- 415 2082
- 415 2101
- 415 2112
-
International Classifications
-
Abstract
There is provided an axial-flow turbine comprising an exhaust chamber; a turbine including multiple stage rotor blades, said multiple stage rotor blade including terminal stage rotor blades; an annular diffuser located between the turbine and the exhaust chamber; and an annular axial-flow turbine passage defined by the turbine, the diffuser and the exhaust chamber, wherein fluid flows through the axial-flow turbine passage toward the exhaust chamber, and an annular stepped portion which inwardly projects in a radial direction is formed on the portion of an inner wall of the axial-flow turbine passage that is located on the downstream side of a trailing edge of a tip portion of the terminal stage rotor blades provided in the flow direction of the fluid. In the stepped portion, a projecting portion which inwardly projects in a radial direction may be provided.
Description
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to an axial-flow turbine and, particularly, to a gas turbine in which the pressure between a turbine and a diffuser is locally increased so that the thermal efficiency is increased.
2. Description of the Related Art
In general, it has been required that the temperature in a turbine entrance and pressure ratio are further increased to improve the thermal efficiency of an axial-flow turbine, e.g. gas turbine.
Japanese Unexamined Patent Publications (Kokai) No. 5-321896 and No. 11-148497 disclose a solution in which the shape of the front side or the back side of a blade is modified so that the pressure loss caused by shock waves is decreased. In Kokai No. 5-321896, a blade, for example, a rotor blade in which the shape of the front side or the back side thereof is modified, is disclosed. In Kokai No. 11-148497, a blade, for example, a rotor blade in which the maximum thickness portion of the blade is changed from a position of 40% of a chord length to a position of 60% of the chord length, is disclosed.
However, in the above-described two related arts, only a part of the shape of a blade and, especially, only the shape of the front side or the back side of the blade is taken into account, and the shape of the tip portion of the blade is not taken into account. In general, a space between the tip portion of a blade, especially, a rotor blade and the inner wall of an axial-flow turbine passage e.g. a gas turbine passage, substantially does not exist, and they are located in contact with each other. Therefore, in order to further reduce the pressure loss caused by shock waves to increase the efficiency, not only the shape of the front side or the back side of the blade but also the shape of the tip portion of the blade and the inner wall of the axial-flow turbine passage adjacent to the tip portion should be taken into account.
Accordingly, the object of the present invention is to further reduce the pressure loss, caused by shock waves in the vicinity of a tip portion trailing edge of terminal stage rotor blades, so as to improve the efficiency of the axial-flow turbine by modifying the shape of the tip portion of the blades and the shape of the axial-flow turbine passage e.g. the gas turbine passage.
SUMMARY OF THE INVENTION
According to an embodiment of the present invention, there is provided an axial-flow turbine comprising an exhaust chamber; a turbine including multiple stage rotor blades, said multiple stage rotor blades including terminal stage rotor blades; an annular diffuser located between the turbine and the exhaust chamber; and an annular axial-flow turbine passage defined by the turbine, the diffuser and the exhaust chamber, wherein fluid flows through the axial-flow turbine passage toward the exhaust chamber, and an annular stepped portion which inwardly projects in a radial direction is formed on the portion of an inner wall of the axial-flow turbine passage that is located on the downstream side of a trailing edge of a tip portion of the terminal stage rotor blades provided in the flow direction of the fluid.
In other words, according to the embodiment of the present invention, the streamline of a fluid passing through the axial-flow turbine passage is inwardly curved between the tip portion trailing edge and the upstream end portion of the stepped portion so that variations in the streamline occurs. Therefore, the pressure is increased to reduce the Mach number, and the pressure loss is decreased to improve the turbine efficiency. Additionally, the Mach number is decreased to reduce the occurrence of shock waves and, thus, damage to the tip portion of the rotor blade can be prevented.
These and other objects, features and advantages of the present invention will be more apparent in light of the detailed description of exemplary embodiments thereof as illustrated by the drawings.
BRIEF DESCRIPTION OF THE DRAWING
The present invention will be more clearly understood from the description as set below with reference to the accompanying drawings, wherein:
FIG. 1
is a longitudinal partly sectional view of a gas turbine in a related art;
FIG. 2
is an enlarged view of the surroundings of a turbine and a diffuser of a gas turbine in a related art;
FIG. 3
is a longitudinal partly sectional view of a first embodiment of a gas turbine according to the present invention;
FIG. 4
is a longitudinal partly sectional view of a second embodiment of a gas turbine according to the present invention;
FIG. 5
is an enlarged view of another embodiment of the surroundings of the tip portion of a terminal stage rotor blade of a gas turbine according to the present invention;
FIG. 6
is a view showing the shape of a gas turbine according to the present invention; and
FIG. 7
is a view showing the rising rate of the turbine efficiency of a gas turbine.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
Before proceeding to a detailed description of the preferred embodiments, a prior art will be described with reference to the accompanying relating thereto for a clearer understanding of the difference between the prior art and the present invention.
FIG. 1
shows a longitudinal partly sectional view of an axial-flow turbine, e.g. a gas turbine in a related art. An axial-flow turbine, e.g. a gas turbine
110
contains a compressor
130
to compress intaken air, at least one combustor
140
provided on the downstream side of the compressor
130
in the direction of the air flow, a turbine
150
provided on the downstream side of the combustor
140
, a diffuser
160
provided on the downstream side of the turbine and an exhaust chamber
170
provided on the downstream side of the diffuser
160
. In the axial-flow turbine e.g. the gas turbine
110
, the compressor
130
, the turbine
150
, the diffuser
160
and the exhaust chamber
170
define an annular axial-flow turbine passage e.g. gas turbine passage
180
.
The compressor contains, in a compressor casing
139
, compressor rotor blades and compressor stay blades composed of multiple-stages. The turbine
150
contains, in the turbine casing
159
, rotor blades and stay blades composed of multiple-stages. As shown in the drawing, the compressor
130
and the turbine
150
are provided on a rotating shaft
190
. The turbine
150
has the multiple-stage stay blades which is provided on the inner wall of the gas turbine passage
180
and the multiple-stage rotor blades provided on the rotating shaft
190
. At each stage of the multiple-stage rotor blades, a plurality of rotor blades are spaced substantially at an equal distance, in the circumferential direction, around the rotating shaft
190
.
Fluid, for example, air enters through the inlet (not shown) of the compressor
130
and passes through the compressor
130
to be compressed. The fluid is mixed, in the combustor
140
, with the fuel to be burnt, and passes through the turbine
150
provided with multiple-stage blades, for example, four-stage blades. Then, the fluid is discharged through the exhaust chamber
170
via the diffuser
160
.
FIG. 2
shows an enlarged view of surroundings of the turbine
150
and the diffuser
160
of the gas turbine
110
. In
FIG. 2
, a rotor blade
151
of the terminal stage rotor blades of the turbine
150
is shown. For the purpose of understanding, blades other than the terminal stage rotor blades are omitted. As shown in
FIG. 2
, the tip portion of the rotor blade
151
substantially linearly extends along the inner wall of the gas turbine passage
180
. As shown in
FIG. 2
, the inner wall of the gas turbine passage
180
in the turbine
150
is formed so that the radius of the inner wall is increased toward the downstream side in the direction of the air flow (indicated by an arrow âFâ). Likewise, the inner wall of the gas turbine passage
180
in the diffuser
160
is formed so that the radius of the inner wall is increased toward the downstream side. Therefore, the fluid which passes through the turbine
150
enters into the diffuser
160
while outwardly and radially spreading from the rotating shaft
190
.
If the operating temperature and pressure of the gas turbine is enhanced to improve the thermal efficiency, the mechanical load of the turbine itself is increased. In other words, the velocity of the fluid increases and the Mach number increases in the vicinity of the tip portion of the rotor blade
151
. Particularly, in the vicinity of the trailing edge of the tip portion
156
of the terminal stage rotor blade
151
as shown in
FIG. 2
, the Mach number is extremely increased. As a result, pressure loss caused by shock waves tends to increase. Moreover, the tip portion of the rotor blades may be partially broken by the shock wave produced by increasing the Mach number as described above.
FIG. 3
shows a longitudinal partly sectional view of a first embodiment of the axial-flow turbine, e.g. a gas turbine according to the present invention. As described above, in
FIG. 3
, the surroundings of a turbine
50
and a diffuser
60
are enlarged. The turbine
50
contains a terminal stage rotor blade
51
of terminal stage rotor blades. For the purpose of understanding, blades other than the terminal stage rotor blade are omitted in the drawing. As shown in
FIG. 3
, the inner wall of the axial-flow turbine passage e.g. a gas turbine passage
80
in the turbine
50
, is formed so that the radius of the inner wall is increased toward the downstream side in the direction of the air flow (indicated by an arrow âFâ). The inner wall of the gas turbine passage
80
in the diffuser
60
is formed so that the radius of the inner wall is increased toward the downstream side.
On the inner wall of the gas turbine passage
80
in the diffuser
60
, an annular stepped portion
20
is provided on the downstream side of the tip portion leading edge
56
of the rotor blade
51
. In the embodiment shown in
FIG. 3
, the stepped portion
20
inwardly and radially projects from a part of the inner wall of the gas turbine passage
80
, which is nearest to the tip portion trailing edge
56
of the rotor blade
51
, to the tip portion trailing edge
56
. An upstream end portion
21
of the stepped portion
20
and the tip portion trailing edge
56
are not in contact with each other. The stepped portion
20
extends from the upstream end portion
21
of the stepped portion
20
toward the downstream side and the exhaust chamber
70
(not shown) in the gas turbine passage
80
in the diffuser
60
. In the first embodiment, the stepped portion
20
has a linear portion
22
extending substantially in parallel with the central axis of a rotating shaft (not shown). If the stepped portion
20
has the linear portion
22
, the stepped portion
20
can be easily formed. The stepped portion
20
is slightly outwardly curved at a curved portion
23
, and outwardly extends, toward the downstream side, along the inner wall of the gas turbine passage
80
in the diffuser
60
.
In other words, in the first embodiment, the distance between the central axis of the rotating shaft and the upstream end portion
21
of the stepped portion
20
is substantially identical to that between the central axis and the tip portion trailing edge
56
of the rotor blade
51
. Thus, the stepped portion
20
causes the streamline which represents a flow direction of the fluid to vary so that the streamline is strongly curved between the stepped portion
20
and the tip portion trailing edge
56
and, especially, between the upstream side end portion
21
and the tip portion trailing edge
56
. Therefore, the pressure is locally increased at a portion in which the above-described variations in streamline are produced. Consequently, the Mach number is decreased between the stepped portion
20
and the tip portion trailing edge
56
and, especially, between the upstream end portion
21
and the tip portion trailing edge
56
, thus resulting in reduction of the pressure loss.
As described above, in the first embodiment, the distance between the central axis and the upstream end portion
21
is substantially identical to that between the central axis and the tip portion trailing edge
56
. However, as there is a possibility that variations in streamline may occur even if the distance between the central axis and the upstream end portion
21
is smaller than that between the central axis and the tip portion trailing edge
56
, the Mach number can be decreased to reduce the pressure loss. Additionally, as there is a possibility that variations in streamline may occur even if the distance between the central axis and the upstream end portion
21
is larger than that between the central axis and the tip portion trailing edge
56
and is smaller than that between the central axis and the inner wall of the gas turbine passage
80
in the diffuser
60
, the Mach number can be decreased to reduce the pressure loss.
FIG. 4
shows a longitudinal partly sectional view of a second embodiment of an axial-flow turbine, e.g. a gas turbine, according to the present invention. In the stepped portion
20
in the above-described embodiment, a linear portion
22
, extending from the upstream end portion
21
substantially in parallel with the central axis, is formed. However, in the second embodiment, the stepped portion
20
has a projecting portion
24
which further projects toward the inside. In other words, in the stepped portion
20
, there is a projecting portion in which the distance between the central axis and the upstream end portion
21
is smaller than that between the central axis and the tip portion trailing edge
56
. In the second embodiment, the projecting portion
24
exists on the downstream side of the linear portion
22
of the stepped portion
20
.
Similar to the first embodiment, the stepped portion
20
causes the streamline which represents the flow direction of the fluid to vary so that the streamline is strongly inwardly curved between the stepped portion
20
and the tip portion trailing edge
56
, along the projecting portion
24
. Therefore, the pressure is locally increased at a portion in which variations in streamline occurs. Consequently, the Mach number is further decreased between the stepped portion
20
and the tip portion trailing edge
56
, thus resulting in a reduction in the pressure loss.
As a matter of course, the projecting portion
24
can be disposed to be adjacent to the upstream end portion
21
without having the linear portion
22
in the second embodiment. In this case, since larger variations in the streamline occur, the pressure loss can be further decreased and the turbine efficiency can be further increased. Similar to the first embodiment, if the distance between the central axis and the upstream end portion
21
is smaller than that between the central axis and the tip portion trailing edge
56
, and if the distance between the central axis and the upstream end portion
21
is larger than that between the central axis and the tip portion trailing edge
56
and is smaller than that between the central axis and the inner wall of the diffuser
60
, there is a possibility that a variation in streamline may occur. Therefore, the Mach number can be decreased to decrease the pressure loss, and the turbine efficiency can be increased.
FIG. 5
shows an enlarged view of another embodiment of surroundings of the tip portion of a terminal stage rotor blade of an axial-flow turbine, e.g. a gas turbine, according to the present invention. In a related art, a portion between the tip portion leading edge and the tip portion trailing edge of the terminal stage rotor blade
151
substantially linearly extends. However, in this embodiment, a curved portion
57
which is outwardly curved in a radial direction is provided between the tip portion leading edge
54
and the tip portion trailing edge
56
of the terminal stage rotor blade
51
.
When fluid is introduced into the axial-flow turbine passage e.g. a gas turbine passage
80
, the streamline of the fluid is inwardly curved in a radial direction on the downstream side of the curved portion
57
. Therefore, the streamline in the vicinity of the tip portion trailing edge
56
is curved more than that of a related art. Consequently, Mach number is decreased as the pressure is increased, and the pressure loss can be decreased.
In this embodiment, a maximum curvature point
58
in which a curvature of the curved portion
57
reaches maximum is located on the downstream side of an axial direction center line
59
of the terminal stage rotor blade
51
in the flow direction of the fluid. Therefore, the variations in streamline in this embodiment are larger than that in case of the maximum curvature point
58
in the curved portion
57
located on the upstream side of the axial direction center line
59
or located on the axial direction center line
59
. Accordingly, in this embodiment, the Mach number can be further decreased and the pressure loss can be further decreased.
As a matter of course, the first embodiment or the second embodiment can be combined with this embodiment, so that the pressure loss can be further decreased to further increase the turbine efficiency. Additionally, the shape of turbine blades and a gas turbine passage in a diffuser can be applied to the shape of a compressor blades and a gas turbine passage in a compressor.
EXAMPLE
FIG. 6
is a view showing the shape of an axial-flow turbine, e.g. a gas turbine, according to the present invention. In
FIG. 6
, the horizontal axis represents an axial length of a gas turbine, and the vertical axis represents a distance from the central axis of a rotating shaft. In
FIG. 6
, the thick line represents a gas turbine in a related art, the thin line represents a gas turbine (having only a linear portion
22
) based on the first embodiment, and the dotted line represents a gas turbine (having a projecting portion
24
on the downstream side of the linear portion
22
) based on the second embodiment, respectively.
FIG. 7
shows the rising rate of turbine efficiency of an axial-flow turbine, e.g. a gas turbine, for each of these embodiments. According to the present invention, the gas turbine efficiency can be improved by 0.13% in the first embodiment, and by 0.20% in the second embodiment.
Further, it will be apparent to those skilled in the art that the present invention can be applied to steam turbines.
According to the present invention, there can be obtained common effects in which the streamline of the fluid which flows through an axial-flow turbine passage e.g. a gas turbine passage, is curved so that the Mach number can be decreased to decrease the pressure loss, and the turbine efficiency can be increased. Additionally, there can be obtained common effects in which the Mach number is decreased to decrease the shock waves so that damage to the tip portions of rotor blades can be decreased.
Moreover, according to the present invention, there can be obtained effects in which the shape of a stepped portion is modified to further curve the streamline of the fluid so that the pressure loss can be further decreased and the turbine efficiency can be further increased.
Moreover, according to the present invention, can be obtained effects in which the streamline that passes between the upstream end portion and the tip portion trailing edge is curved along the projecting portion so that the Mach number and the pressure loss can be decreased to increase the turbine efficiency.
Moreover, according to the present invention, there can be obtained effects in which the streamline of the fluid is inwardly curved, in a radial direction, on the downstream side of the tip portion trailing edges of the terminal stage rotor blades so that the pressure loss can be decreased and the turbine efficiency can be increased.
Although the invention has been shown and described with exemplary embodiments thereof, it will be understood by those skilled in the art that various changes, omissions and additions may be made therein and thereto without departing from the spirit and the scope of the invention.
Claims
- 1. An axial-flow turbine comprisingan exhaust chamber; a turbine including multiple stage rotor blades, said multiple stage rotor blades including terminal stage rotor blades, an annular diffuser located between the turbine and the exhaust chamber; and an annular axial-flow turbine passage defined by the turbine, the diffuser and the exhaust chamber, wherein fluid flows through the axial-flow turbine passage toward the exhaust chamber, and an annular stepped portion which inwardly projects in a radial direction is formed on the portion of an inner wall of the axial-flow turbine passage that is located on the downstream side of a trailing edge of a tip portion of the terminal stage rotor blades provided in the flow direction of the fluid, and the distance between the central axis of the turbine and the stepped portion is substantially identical to that between the central axis of the turbine and the tip portion trailing edge of the terminal stage rotor blades.
- 2. An axial-flow turbine according to claim 1, wherein the upstream end portion of the stepped portion located on the upstream side in the flow direction of the fluid is located at the inner wall of the axial-flow turbine adjacent to the tip portion trailing edge of the terminal stage rotor blades.
- 3. An axial-flow turbine according to claim 1 or 2, wherein the stepped portion has a linear portion which extends from the upstream end portion of the stepped portion located on the upstream side in the flow direction of the fluid, substantially in parallel with the central axis of the turbine.
- 4. An axial-flow turbine according to claim 1 or 2, wherein the stepped portion has a projecting portion which radially projects from the inner wall of the axial-flow turbine more inwardly than the tip portion trailing edge of the terminal stage rotor blades.
- 5. An axial-flow turbine according to claim 4, wherein the projecting portion is disposed downstream of the linear portion.
- 6. An axial-flow turbine according to claim 1 or 2, wherein the terminal stage rotor blades have a curved portion which is radially and outwardly curved between a tip portion leading edge and the tip portion trailing edge of the terminal stage rotor blades.
- 7. An axial-flow turbine according to claim 6, wherein the maximum curvature point of the curved portion is located on the downstream side of a center line of the terminal stage rotor blades in the axial direction in the flow direction of the fluid.
Priority Claims (1)
Number |
Date |
Country |
Kind |
2001-132962 |
Apr 2001 |
JP |
|
US Referenced Citations (1)
Number |
Name |
Date |
Kind |
3625630 |
Soo |
Dec 1971 |
A |
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Number |
Date |
Country |
216489 |
Aug 1941 |
CH |
1227217 |
Jul 2002 |
EP |
996 967 |
Dec 1951 |
FR |
1338515 |
Sep 1963 |
FR |
5-321896 |
Dec 1993 |
JP |
8-260905 |
Oct 1996 |
JP |
11-148497 |
Jun 1999 |
JP |