Information
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Patent Grant
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6390775
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Patent Number
6,390,775
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Date Filed
Wednesday, December 27, 200023 years ago
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Date Issued
Tuesday, May 21, 200222 years ago
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Inventors
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Original Assignees
-
Examiners
Agents
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CPC
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US Classifications
Field of Search
US
- 416 193 A
- 416 223 R
- 416 234
- 416 235
- 416 236 R
- 416 248
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International Classifications
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Abstract
A gas turbine moving blade includes a platform which is undercut with a groove. The groove extends from the concave side to the trailing edge side of the platform, where the groove exits the platform. The groove has a depth which will enter a stress line causing a change to the load path direction away from the trailing edge. The location and depth of the groove reduces both high thermal stress and mechanical stress arising at a connection portion of a blade trailing edge and the platform of the gas turbine air cooled moving blade during transient engine operation as well as steady state, full speed, full load conditions.
Description
BACKGROUND OF THE INVENTION
The present invention relates to a gas turbine moving blade, and, more particularly, to a gas turbine blade having a platform undercut with improved thermal stress relief.
Gas turbine blades, also referred to as buckets, are exposed to high temperature combustion gases, and, consequently, are subject to high thermal stresses. Methods are known in the art for cooling the blades and reducing the thermal stresses.
FIGS. 1-3
show one example of a prior art air-cooled moving blade. High pressure air
2
, discharged from a compressor, is introduced into an interior of an air-cooled blade from a blade root bottom portion
4
. The high pressure air, after cooling a shank portion
6
, a platform
8
and a blade profile portion (or airfoil)
10
, flows out of fine holes
12
provided at a blade face, or out of fine holes
14
provided at a blade tip portion. Also, fine holes
12
are provided at a blade trailing edge portion
13
of the blade, through which the high pressure air flows to cool the trailing edge of the blade. Thus, the high pressure air cools the metal temperature of the moving blade.
Highly cooled gas turbine buckets experience high temperature mismatches at the interface of the hot airfoil and the relatively cooler shank portion of the bucket platform. These high temperature differences produce thermal deformations at the bucket platform, which are incompatible with those of the airfoil. In the prior art, the airfoil is attached to a bucket platform that is of greater stiffness than the airfoil. When the airfoil is forced to follow the displacement of the shank and platform, high thermal stresses occur on the airfoil, particularly in the thin trailing edge region. These high thermal stresses are present during transient engine operation as well as steady state, full speed, full load conditions, and can lead to crack initiation and propagation. These cracks potentially can ultimately lead to catastrophic failure of the component.
U.S. Pat. No. 5,947,687 discloses a gas turbine moving blade (
FIGS. 1-3
) having a groove
16
on the trailing side
18
of the platform of a turbine blade, designed to suppress a high thermal stress at the attachment point of the airfoil trailing edge and platform that occurs during transient operating conditions, i.e., starting and stopping of the turbine. However, the groove has a depth which does not enter a stress line of the platform caused by the load on the airfoil. Since the groove does not enter a stress line, it does not affect the load path through the trailing edge of the airfoil, and the groove is, therefore, not highly stressed. Also, this groove extends along the entire length of the platform, from the concave side
20
of the blade to the convex side
24
, along a circumference of the turbine, parallel to a plane of rotation of the turbine. In this configuration, the groove affects blade natural frequencies, thereby potentially inducing additional mechanical vibratory stress on the blade.
BRIEF SUMMARY OF THE INVENTION
It is therefore seen to be desirable to reduce the likelihood of initiating cracks in the root trailing edge region of the airfoil by reducing the thermal and mechanical stresses that occur due to the mismatch between the airfoil and the shank.
The present invention provides a gas turbine moving blade in which a groove is introduced in the bucket platform, at an angle with respect to a mean camber line of the airfoil, such that the groove begins on the concave side of the platform and exits the platform on the trailing edge side of the bucket shank cover plate. In alternative embodiments, the cross-section of the groove may be circular, elliptical, or square with simple or compound radii, rectangular, or polygonal, in which the groove is defined by two or more planes. This groove has a depth which will enter a stress line of the platform caused by a load encountered by the blade, and will change the load path direction away from the trailing edge.
The location and depth of the groove of the present invention results in a reduced mechanical as well as thermal stress condition in the airfoil root trailing edge and a higher stressed condition in the groove. An increase in the fatigue capability of this region of the component is possible because the groove is located in a region of cooler metal temperatures having greater material fatigue strength. This groove, additionally, provides a decrease in the mechanical stress in the trailing edge by cutting into the load path of the airfoil, thus having an overall greater benefit in the fatigue life of the region.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1
is a perspective view of a prior art turbine blade.
FIG. 2
is a front side view showing an example of a prior artturbine blade.
FIG. 3
is right side view of the example of a prior art turbine blade illustrated in FIG.
2
.
FIG. 4
is a front side view showing a preferred embodiment of a turbine blade according to the present invention.
FIG. 5
is a right side view of the turbine blade illustrated in FIG.
4
.
FIG. 6
is a cross sectional view, taken along line A—A of
FIG. 4
, of the turbine blade of the present invention.
FIG. 7
is a front side view showing the stress line in a prior art turbine blade.
FIG. 8
is a front side view showing the stress line in a preferred embodiment of a turbine blade according to the present invention.
FIG. 9
is an elevation view of another preferred embodiment of the turbine blade of the present invention.
DETAILED DESCRIPTION OF THE INVENTION
In a preferred embodiment of the present invention, as seen in
FIGS. 4-5
, a turbine blade
30
has a blade root portion
34
, a shank portion
36
, a blade platform
38
, and a blade profile portion (or airfoil)
40
. The platform has a trailing edge side
48
, a concave side
50
, a leading edge side
52
, and a convex side
54
, where the sides are labeled according to their position relative to the blade profile portion
40
. A groove
46
is provided in the platform
38
, such that the groove
46
extends from the concave side
50
to the trailing edge side
48
of the platform
38
, where the groove exits the platform.
As seen in
FIG. 6
, the preferred orientation of groove
46
is at an angle of about 90 degrees from the mean camber line
60
at the trailing edge
43
of the airfoil
40
. A prior art turbine blade
28
shown in
FIG. 7
has a stress line
26
encountered by blade
28
, or blade load, that includes stress distribution along the airfoil root trailing edge
18
. As seen in
FIG. 8
, groove
46
has a depth
68
that will enter a stress line
70
(shown after alteration by groove
46
) of turbine blade
30
caused by a load encountered by blade
30
, or blade load. Thus, groove
46
causes a change to the load path direction away from the trailing edge
48
. Consequently, the groove location and depth results in a reduced mechanical as well as thermal stress condition in the airfoil root trailing edge
48
and a higher stressed condition in the groove
46
. An increase in the fatigue capability of this region of the component is possible because the groove
46
is located in a region of cooler metal temperatures having greater material fatigue strength. This groove
46
additionally provides a decrease in the mechanical stress in the trailing edge
48
by cutting into the load path of the airfoil, thus having an overall greater benefit in the fatigue life of the region. Also, the groove
46
is angled, such that the groove
46
begins on the concave side
50
of the platform and exits on the trailing edge side
48
of the bucket shank cover plate
56
. This groove orientation has a significantly smaller effect on blade natural frequencies than a groove that completely extends from the concave side to the convex side of the blade, thereby further reducing the potential for increased mechanical vibratory stress in the airfoil.
In alternative embodiments, the groove
46
may possess any of a number of shapes, such that the cross-section of the groove may be, but is not limited to, circular, elliptical, square, rectangular, or polygonal, in which the groove is defined by two or more planes. In a preferred embodiment of the present invention, the shape of the groove has an elliptical cross-section. In a most-preferred embodiment, as seen in
FIG. 9
, the elliptical groove
46
has a semi-major dimension
62
of 0.237″ and a semi-minor dimension
64
of 0.160″, based on an airfoil
40
height of 5.60″. This embodiment has a preferred radial distance
66
from the groove
46
to the top
39
of the blade platform
38
of 0.085″, and the depth
68
is 1.050″. The depth
68
of the groove
46
is application specific, and controls the distribution of load between the groove and the airfoil trailing edge
48
. Increasing the depth
68
decreases trailing edge stress and increases groove stress, and vice versa.
While the preferred form of the present invention has been described, variations thereof will occur to those skilled in the art within the scope of the present inventive concepts that are delineated by the following claims.
Claims
- 1. A gas turbine blade comprising:a blade platform having a blade trailing edge side, a blade convex side, a blade concave side, and a blade leading edge side; a blade profile portion connected to said blade platform; and a groove formed in said blade trailing edge side of said blade platform, wherein said groove begins on said blade concave side and exits on said blade trailing edge side.
- 2. The groove as claimed in claim 1, said groove being at an angle with respect to a mean camber line of a trailing edge of said blade profile portion.
- 3. The groove as claimed in claim 2, said angle being 90 degrees.
- 4. The groove as claimed in claim 1, said groove having a depth that will enter into a line of stress created by a blade load.
- 5. The gas turbine blade as claimed in claim 1, said groove having a substantially elliptical cross-section.
- 6. The gas turbine blade as claimed in claim 1, said groove having a substantially round cross-section.
- 7. A gas turbine blade comprising:a blade platform having a blade trailing edge side, a blade convex side, a blade concave side, and a blade leading edge side; a blade profile portion connected to the blade platform; and a groove formed in the blade platform, the groove having an elliptical cross-section and extending from the blade concave side to the blade trailing edge side at an angle of 90° with respect to a mean camber line of a trailing edge of the blade profile portion.
- 8. The gas turbine blade as claimed in claim 7, wherein the groove has a depth that will enter into a line of stress created by a blade load.
US Referenced Citations (8)