U.S. patent application number 09/887448 was filed with the patent office on 2002-12-26 for method for repairing cracks in a turbine blade root trailing edge.
Invention is credited to Dolansky, Gregory M., Merry, Brian D., Reinhardt, Gregory E., Surace, Raymond C..
Application Number | 20020194733 09/887448 |
Document ID | / |
Family ID | 25391153 |
Filed Date | 2002-12-26 |
United States Patent
Application |
20020194733 |
Kind Code |
A1 |
Surace, Raymond C. ; et
al. |
December 26, 2002 |
METHOD FOR REPAIRING CRACKS IN A TURBINE BLADE ROOT TRAILING
EDGE
Abstract
The present invention relates to a method for repairing a
turbine blade having a crack in a trailing edge portion of the
blade. The method comprises the steps of cutting back a first
surface of the turbine blade adjacent the blade trailing edge
portion where the crack is located, and cutting back a second
surface of the turbine blade adjacent the blade trailing edge
portion where the crack is located. Each cut back step comprises
cutting back the respective surface by a depth greater than the
length of the crack and less than the trailing edge radius to
remove the crack and form a cut back trailing edge portion. A
compound radius is used to prevent a blunt transition into the
trailing edge that would result in aerodynamic losses and to reduce
the airfoil root stresses. The method also includes applying a
thermal barrier coating to the turbine blade to increase service
life. Prior to applying the coating, the tip portion of the turbine
blade is modified to account for the change in the thermal
characteristics of the turbine blade.
Inventors: |
Surace, Raymond C.;
(Middletown, CT) ; Merry, Brian D.; (Andover,
CT) ; Dolansky, Gregory M.; (Higganum, CT) ;
Reinhardt, Gregory E.; (Glastonbury, CT) |
Correspondence
Address: |
Barry L. Kelmachter
BACHMAN & LaPOINTE, P.C.
Suite 1201
900 Chapel Street
New Haven
CT
06510-2802
US
|
Family ID: |
25391153 |
Appl. No.: |
09/887448 |
Filed: |
June 22, 2001 |
Current U.S.
Class: |
29/889.1 |
Current CPC
Class: |
F01D 5/005 20130101;
Y10T 29/49318 20150115; Y10T 29/49996 20150115; Y10T 29/49746
20150115; Y10T 29/49336 20150115; B23P 6/007 20130101; B23P 6/045
20130101; Y10T 29/49726 20150115 |
Class at
Publication: |
29/889.1 |
International
Class: |
B23P 006/00 |
Claims
What is claimed is:
1. A method for repairing a turbine blade having a crack in a
trailing edge portion of the blade, said method comprising the
steps of: cutting back a first surface of the turbine blade
adjacent the blade trailing edge portion where said crack is
located; cutting back a second surface of the turbine blade
adjacent the blade trailing edge portion where said crack is
located; and each of said cutting back steps comprising cutting
back said respective surface by a depth greater than the length of
said crack and less than the trailing edge radius to remove said
crack and form a cut back trailing edge portion.
2. A method according to claim 1, wherein each of said cutting back
steps comprises cutting back said respective surface from a first
point adjacent a root portion of said trailing edge portion to a
second point at the approximate mid-span of the turbine blade.
3. A method according to claim 2, further comprising blending said
cut back trailing edge portion into a portion of the original
trailing edge portion.
4. A method according to claim 1, blending said cut back trailing
portion into a platform portion of said turbine blade.
5. A method according to claim 4, wherein said blending step
comprises using a compound radius to eliminate any cusp on the
trailing edge of said turbine blade.
6. A method according to claim 4, further comprising blending
remaining edges to a smooth radius to minimize stress
concentrations.
7. A method according to claim 1, further comprising applying a
thermal barrier coating to said turbine blade after said cutting
back steps.
8. A method according to claim 7, further comprising modifying the
tip length of the turbine blade prior to the thermal barrier
coating applying step to account for reduced substrate
temperatures.
9. A method according to claim 8, wherein said modifying step
comprises applying weld material to a tip portion of said turbine
blade prior to said thermal barrier coating applying step and
machining said turbine blade to a predetermined length.
10. A method according to claim 9, wherein said weld material
applying step comprises applying a nickel base alloy welding
material to said tip portion.
11. A method according to claim 7, wherein said thermal barrier
coating step comprises applying a thermally insulating ceramic
coating.
12. A method according to claim 7, wherein said thermal barrier
coating applying step comprises applying a MCrAlY coating where M
is selected from the group consisting of iron, nickel, cobalt and
mixtures of nickel and cobalt.
13. A method according to claim 7, wherein said thermal barrier
coating applying step comprises applying a MCrAlY coating where M
is nickel or cobalt and which contains silicon and hafnium.
14. A method according to claim 7, wherein said thermal barrier
coating applying step comprises applying a thermally insulating
ceramic coating.
15. A method according to claim 14, wherein said thermally
insulating ceramic coating has a cubic pyrochlore structure.
16. A method according to claim 7, wherein said thermal barrier
coating applying step comprises applying a ceramic thermal barrier
coating composed of gadolina and zirconia.
Description
BACKGROUND OF THE INVENTION
[0001] The present invention relates to a method for repairing
cracks in a trailing edge portion of a turbine blade.
[0002] Axial cracks initiating at the root trailing edge cooling
hole occur on turbine blades used in industrial applications. The
cracks are caused by thermal mechanical fatigue. Typically, the
cracks initiate from both the concave and the convex side of the
root trailing edge cooling hole and run axially towards the leading
edge of the blade. Since the turbine blades are otherwise
serviceable, a method for effectively repairing these cracks is
needed.
SUMMARY OF THE INVENTION
[0003] Accordingly, it is an object of the present invention to
provide a method for repairing cracks in a trailing edge portion of
a turbine blade.
[0004] It is a further object of the present invention to provide a
repair method as above which has particular utility in the repair
of cracks initiating at a root trailing edge cooling hole.
[0005] It is yet a further object of the present invention to
provide a method as above which increases the service life of the
repaired turbine blade.
[0006] The foregoing objects are attained by the method of the
present invention.
[0007] In accordance with the present invention, a method for
repairing a turbine blade having a crack in a trailing edge portion
of the turbine blade is provided. The method broadly comprises
cutting back the trailing edge portion of the concave and convex
surfaces adjoining the trailing edge portion to a depth greater
than the length of the crack. Concurrent with the cut back
procedure, the portion of the turbine blade between the platform
and the cut back trailing edge portion is shaped using a compound
radius to eliminate the presence of any cusp on the trailing edge.
Further, those edges remaining after the cut back procedure are
blended to a smooth radius to minimize stress concentration and
aerodynamic losses. The cut back trailing edge portion is also
faired into the original trailing edge profile, preferably at the
approximate mid-span, to minimize aerodynamic impact.
[0008] In accordance with the present invention, a thermal barrier
coat is applied to the repaired turbine blade to increase its
service life. Prior to the application of the thermal barrier
coating, the tip length of the turbine blade is modified to account
for reduced substrate temperature of the repaired turbine
blade.
[0009] Other details of the repair method of the present invention,
as well as other objects and advantages attendant thereto, are set
forth in the following detailed description and the accompanying
drawings wherein like reference numerals depict like elements.
BRIEF DESCRIPTION OF THE DRAWINGS
[0010] FIG. 1 is a perspective view of a root portion of a turbine
blade to be repaired;
[0011] FIG. 2 is a side view of the root trailing edge portion on
the concave airfoil side of the turbine blade;
[0012] FIG. 3 is a side view of the root trailing edge portion on
the convex airfoil side of the turbine blade;
[0013] FIG. 4 is a side vide of the root trailing edge portion of
the turbine blade showing the compound radius curve used to blend
the cut back trailing edge portion to the platform portion of the
turbine blade;
[0014] FIG. 5 is a side view of a turbine blade repaired in
accordance with the present invention;
[0015] FIG. 6 is a perspective view of the turbine blade of FIG. 5;
and
[0016] FIG. 7 is a rear view of the turbine blade of FIG. 5.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT(S)
[0017] Referring now to the drawings, FIG. 1 shows a portion of a
turbine blade 10 that requires repair. As can be seen from the
figure, the trailing edge 12 of the turbine blade 10 is provided
with a plurality of cooling holes 14. In service, one or more
cracks 16 form in the vicinity of the lowermost one 14' of the
cooling holes 14 known as the root trailing edge cooling hole.
Typically, the cracks initiate from both the concave side 18 and
the convex side 20 of the airfoil portion 22 of the turbine blade
10. Each crack 16 extends axially toward the leading edge 24 of the
blade 10. It has been found that the cracks 16, that fall within
acceptable serviceable limits, preferably that extend less than
approximately about 0.05 inches and as determined for a given
blade, may be repaired using the method of the present
invention.
[0018] To repair the cracks 16, a portion 26 of the trailing edge
known as the root trailing edge portion is cut back on both the
concave side 18 and the convex side 20. This cut back is shown in
FIGS. 2 and 3. As can be seen in each of these figures, the
original trailing edge 12 is cut back by a distance or depth L to
form a cut back trailing edge portion 28. In a preferred embodiment
of the present invention, the distance or depth L is greater than
the length of the crack 16. This cut back removes material in the
area where the crack 16 is located and reach fresh material where
there are no cracks or microcracks. The distance or depth L is
preferably less than the radius of the trailing edge 12.
[0019] The cutting back of the root trailing edge portion 26 may be
carried out using any suitable means known in the art. Preferably,
it is carried out by grinding or milling each of the concave side
18 and the convex side 20.
[0020] Concurrent with the cut back of each of the sides 18 and 20,
it is necessary to blend the cut back trailing edge portion 28 into
the platform 30 on the turbine blade. The blending must be carried
out so that there is a smooth transition between the cut back
trailing edge portion 28 and the platform 30. A unique feature of
this invention is the use of a compound radius approach to achieve
this smooth transition. The compound radius provides a large radius
in the high stress location, while rapidly transitioning into the
existing platform profile. The rapid transition eliminates a large
trailing edge 12 blunt area which would increase aerodynamic
losses. As can be seen in FIG. 4, a compound radius having a major
radius R.sub.1 and a minor radius R.sub.2, which compound radius
preferably varies from approximately about 0.375 inches to
approximately about 0.1875 inches, is used in the transition area
between the cut back trailing edge 28 and the platform 30. By using
this compound radius blending approach, the formation of a cusp on
the trailing edge is avoided.
[0021] When the concave and convex sides are cut back, they are not
cut back along the entire span of the airfoil portion 22 of the
turbine blade 10. Preferably, the top half 32 of the original
trailing edge 12 is left alone. It is then necessary to fair the
cut back trailing edge portion 28 into the original trailing edge
profile to minimize aerodynamic impact. Any suitable technique
known in the art which minimizes abrupt changes/discontinuities in
the trailing edge geometry of the turbine blade 10 and which avoids
adverse effects on the flow field, vibrations, and structural
integrity may be used to fair the cut back trailing edge portion 28
into the original trailing edge profile. As shown in FIGS. 5 and 6,
preferably, the cut back trailing edge portion 28 is faired into
the original trailing edge profile at approximately about the 50%
span.
[0022] After the cut back, platform blending, and trailing edge
fairing steps have been completed to a desired depth, any remaining
edges are blended, either by machine or by hand, to a smooth radius
to minimize stress concentration and aerodynamic losses. The edges
34, 36, 38, 40, 42, and/or 44 which typically require the blending
are shown in FIG. 7. In a preferred embodiment of the repair method
of the present invention, the edges 34, 36, 38, 40, 42, and/or 44
are blended to a smooth radius of approximately about 0.005 to
approximately about 0.015 inches.
[0023] If needed, the cooling holes 14 may be refurbished using any
suitable technique known in the art.
[0024] It has been found that the service life of a turbine blade
10 repaired as above can be increased by approximately about 2X or
more by applying a thermal barrier coating to the turbine blade 10.
The thermal barrier coating may comprise any suitable thermal
barrier coating known in the art and may be applied using any
suitable means known in the art. For example, the thermal barrier
coating may be a MCrAlY coating where M is selected from the group
consisting of iron, nickel, cobalt, and mixtures of nickel and
cobalt such as that shown in U.S. Pat. No. 4,321,311, which is
hereby incorporated by reference herein. Alternatively, the thermal
barrier coating may be a MCrAlY type coating where M is nickel or
cobalt and which is improved by the addition of from 0.1 to 7.0% by
weight silicon and 0.1 to 2.0% by weight hafnium such as that shown
in U.S. Pat. No. 4,585,481, which is hereby incorporated by
reference herein. The thermal barrier coating could also be a
thermally insulating ceramic coating having a pyrochlore structure
such as that shown in U.S. Pat. No. 6,117,560, which is hereby
incorporated by reference herein. The thermal barrier coating could
also be a thermally insulating ceramic coating containing gadolinia
and zirconia such as that shown in U.S. Pat. No. 6,177,200, which
is hereby incorporated by reference herein. The thermal barrier
coating may be applied to the turbine blade using any of the
techniques shown in the aforementioned U.S. Patents.
[0025] It has been found desirable to modify the tip portion 46 of
the turbine blade 10 to increase its length prior to applying the
thermally insulating ceramic coating to the turbine blade 10. This
is to account for the reduced substrate temperatures which will be
encountered by the turbine blade 10 as a result of the thermally
insulating ceramic coating. The tip portion 46 is preferably
modified by applying a weld material to the tip portion 46 and
machining the turbine blade 10 to a predetermined length. While the
weld material to be applied to the tip portion 46 may comprise any
suitable welding or brazing material known in the art, it is
preferred to add a nickel based alloy weld material to the tip
portion 46. The added weld material helps avoid any negative
clearance effects caused by the application of the thermally
insulating ceramic coating. With the thermally insulating ceramic
coating, the turbine blade 10 will not thermally expand as
originally designed. The turbine blade 10 will expand less, as it
is cooler. This in turn creates a larger gap at the tip portion 46
in the radial direction, than is desirable from a leakage/sealing
standpoint and from a performance standpoint. Thus, to accommodate
the reduced tip growth, the tip portion 46 is built up with weld
material and then machined to a desired predetermined length.
[0026] One of the principal advantages to adding the thermal
barrier coating to the repaired turbine blade 10 is that it reduces
blade metal temperature gradients. The reduced thermal gradient
combined with thicker trailing edge walls and increased fillet
radius reduce airfoil root stresses and increases blade service
life. In particular the thermal mechanical fatigue life, of the
repaired blade is increased by 2 times or more.
[0027] Further, it is within the scope of this invention, and
understood by those skilled in the art, that the method described
herein may be utilized to repair a variety of blades thus
advantageously providing a repaired blade with enhanced service
life, as compared to that of the original blade.
[0028] It is apparent that there has been provided in accordance
with the present invention a method for repairing trailing edge
cracks in turbine blades which fully satisfies the objects, means
and advantages set forth hereinbefore. While the present invention
has been described in the context of specific embodiments thereof,
other alternatives, modifications, and variations will become
apparent to those skilled in the art having read the foregoing
description. Therefore, it is intended to embrace those
alternatives, modifications, and variations which fall within the
broad scope of the appended claims.
* * * * *