Information
-
Patent Grant
-
6612811
-
Patent Number
6,612,811
-
Date Filed
Wednesday, December 12, 200122 years ago
-
Date Issued
Tuesday, September 2, 200321 years ago
-
Inventors
-
Original Assignees
-
Examiners
- Look; Edward K.
- White; Dwayne J.
Agents
- Andes; William Scott
- Davidson; James P.
-
CPC
-
US Classifications
Field of Search
US
- 415 115
- 415 116
- 416 96 R
- 416 96 A
- 416 97 R
- 164 369
- 164 132
- 029 8892
- 029 8897
- 029 88972
- 029 88971
-
International Classifications
-
Abstract
An airfoil for a turbine nozzle assembly of a gas turbine engine includes an outer side wall, an inner side wall, a leading edge extending from the outer side wall to the inner side wall, a trailing edge extending from the outer side wall to the inner side wall, a concave surface extending from the leading edge to the trailing edge on a pressure side of the airfoil, a convex surface extending from the leading edge to the trailing edge on a suction side of the airfoil, an outer cooling slot, an inner cooling slot, and at least one middle cooling slot formed in the concave side of the airfoil adjacent the trailing edge. Each of the cooling slots further includes a recessed wall, an inner slot side wall, an outer slot side wall, an inner corner fillet located between the inner slot side wall and the recessed wall, and an outer corner fillet located between the outer slot side wall and the recessed wall, wherein one of the inner and outer corner fillets for at least one of the inner and outer cooling slots forms a variable contour from an opening in the concave surface to an exit plane of the trailing edge cooling slots.
Description
BACKGROUND OF THE INVENTION
The present invention relates generally to a turbine nozzle for a gas turbine engine and, in particular, to an airfoil utilized therein having at least one of an inner cooling slot and an outer cooling slot at the trailing edge thereof configured to have a variable fillet between a recessed wall and a side wall so as to reduce stress on the airfoil.
It will be appreciated that a nozzle segment for the high pressure turbine of a gas turbine engine typically includes a pair of hollow airfoils with integral inner and outer flowpath bands. These pieces are cast separately, partially machined, brazed together, and subsequently finish machined to form the nozzle segment. The hollow airfoil is fed internally with cooling air which then flows through trailing edge slots that exit the aft cavity of the airfoil and discharges through openings in the trailing edge of the airfoil. This cooling air then performs convection cooling as it passes along the trailing edge slot within the airfoil. When such air discharges to the flowpath through the openings in the airfoil trailing edge, it provides film cooling for the airfoil trailing edge.
Turbine airfoils with trailing edge cooling slots inherently have a step between the slot and the rib between the slots. It has been found that the step in the cooling slot closest to the nozzle bands at the inner and outer airfoil/flowpath intersection causes a large stress concentration with high thermal stresses present, which can then result in trailing edge axial cracks. The cracks ultimately propagate through the airfoil section and lead to premature failure of the turbine nozzles. The cooling slot itself cannot be removed since overheating of the trailing edge of the airfoil would result.
Moreover, the step is difficult to grind smooth because of its proximity to the airfoil/band junction.
It will be understood that the hollow airfoil cavities and trailing edge cooling slots are formed during a casting process by ceramic core which is produced separately and combined with a wax pattern prior to casting. On previous designs, corner fillets for the trailing edge slot are created by the ceramic core and minimized in order to reduce slot blockage and maintain cooling flow area. During manufacturing, however, the ceramic core is subjected to auto-finishing to remove unwanted core material around the core die splitline. It has been found that this process often removes some, if not all, of the external corner fillet on the core and results in a sharp internal corner in the finished casting. This corner acts as a stress concentration and can initiate cracking of the airfoil trailing edge.
It will be recognized that an attempt to address a similar problem for a turbine blade in a gas turbine engine is disclosed in U.S. Pat. No. 6,062,817, entitled “Apparatus and Methods For Cooling Slot Step Elimination,” which is also owned by the assignee of the present invention. A turbine blade is disclosed therein where at least a portion of a step between an airfoil trailing edge slot and a platform is eliminated. An airfoil core utilized to cast the turbine blade includes a tab for forming a continuous and smooth contour from a first trailing edge slot recessed wall to a juncture of the airfoil. In this way, stress concentration is reduced, thereby improving the longevity and performance of the turbine blade.
Thus, in light of the foregoing, it would be desirable for an improved airfoil design to be developed for use with a turbine nozzle which reduces stress concentrations at the steps of the cooling slots located adjacent the inner and outer nozzle bands without adversely affecting the cooling flow from such slots. It would also be desirable to modify the core utilized so as to eliminate the opportunity for additional stress concentrations created by the auto-finishing manufacturing process.
BRIEF SUMMARY OF THE INVENTION
In a first exemplary embodiment of the invention, an airfoil for a turbine nozzle assembly of a gas turbine engine is disclosed as including an outer side wall, an inner side wall, a leading edge extending from the outer side wall to the inner side wall, a trailing edge extending from the outer side wall to the inner side wall, a concave surface extending from the leading edge to the trailing edge on a pressure side of the airfoil, a convex surface extending from the leading edge to the trailing edge on a suction side of the airfoil, an outer cooling slot, an inner cooling slot, and at least one middle cooling slot formed in the concave side of the airfoil adjacent the trailing edge. Each of the cooling slots also includes a recessed wall, an inner slot side wall, an outer slot side wall, an inner corner fillet located between the inner slot side wall and the recessed wall, and an outer corner fillet located between the outer slot side wall and the recessed wall, wherein one of the inner and outer corner fillets of at least one of the inner and outer cooling slots forms a variable contour from an opening in the concave surface to an exit plane of the trailing edge cooling slots. More specifically, the corner fillet forming the variable contour is radiused in a first plane substantially perpendicular to the slot exit plane from the opening to the exit plane. The airfoil also includes a junction between the corner fillet forming the variable contour and an end portion of the airfoil, wherein the junction is radiused in a second plane substantially perpendicular to the slot exit plane from the opening to the exit plane.
In a second exemplary embodiment of the invention, an airfoil core for a turbine airfoil is disclosed as including a wedge channel for forming a hollow portion of an airfoil and a plurality of fingers extending from the wedge channel, wherein at least one of the fingers located at an end is configured to have a distal portion with a predetermined radius from a first side wall to a second side wall. The distal portion of the finger is radiused in a first plane substantially perpendicular to an axis through the finger and radiused in a second plane substantially parallel to the axis through the finger.
In a third exemplary embodiment of the invention, a method of fabricating an airfoil of a turbine nozzle is disclosed as including the steps of inserting a mold within a die and injecting a slurry into the die. An airfoil is formed that includes an outer side wall, an inner side wall, a leading edge extending from the outer side wall to the inner side wall, a trailing edge extending from the outer side wall to the inner side wall, a concave surface extending from the leading edge to the trailing edge on a pressure side of the airfoil, a convex surface extending from the leading edge to the trailing edge on a suction side of the airfoil, and a plurality of cooling slots formed in the concave side of the airfoil adjacent the trailing edge, each of the cooling slots further including a recessed wall and a pair of slot side walls, and a variable contour for a corner fillet between the recessed wall and one of the slot side walls of a cooling slot adjacent at least one of the inner and outer side walls of the airfoil from an opening in the concave surface to an exit plane of the trailing edge cooling slots. In this way, the corner fillet is formed with a radius in a first plane substantially perpendicular to the slot exit plane that gradually increases from a minimum radius at the opening to a maximum radius at the slot exit plane. The method also includes the step of forming a junction between the corner fillet and an end portion of the airfoil, wherein the junction is radiused in a second plane substantially perpendicular to the slot exit plane from the opening to the exit plane.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1
is a cross-sectional view of a gas turbine engine including a turbine nozzle in accordance with the present invention;
FIG. 2
is an enlarged, perspective view of a segment of the turbine nozzle depicted in
FIG. 1
;
FIG. 3
is an enlarged, partial perspective view of an airfoil and the inner band of the turbine nozzle depicted in
FIG. 2
;
FIG. 4
is a partial sectional view of the airfoil depicted in
FIG. 3
taken along line
4
—
4
;
FIG. 5
is a partial plan view of the airfoil depicted in
FIG. 3
taken along line
5
—
5
;
FIG. 6
is a partial sectional view of the airfoil depicted in
FIG. 3
taken along line
6
—
6
;
FIG. 7
is an enlarged, partial top perspective view of the airfoil depicted in
FIGS. 2-6
including a core portion defining the trailing edge cooling slots in the airfoil; and,
FIG. 8
is a bottom perspective view of the core utilized to define the hollow inner portion and the trailing edge cooling slots of the airfoil.
DETAILED DESCRIPTION OF THE INVENTION
Referring now to the drawings in detail, wherein identical numerals indicate the same elements throughout the figures,
FIG. 1
depicts an exemplary turbofan gas turbine engine
10
having in serial flow communication a conventional fan
12
, a high pressure compressor
14
, and a combustor
16
. Combustor
16
conventionally generates combustion gases that are discharged therefrom through a high pressure turbine nozzle assembly
18
, from which the combustion gases are channeled to a conventional high pressure turbine
20
and, in turn, to a conventional low pressure turbine
22
. High pressure turbine
20
drives high pressure compressor
14
through a suitable shaft
24
, while low pressure turbine
22
drives fan
12
through another suitable shaft
26
, all disposed coaxially about a longitudinal or axial centerline axis
28
.
Referring now to
FIG. 2
, it will be understood that turbine nozzle
18
preferably includes a plurality of circumferentially adjoining nozzle segments
30
to collectively form a complete 360° assembly. Each nozzle segment
30
preferably has two or more circumferentially spaced airfoils
32
which are connected to an arcuate radially outer band
34
and an arcuate radially inner band
36
. More specifically, each airfoil
32
includes an outer side wall
38
whose surface lies adjacent to outer band
34
, an inner side wall
40
whose surface lies adjacent to inner band
36
, a leading edge
42
extending from outer side wall
38
to inner side wall
40
, a trailing edge
44
extending from outer side wall
38
to inner side wall
40
, a concave surface
46
extending from leading edge
42
to trailing edge
44
on a pressure side of airfoil
32
, and a convex surface
48
extending from leading edge
42
to trailing edge
44
on a suction side of airfoil
32
.
As seen in
FIG. 2
, airfoils
32
further include an outer cooling slot
50
located adjacent outer band
34
, an inner cooling slot
52
located adjacent inner band
36
, and at least one middle cooling slot
54
located between outer and inner cooling slots
50
and
52
, respectively. It will be appreciated from
FIGS. 3-6
that each of cooling slots
50
,
52
and
54
is formed by a recessed wall
56
, an inner slot side wall
58
, an outer slot side wall
60
, an inner corner fillet
62
located between inner slot side wall
58
and recessed wall
56
, and an outer corner fillet
64
located between outer slot side wall
60
and recessed wall
56
. The inner and outer slot walls
58
and
60
are generally provided by adjacent ribs
61
interposed between each cooling slot, but it will be seen that a rib
63
is used to provide outer slot side wall
60
for inner cooling slot
52
and an inner portion
78
of airfoil
32
(discussed in greater detail hereinafter) provides inner slot side wall
58
thereof.
In accordance with the present invention, it is preferred that at least one of inner corner fillet
62
for inner cooling slot
52
and outer corner fillet
64
for outer cooling slot
50
form a variable contour (as designated by surface
66
in
FIG. 3
) from an opening
68
in concave surface
46
(known in the art as the breakout) to an exit plane
70
which extends substantially perpendicular to cooling slots
50
,
52
and
54
. It will be seen that a coordinate system defined by an x axis
71
, a y axis
73
and a z axis.
75
is depicted in
FIG. 3
which will be utilized to define various planes discussed herein. As such, exit plane
70
is defined as the extending in the y-z plane thereof.
Although depicted and described herein with respect to inner corner fillet
62
for inner cooling slot
52
, the present invention can be, and preferably is, applied in mirror image to outer corner fillet
64
for outer cooling slot
50
. As evidenced by contour lines
72
in
FIG. 3
, surface
66
(which may also be considered inner slot side wall
58
for inner cooling slot
52
) is radiused in a first plane
74
(defined as extending in the x-z plane) which extends substantially perpendicular to slot exit plane
70
from opening
68
to slot exit plane
70
. It will be appreciated from the curvature of such contour lines
72
that the radius of inner corner fillet
62
forming the variable contour gradually increases from a minimum radius R
min
at opening
68
to a maximum radius R
max
at slot exit plane
70
. This is done in order to maintain the slot area, footprint and cooling characteristics for inner cooling slot
52
.
Further, airfoil
32
includes a junction
76
between inner corner fillet
62
and an inner portion
78
of concave surface
46
, wherein junction
76
is radiused in a second plane
80
(defined as extending in the x-y plane) which extends substantially perpendicular to slot exit plane
70
(and first plane
74
) from opening
68
to slot exit plane
72
. As seen in
FIG. 6
, an angle θ between inner corner fillet
62
and inner portion
78
of airfoil
32
is established at junction
76
, where such angel θ gradually decreases from a maximum angle θ
max
at opening
68
to a minimum angle θ
min
at slot exit plane
72
. It is preferred that maximum angle θ
max
be approximately 65°-85° and minimum angle θ
min
be approximately 0°-10°. It will be seen that angle θ is approximately 45° at the approximate mid-point between opening
68
and slot exit plane
70
shown in FIG.
6
.
In order for inner corner fillet
62
to establish the variable contour of surface
66
, it will be understood that inner slot side wall
58
and recessed wall
56
of inner cooling slot
52
preferably form a continuous curve having a predetermined radius from opening
68
in concave surface
46
to slot exit plane
70
(best seen in FIG.
6
). Similarly, in the case of outer cooling slot
50
, outer slot side wall
60
and recessed wall
56
will preferably form a continuous curve having a predetermined radius from opening
68
in concave surface
46
to slot exit plane
70
.
It will be understood that an airfoil core
100
is utilized to form the interior hollow portions and trailing edge cooling slots
50
,
52
and
54
of airfoil
32
. As seen in
FIG. 8
, airfoil core
100
includes a wedge channel
104
, an outer finger
105
, a plurality of middle fingers
106
, and an inner finger
108
extending from wedge channel
104
. It will be noted that inner finger
108
is utilized to form inner cooling slot
52
of airfoil
32
, outer finger
105
forms outer cooling slot
50
, and middle fingers
106
form middle cooling slots
54
. More specifically, inner finger
108
is configured to have a stem portion
109
connected to wedge channel
104
and a distal portion
110
which has a predetermined radius from a first side wall
112
to a second side wall
114
when viewed in section (see FIGS.
6
-
8
). Contrary to the substantially rectangular distal portions
111
of middle fingers
106
, a continuous curve is established by recessed wall
56
and inner slot side wall
58
of inner cooling slot
52
as described hereinabove. Likewise, a continuous curve is established by recessed wall
56
and outer slot side wall
60
for outer cooling slot
50
in airfoil
32
since distal portion
115
of outer finger
105
preferably has a predetermined radius from a first side wall
117
to a second side wall
119
(see FIG.
8
).
Accordingly, distal portion
110
of inner finger
108
is radiused in a first plane
116
(corresponding to first plane
74
) substantially perpendicular to an axis
118
through inner finger
108
, as well as a second plane
120
(corresponding to second plane
80
) substantially parallel to axis
118
. Although airfoil core
100
is discussed with respect to inner finger
108
, it will be appreciated that a mirror image thereof is preferably utilized for outer finger
105
to form the preferred configuration of outer cooling slot
50
in airfoil
32
.
As noted hereinabove, the nature of the forming process for airfoil core
100
results in “flash,” where ceramic material escapes between two mating pieces of the die. Airfoil core
100
is then preferably finished using a small computer controlled milling machine to remove the flash. As demonstrated by dashed line
122
in
FIG. 6
, this finishing process can also remove a portion of the radius for finger side walls that eventually form inner and outer corner fillets
62
and
64
, which has created sharp corners in previous designs. By providing fillets of variable contour in inner slot side wall
58
of inner cooling slot
52
and outer slot side wall
60
of outer cooling slot
50
in the present invention, the radius for inner corner fillet
62
and outer corner fillet
64
, respectively, for such cooling slots
52
and
50
are better maintained since such corner fillets are present outside a nominal casting geometry of airfoil
32
.
In accordance with a method of fabricating airfoil
32
of turbine nozzle
18
, it will be understood that airfoil core
100
is held within a die so that a wax encapsulates it. A final wax pattern is produced which is a replica of the metal casting for airfoil
32
, with airfoil core
100
taking the place of cavities formed in the finished part. It will be appreciated that the wax pattern is dipped in a ceramic solution and dried a number of times to build up layers which form a strong shell mold. The mold is then heated to melt out the wax and cure the ceramic so that airfoil core
100
remains within the shell to form the cavities of airfoil
32
when the mold is filled with molten metal. A molten alloy is poured into the mold, taking up the form left by the wax, with airfoil core
100
preventing the metal from entering areas that are to be cavities in the finished casting and creating the internal features. Finally, the ceramic shell is broken off the casting and the internal ceramic core
100
is leached out using a dissolving solution. The final casting of airfoil
32
thus has the external form of the wax pattern and the internal features of airfoil core
100
, which preferably includes inner corner fillet
62
of inner cooling slot
52
and outer corner fillet
64
of outer cooling slot
50
as described above.
Having shown and described the preferred embodiment of the present invention, further adaptations of the airfoil
32
for a turbine nozzle
18
, airfoil core
100
, and the method for making such airfoil can be accomplished by appropriate modifications by one of ordinary skill in the art without departing from the scope of the invention. In particular, it will be understood that the concepts described and claimed herein could be utilized in a turbine blade and still be compatible with the present invention.
Claims
- 1. An airfoil for a turbine nozzle assembly of a gas turbine engine, comprising:(a) an outer side wall; (b) an inner side wall; (c) a leading edge extending from said outer side wall to said inner side wall; (d) a trailing edge extending from said outer side wall to said inner side wall; (e) a concave surface extending from said leading edge to said trailing edge on a pressure side of said airfoil; (f) a convex surface extending from said leading edge to said trailing edge on a suction side of said airfoil; (g) an outer cooling slot, an inner cooling slot, and at least one middle cooling slot formed in said concave side of said airfoil adjacent said trailing edge, each of said cooling slots further including: (1) a recessed wall; (2) an inner slot side wall; (3) an outer slot side wall; (4) an inner corner fillet located between said inner slot side wall and said recessed wall; and, (5) an outer corner fillet located between said outer slot side wall and said recessed wall; wherein one of said inner and outer corner fillets for at least one of said inner and outer cooling slots forms a variable contour from an opening in said concave surface to an exit plane of said trailing edge cooling slots.
- 2. The turbine nozzle of claim 1, wherein said corner fillet forming a variable contour is radiused in a first plane substantially perpendicular to said slot exit plane from said opening to said exit plane.
- 3. The turbine nozzle of claim 2, wherein said radius of said corner fillet forming a variable contour gradually increases from a minimum radius at said opening to a maximum radius at said exit plane.
- 4. The turbine nozzle of claim 1, said airfoil including a junction between said corner fillet forming a variable contour and an end portion of said airfoil, wherein said junction is radiused in a second plane substantially perpendicular to said slot exit plane from said opening to said exit plane.
- 5. The turbine nozzle of claim 4, wherein an angle between said corner fillet and said end portion of said airfoil at said junction gradually decreases from a maximum angle at said opening to a minimum angle at said exit plane.
- 6. The turbine nozzle of claim 5, wherein said maximum angle is approximately 65°-85°.
- 7. The turbine nozzle of claim 5, wherein said minimum angle is approximately 0°-10°.
- 8. The turbine nozzle of claim 1, wherein said corner fillet forming a variable contour is said outer corner fillet in said outer cooling slot.
- 9. The turbine nozzle of claim 1, wherein said corner fillet forming a variable contour is said inner corner fillet in said inner cooling slot.
- 10. The turbine nozzle of claim 8, wherein said outer side wall and said recessed wall of said outer cooling slot form a continuous curve having a predetermined radius from an opening in said concave surface to said slot exit plane.
- 11. The turbine nozzle of claim 9, wherein said inner side wall and said recessed wall of said inner cooling slot form a continuous curve having a predetermined radius from an opening in said concave surface to said slot exit plane.
- 12. An airfoil core for a turbine airfoil, comprising:(a) a wedge channel; and (b) a plurality of fingers extending from said wedge channel, wherein at least one of said fingers located at an end is configured to have a distal portion with a predetermined radius from a first side wall to a second side wall.
- 13. The airfoil core of claim 12, wherein said distal portion of said end finger is radiused in a first plane substantially perpendicular to an axis through said finger.
- 14. The airfoil core of claim 12, wherein said distal portion of said end finger is radiused in a second plane substantially parallel to an axis through said end finger.
- 15. The airfoil core of claim 12, wherein said end finger is located at an outer end of said airfoil core.
- 16. The airfoil core of claim 12, wherein said end finger is located at an inner end of said airfoil core.
- 17. The airfoil core of claim 12, said radius between said end finger first and second walls being maintained after auto-finishing so that any sharp corner for a cooling slot formed therefrom is outside a nominal casting geometry of said turbine airfoil.
- 18. A method of fabricating an airfoil of a turbine nozzle, comprising the steps of:(a) inserting a mold within a die; (b) injecting a slurry into the die to form an airfoil that includes an outer side wall, an inner side wall, a leading edge extending from said outer side wall to said inner side wall, a trailing edge extending from said outer side wall to said inner side wall, a concave surface extending from said leading edge to said trailing edge on a pressure side of said airfoil, a convex surface extending from said leading edge to said trailing edge on a suction side of said airfoil, and a plurality of cooling slots formed in said concave side of said airfoil adjacent said trailing edge, each of said cooling slots further including a recessed wall and a pair of slot side walls, and a variable contour for a corner fillet between said recessed wall and one of said slot side walls of a cooling slot adjacent at least one of said inner and outer side walls from an opening in said concave surface to an exit plane of said trailing edge cooling slots.
- 19. The method of claim 18, wherein said corner fillet is formed with a radius in a first plane substantially perpendicular to said slot exit plane that gradually increases from a minimum radius at said opening to a maximum radius at said slot exit plane.
- 20. The method of claim 18, further comprising the step of forming a junction between said corner fillet and an end portion of said airfoil, wherein said junction is radiused in a second plane substantially perpendicular to said slot exit plane from said opening to said exit plane.
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