U.S. patent application number 11/606598 was filed with the patent office on 2008-06-05 for rmc-defined tip blowing slots for turbine blades.
This patent application is currently assigned to United Technologies Corporation. Invention is credited to Jason Edward Albert, Jeffrey S. Beattie, Francisco J. Cunha.
Application Number | 20080131285 11/606598 |
Document ID | / |
Family ID | 39106243 |
Filed Date | 2008-06-05 |
United States Patent
Application |
20080131285 |
Kind Code |
A1 |
Albert; Jason Edward ; et
al. |
June 5, 2008 |
RMC-defined tip blowing slots for turbine blades
Abstract
A process for forming an airfoil portion of a turbine engine
component, such as a turbine blade, is described. The process
comprises the steps of placing a ceramic core having a
configuration of a passageway to be formed in the airfoil portion
within a mold, attaching a refractory metal core element to the
ceramic core to stabilize a tip region of the ceramic core, and
casting the airfoil portion.
Inventors: |
Albert; Jason Edward; (West
Hartford, CT) ; Cunha; Francisco J.; (Avon, CT)
; Beattie; Jeffrey S.; (West Hartford, CT) |
Correspondence
Address: |
BACHMAN & LAPOINTE, P.C.
900 CHAPEL STREET, SUITE 1201
NEW HAVEN
CT
06510
US
|
Assignee: |
United Technologies
Corporation
|
Family ID: |
39106243 |
Appl. No.: |
11/606598 |
Filed: |
November 30, 2006 |
Current U.S.
Class: |
416/96R ;
164/159; 164/6; 29/889.71; 416/241B |
Current CPC
Class: |
Y10T 29/49337 20150115;
F05D 2230/211 20130101; F01D 5/147 20130101; F01D 5/187 20130101;
F05D 2300/13 20130101; B22C 9/04 20130101; B22C 9/103 20130101 |
Class at
Publication: |
416/96.R ;
416/241.B; 164/6; 164/159; 29/889.71 |
International
Class: |
F01D 5/18 20060101
F01D005/18; F01D 25/12 20060101 F01D025/12; B22D 15/00 20060101
B22D015/00; B22D 19/00 20060101 B22D019/00; B21D 53/78 20060101
B21D053/78 |
Claims
1. A process for forming an airfoil portion of a turbine engine
component comprising the steps of: positioning a ceramic core
having a configuration of a passageway to be formed in said airfoil
portion; attaching a refractory metal core element to said ceramic
core to stabilize a tip region of said ceramic core; creating a
mold about the ceramic core and refractory metal core element; and
casting said airfoil portion.
2. The process of claim 1, further comprising locating said ceramic
core relative to said mold with said refractory metal core
element.
3. The process of claim 2, wherein said locating step comprises
providing a refractory metal core element having at least one
leg.
4. The process of claim 3, wherein said refractory metal core
element providing step comprises providing a refractory metal core
element having a plurality of legs.
5. The process of claim 3, further comprising removing said ceramic
core so as to form said passageway and subsequently removing said
refractory metal core element and thereby leaving at least one
cooling slot in a tip region of said airfoil portion.
6. The process of claim 5, wherein said removing step comprises
leaving a plurality of cooling slots in said tip region.
7. The process of claim 5, further comprising machining a plurality
of film cooling holes in said airfoil portion in the vicinity of
said passageway formed by said ceramic core.
8. In combination, a ceramic core for forming a passageway in a
cast airfoil portion and means for stabilizing a tip region of said
ceramic core, said stabilizing means comprising a refractory metal
core element.
9. The combination of claim 8, wherein said refractory metal core
element comprises a solid portion and a plurality of legs depending
from said solid portion.
10. The combination of claim 9, wherein each said legs has an
angled portion and a base portion and said base portion of said
legs is joined together by a lower portion.
11. A refractory metal core element comprising a solid portion and
a plurality of spaced apart legs depending from said solid portion,
each of said legs having a first portion adjacent said solid
portion, a base portion, and an angled portion intermediate said
first portion and said base portion so that said base portion is
laterally offset from said solid portion.
12. A refractory metal core element according to claim 11, further
comprising a lower portion connecting each base portion.
13. A turbine engine component having an airfoil portion with a tip
region, a recessed shelf in said tip region, and a plurality of
slots in said recessed shelf through which a cooling fluid flows,
said slots being located along a pressure side of said tip
region.
14. The turbine engine component according to claim 13, further
comprising said slots being oriented primarily radially outwards
and being angled towards said pressure side.
15. The turbine engine component according to claim 13, further
comprising a passageway within said tip region and each of said
slots communicating with said passageway.
16. The turbine engine component according to claim 15, further
comprising a plurality of cooling holes machined in said pressure
side and communicating with said passageway.
17. The turbine engine component according to claim 15, wherein
said passageway comprises a laterally-oriented cavity.
Description
BACKGROUND
[0001] (1) Field of the Invention
[0002] The present invention relates to a process for forming a
turbine engine component, such as a turbine blade, having a
plurality of as-cast blowing slots in a tip region using a
refractory core element.
[0003] (2) Prior Art
[0004] One of the typical failure modes for high pressure turbine
(HPT) rotor airfoils (blades) is tip distress via oxidation and
erosion. It is particularly challenging to design a cooling
configuration for a tip region for a variety of reasons. First, it
is very difficult to determine the external thermal boundary
conditions near the tip due to the highly-three dimensional nature
of the gaspath flow. Also, the tip region of a turbine blade is
typically the thinnest portion of the airfoil, which makes it more
difficult to package the desired cooling features. Furthermore, the
tip region of a turbine blade is typically difficult to accurately
produce with investment casting processes because the internal
ceramic core is thin and weak near the tip. Further, it is
cantilevered relatively far from the core-locating fixture at the
blade root. Considering these points, it is desirable to have
methods to create intricate cooling features near the tip capable
of being targeted at specific regions of high heat load, while also
allowing for greater control during the investment casting
process.
[0005] An existing HPT blade tip cooling design is shown in FIG. 1.
A radially oriented cavity supplies cooling air to a leading edge
impingement cooling scheme as well as a laterally-oriented cavity,
known as a tip flag, that helps cool the tip before exiting the
blade at the trailing edge near the tip. FIG. 1 also shows a
midbody three-pass serpentine cooling arrangement and a trailing
edge double-impingement system.
[0006] The tip of the core in FIG. 1 includes an appendage that
creates a recess blade tip known as a squealer pocket. That
appendage is connected to the leading edge and tip flag core by
means of two cylindrical connections ("print-outs") that form open
holes in the finished casting ("print-out holes"). The core is
fixed at the root of the blade during the casting process. The
squealer pocket core is located laterally during the casting
process, allowing the tip print outs to stabilize the tip region of
the core. In order to prevent core breakage during the casting
process, these tip print-outs should be as large as possible,
especially considering that they are constructed from the brittle
ceramic core material. One of the primary purposes of the squealer
pocket is to allow for a shorter distance that the tip print-outs
must span. However, it is desirable to have the tip print-out holes
be smaller so that they do not flow an excessive amount of cooling
air in the finished part, which results in inefficiency in the
cooling design and, therefore, the turbine performance.
SUMMARY OF THE INVENTION
[0007] In accordance with the present invention, there is provided
a new tip cooling design that utilizes refractory metal core (RMC)
technology in order to create a tip cooling scheme for a turbine
engine component that is capable of more efficient use of cooling
air and a more reliable casting process.
[0008] In accordance with the present invention, a process for
forming an airfoil portion of a turbine engine component is
provided. The process comprises the steps of placing a ceramic core
having a configuration of a passageway to be formed in the airfoil
portion within a mold; attaching a refractory metal core element to
the ceramic core to stabilize a tip region of the ceramic core
during casting; and casting the airfoil portion.
[0009] Further, in accordance with the present invention, there is
in combination, a ceramic core for forming a passageway in a cast
airfoil portion and means for stabilizing a tip region of the
ceramic core. The stabilizing means comprises a refractory metal
core element.
[0010] Still further, in accordance with the present invention,
there is provided a refractory metal core element comprising a
solid portion and a plurality of spaced apart legs depending from
the solid portion. Each of the legs has a first portion adjacent
the solid portion, a base portion, and an angled portion
intermediate the first portion and the base portion so that the
base portion is laterally offset from the solid portion. The base
portions of the legs are preferably joined together by a lower
portion.
[0011] Still further, in accordance with the present invention,
there is provided a turbine engine component having an airfoil
portion with a tip region, a shelf portion in said tip region, and
a plurality of as-cast slots in the shelf portion through which a
cooling fluid flows. The slots are located along a pressure side of
the tip region.
[0012] Other details of the RMC-defined tip blowing slots for
turbine blade 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
[0013] FIG. 1 is a schematic representation of a cooling design
used in a prior art turbine blade;
[0014] FIG. 2 is an external view of a tip region of a casting;
[0015] FIG. 3 is a schematic representation of a tip region of a
cast airfoil portion of a turbine blade;
[0016] FIG. 4 is a view of a refractory metal core element from the
pressure side; and
[0017] FIG. 5 is a view of the refractory metal core element from
the trailing edge.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT(S)
[0018] As noted before, a new tip cooling design for a turbine
blade is proposed here that utilizes refractory metal core
technology in order to help create a tip cooling scheme that is
capable of more efficient use of cooling air and a more reliable
casting process.
[0019] Referring now to FIG. 2, a relatively thin, approximately
0.015'', refractory metal core element 10 is used to stabilize a
tip region 12 of a ceramic core 14 during the casting process. The
ceramic core 14 is positioned within a mold 80, only a portion of
which has been shown. The ceramic core 14 may have the
configuration of a laterally oriented passageway 15 to be formed in
the airfoil tip region 34. The refractory metal core element 10 is
printed out of the airfoil tip region 34 during casting and is
located laterally of the ceramic core 14. Preferably, the
refractory metal core element 10 is positioned adjacent a side of
the mold which forms the pressure side 40 of the airfoil portion
42. The refractory metal core element 10 is a metal piece which is
much more rugged than typically brittle core print-outs. Thus,
there is no manufacturing requirement for relatively large core
print-out. A core print-out hole (not shown) may still be included
if it is required for cooling purposes. In the present case, the
core print-out hole can be made smaller than it previously could
because it is not required to have as high of a strength. This
configuration also allows for multiple ceramic core features to be
stabilized by the same refractory metal core element. Furthermore,
because this new tip design provides more stability and strength
for the ceramic core 14 near the tip, the size of the trailing edge
print-out of the tip flag cavity can be reduced, enabling lower
cooling air flow out the tip flag exit.
[0020] The refractory metal core element 10 may be formed from any
suitable refractory material known in the art such as molybdenum or
a molybdenum alloy. The refractory metal core element 10, as shown
in FIGS. 2, 4 and 5 may have a solid portion 46 and a plurality of
spaced apart legs 48 depending downwardly from the solid portion
46. Each leg 48 preferably has a first leg portion 50, a base
portion 52, and an angled portion 54 between the first portion 50
and the base portion 52. The base portion 52 of the legs may be
joined together by a lower portion 53. The refractory metal core
element 10 may be attached to the ceramic core 14 using any
suitable means known in the art such as an adhesive or a mechanical
fit connection. Because the refractory metal core element 10 and
the ceramic core 14 are attached, inside the casting, the
refractory metal core element can be used to control the location
of both the refractory metal core and the ceramic core, relative to
the external mold. In an alternative embodiment of the refractory
metal core element 10, the angled portion 54 may be omitted. Still
further, the legs 48 can be arranged in any way that makes sense
for the cooling design. Furthermore, the legs 48 only need to be
connected at one end (inside or outside the casting), whichever
makes sense for the cooling design and the casting process.
[0021] As shown in FIG. 3, the refractory metal core element 10 is
printed out in such a way as to produce a row of aligned open slots
30 in the finished casting, along the pressure side edge 32 of the
tip 34. Cooling air may be ejected from the slots 30 in whichever
direction the slots 30 are oriented. As shown in FIG. 3, the slots
30 may be oriented primarily radially outwards towards an outer
circumference of the gaspath. The slots 30 may also be slightly
angled towards the pressure side 40 of the turbine blade airfoil
portion 42. The slots 30 may be purely radial or leaned in any
combination of directions--forward/aft and/or towards
pressure/suction side. The slots 30 may be in fluid communication
with the passageway 15. As shown in FIG. 3, the slots 30 may be
located in a recessed shelf 36 in the tip 34. The recessed shelf 36
may be a cast feature, or it may be machined into the finished
casting in a later process.
[0022] When the cooling air exits the RMC defined tip slots 30, the
cooling air immediately flows into a tip gap between the blade tip
34 and the blade outer air seal (BOAS)(not shown) due to the strong
pressure gradient towards the suction side 60 of the airfoil
portion 42. Injecting the cooling air into the tip gap
significantly reduces the gaspath temperature in the tip gap
downstream of the slots 30, resulting in lower heat load to the tip
region of the blade. This is a similar effect to film cooling on
the body of an airfoil. Conventional tip print-out holes provide
some film cooling benefit on the tip surface, but they are
significantly less efficient than this new design because the
conventional tip print-out holes are so large that they can only be
located at one or two locations along the mid-thickness of the
tip.
[0023] Another cooling benefit of the RMC-defined tip slots 30 is
the substantial convective cooling of the pressure side region of
the tip 34 due to the high-velocity cooling air flowing through the
tip slots 30. This convective cooling is very effective at
preventing oxidation and erosion along the pressure side edge 32 of
the tip 34, which is a common location of tip distress. As a result
of this increased convective cooling along the pressure side edge
32 of the tip 34, it is feasible to use fewer film cooling holes on
the pressure side edge of the airfoil near the tip. In a prior art
design, two rows of shaped cooling holes are provided along the
pressure side near the tip. The purpose of these holes is to cool
the tip region via film cooling and convective cooling. FIG. 3
shows a tip cooling design in accordance with the present invention
which has only a single row of shaped cooling holes 70. The
reduction of two rows of pressure side film cooling to one row is a
benefit of the present invention, but it is not a necessary aspect
of it.
[0024] The flexibility of the convective and film cooling aspects
of the RMC-defined tip slots lends itself well to the challenge of
designing a tip cooling configuration when the external boundary
conditions are difficult to determine. Furthermore, the inherent
strength of the refractory metal core element 10 during the casting
process allows for increased design flexibility in the tip region.
As a result, this new tip cooling configuration allows for more
efficient use of cooling air and more predictable casting yields,
resulting in a more cost-effective product.
[0025] Another advantage of this tip cooling configuration is that
it is complimentary to tip blowing technology for aerodynamic
performance benefits. Tip blowing utilizes a row of cooling air
jets or holes 30 along the pressure side edge 32 of the blade tip
34, which act to improve aerodynamic efficiency by reducing endwall
losses associated with gaspath leakage across the tip gap. The
cooling holes 70 may be machined in the pressure side edge 32 after
the blade and its airfoil portion have been cast. The cooling holes
70 may be machined using any suitable technique known in the art.
The cooling holes 70 are preferably in fluid communication with the
passageway 15. The RMC-defined cooling slots 30 may be situated
along the recessed shelf 36 along the pressure side of the tip 34.
The recessed shelf 36 will prevent the slots 30 from being
unexpectedly closed during engine operation when the blade tip 34
rubs against the outer circumference of the gaspath. The recessed
shelf 36 also allows for easier masking when applying abradable
coating to the tip surface.
[0026] The tip portion 34 of the airfoil portion 42 of the turbine
engine blade is a cast structure and is formed at the same time as
the remainder of the cast portions of the turbine engine blade. For
simplicity sake, only a portion of the mold 80 forming the tip
region 34 of the airfoil portion 42 is illustrated in the drawings.
It should be recognized that the mold 80 has a portion which is in
the shape of the pressure side of the airfoil.
[0027] The tip portion 34 may be formed by placing the ceramic core
14 into a mold 80. After the ceramic core 14, as well as any other
needed ceramic or silica cores, has been positioned, the refractory
metal core element 10 may be attached to the ceramic core 14 using
any suitable means known in the art, such as an adhesive or pins.
The mold 80 is created after the ceramic core 14 and the RMC 10 are
assembled. This is preferably done by first assembling the ceramic
core 14 and RMC 10, then injecting wax around the cores 10 and 14
using a wax die, so that the external surface of the wax is the
same geometry as the external surface of finished casting. Then, a
ceramic shell is applied to the external surface of the wax
pattern. Then, the wax is melted out, leaving the ceramic core 14,
RMC 10 and ceramic shell (not shown). As previously mentioned, the
refractory metal core element 10 serves to stabilize the tip region
of the ceramic core 14. Thereafter the blade with the airfoil
portion may be cast using any suitable technique known in the art.
After casting has been completed, the ceramic core 14 may be
removed using any suitable technique known in the art to leave the
passageway 15. Similarly, the refractory metal core element 10 is
removed, thus leaving the slots 30. The RMC 10 may be leached out
of the casting using any suitable chemical bath known in the art,
very similar to how the ceramic cores are leached. Thereafter, a
plurality of cooling holes 70 may be machined into the tip region
of the airfoil portion 42.
[0028] While the present invention has been described in the
context of turbine blades, it should be apparent to those skilled
in the art that the process of the present invention, as well as
the refractory metal core element of the present invention, may be
used to form tip blowing slots in other turbine engine
components.
[0029] It is apparent that there has been provided in accordance
with the present invention RMC-defined tip blowing slots for a
turbine blade 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 unforeseeable alternatives, modifications and variations may
become apparent to those skilled in the art having read the
foregoing description. Accordingly, it is intended to embrace those
alternatives, modifications, and variations, as fall within the
broad scope of the appended claims.
* * * * *