U.S. patent application number 13/222490 was filed with the patent office on 2013-02-28 for airfoil with nonlinear cooling passage.
The applicant listed for this patent is William Abdel-Messeh, Justin D. Piggush. Invention is credited to William Abdel-Messeh, Justin D. Piggush.
Application Number | 20130052037 13/222490 |
Document ID | / |
Family ID | 46888909 |
Filed Date | 2013-02-28 |
United States Patent
Application |
20130052037 |
Kind Code |
A1 |
Abdel-Messeh; William ; et
al. |
February 28, 2013 |
AIRFOIL WITH NONLINEAR COOLING PASSAGE
Abstract
An example method of manufacturing an airfoil includes providing
a ceramic core corresponding to an interior cooling channel. A
refractory metal core is provided that corresponds to a cooling
passage. The cores are arranged in a mold. An airfoil structure is
cast about the cores to provide a turbine engine airfoil. The
turbine engine airfoil includes a wall providing the interior
cooling channel and an exterior airfoil surface. The cooling
passage is provided in the wall and fluidly connects the interior
cooling channel to the exterior airfoil surface. The cooling
passage includes multiple inlets and multiple outlets respectively
adjoining the interior cooling channel and the exterior airfoil
surface. At least one of a first inlet and outlet has a different
structural flow characteristic than at least one of a second inlet
and outlet.
Inventors: |
Abdel-Messeh; William;
(Middletown, CT) ; Piggush; Justin D.; (LaCrosse,
WI) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Abdel-Messeh; William
Piggush; Justin D. |
Middletown
LaCrosse |
CT
WI |
US
US |
|
|
Family ID: |
46888909 |
Appl. No.: |
13/222490 |
Filed: |
August 31, 2011 |
Current U.S.
Class: |
416/97R ;
164/137; 164/6 |
Current CPC
Class: |
F05D 2250/50 20130101;
F05D 2260/221 20130101; F05D 2250/70 20130101; F01D 5/186 20130101;
F01D 5/187 20130101; B22C 9/103 20130101; F05D 2260/204
20130101 |
Class at
Publication: |
416/97.R ;
164/137; 164/6 |
International
Class: |
F01D 5/18 20060101
F01D005/18; B22C 9/10 20060101 B22C009/10; B22D 25/02 20060101
B22D025/02 |
Claims
1. A turbine engine airfoil comprising: an airfoil structure having
a wall providing an interior cooling channel and an exterior
airfoil surface, a cooling passage provided in the wall fluidly
connecting the interior cooling channel to the exterior airfoil
surface, the cooling passage including at least one inlet and
multiple outlets respectively adjoining the interior cooling
channel and the exterior airfoil surface, at least one of a first
inlet and outlet having a different structural flow characteristic
than at least one of a second inlet and outlet.
2. The turbine engine airfoil according to claim 1, wherein the
structural flow characteristic includes at least one of length,
path and shape.
3. The turbine engine airfoil according to claim 2, wherein the
shape includes a cross-sectional area, the first outlet having a
different cross-sectional area than the second outlet.
4. The turbine engine airfoil according to claim 1, wherein the
cooling passage extends generally axially within the wall, and
including a generally axially extending intermediate passage
fluidly connecting the inlets to the outlets.
5. The turbine engine airfoil according to claim 4, wherein
multiple inlets each include an entrance at the interior cooling
channel, and the outlets each include an exit at the exterior
airfoil surface, the entrances having a greater cross-sectional
area than that of the exits.
6. The turbine engine airfoil according to claim 5, wherein a first
entrance includes an area that is greater than a second
entrance.
7. The turbine engine airfoil according to claim 5, wherein a first
exit has an area that is greater than a second exit.
8. The turbine engine airfoil according to claim 1, wherein the
cooling passage includes trip strips.
9. The turbine engine airfoil according to claim 1, wherein the
cooling passage is nonlinear.
10. A method of manufacturing the airfoil of claim 1, comprising
the steps of: providing a ceramic core corresponding to an interior
cooling channel; providing a refractory metal core corresponding to
a cooling passage; arranging the cores in a mold; and casting an
airfoil structure around the cores, wherein the airfoil structure
includes a wall separating the interior cooling channel from an
exterior airfoil surface, the cooling passage provided in the wall
fluidly connects the interior cooling channel to the exterior
airfoil surface, the cooling passage including at least one inlet
and multiple outlets respectively adjoining the interior cooling
channel and the exterior airfoil surface, at least one of a first
inlet and outlet having a different structural flow characteristic
than at least one of a second inlet and outlet.
11. The method according to claim 10, wherein the refractory metal
core providing step includes forming a desired core shape and
bending the formed desired core shape to correspond to the cooling
passage.
12. The method according to claim 11, wherein the refractory metal
core providing step includes providing notches in the cooling
passage corresponding to trip strips.
13. The method according to claim 11, wherein the bending step
includes the bending the cooling passage into generally an S-shape
in a lateral direction.
14. The method according to claim 10, wherein the arranging step
includes locating the refractory metal core relative to the ceramic
core.
15. The method according to claim 10, wherein the casting step
includes forming a diffuser feature in the cooling passage.
Description
BACKGROUND
[0001] This disclosure relates to a cooling passage for an
airfoil.
[0002] Turbine blades are utilized in gas turbine engines. As
known, a turbine blade typically includes a platform having a root
on one side and an airfoil extending from the platform opposite the
root. The root is secured to a turbine rotor. Cooling circuits are
formed within the airfoil to circulate cooling fluid, such as
compressor bleed air. Typically, multiple relatively large cooling
channels extend radially from the root toward a tip of the airfoil.
Air flows through the channels and cools the airfoil, which is
relatively hot during operation of the gas turbine engine.
[0003] Some advanced cooling designs use one or more radial cooling
passages arranged between the cooling channels and an airfoil
exterior surface that extend from the root toward the tip. The
cooling passages provide high convective cooling.
[0004] Other current airfoil tooling designs make use of some
cooling holes drilled through airfoil walls and into internal
cooling passages. In this type of configuration, the geometry of
the holes is limited to a straight hole with the possibility for
some flow diffusing feature added near the exit of the hole. As
holes must be drilled in a straight line, minimal angles with the
airfoil exterior surface must be observed. The length of holes is
dictated by manufacturing constraints.
SUMMARY
[0005] An example method of manufacturing an airfoil includes
providing a ceramic core corresponding to an interior cooling
channel. A refractory metal core is provided that corresponds to a
cooling passage. The cores are arranged in a mold. An airfoil
structure is cast about the cores to provide a turbine engine
airfoil.
[0006] The turbine engine airfoil includes a wall providing the
interior cooling channel and an exterior airfoil surface. The
cooling passage is provided in the wall and fluidly connects the
interior cooling channel to the exterior airfoil surface. The
cooling passage includes multiple inlets and multiple outlets
respectively adjoining the interior cooling channel and the
exterior airfoil surface. At least one of a first inlet and outlet
has a different structural flow characteristic than at least one of
a second inlet and outlet.
[0007] These and other features of the disclosure can be best
understood from the following specification and drawings, the
following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] FIG. 1 is a schematic view of an example gas turbine engine
incorporating the disclosed airfoil.
[0009] FIG. 2 is a perspective view of an example turbine
blade.
[0010] FIG. 3 is a cross-sectional view of a portion of the turbine
blade illustrated in FIG. 2.
[0011] FIG. 4 is an airfoil tip cross-sectional view through a
cooling passage in a wall of the airfoil structure shown in FIG.
3.
[0012] FIG. 5 is a partial cross-sectional view of a core assembly
arranged in a mold prior to casting the airfoil structure.
[0013] FIG. 6 is a perspective view of a portion of a refractory
metal core used to form the cooling passage shown in FIGS. 3 and
4.
[0014] FIGS. 6A-6C are cross-sectional views of portions of cooling
passage outlets illustrated in FIG. 6.
[0015] FIG. 7 is an enlarged top view of a portion of a cooling
passage outlet illustrated in FIG. 6.
DETAILED DESCRIPTION
[0016] A gas turbine engine (GTE) 10 is illustrated schematically
in FIG. 1. The GTE 10 includes a core section downstream from a fan
section 14. The core section 12 includes a compressor section 18
supplying compressed air to a combustor 20. The combusted air
expands over a turbine section 22 that rotationally drives a fan 16
within the fan section 14 about an axis A.
[0017] The turbine section 22 includes turbine blades 24 rotatable
about the axis A and arranged in a circumferential direction C,
shown in FIG. 2. One example turbine blade is illustrated in FIG.
2. The turbine blade 24 has a root 26 that supports a platform 28.
An airfoil structure 30 extends in a radial direction R from the
platform 28 to a tip 32. The airfoil structure 30 provides an
exterior airfoil surface 34 having leading and trailing edges 36,
38 with adjoining spaced apart sides 40.
[0018] Referring to FIGS. 3 and 4, the example turbine blade 24
includes a wall 44 that provides the exterior airfoil surface 34.
One or more interior cooling channels 42 are provided by the wall
44 and supply cooling air, for example, compressor bleed air, for
cooling the turbine blade 24. This cooling fluid is supplied to
various cooling features that ultimately flow through the wall 44
to provide internal convective cooling and a cooling film to the
exterior airfoil surface 34.
[0019] In the example, a cooling passage 46 fluidly interconnects
the interior cooling channel 42 to the exterior airfoil surface 34
and is arranged on the pressure side of the turbine blade 24. The
cooling passage 46 includes multiple inlets 48 adjoining a radially
extending intermediate passage 50. Multiple outlets 52 adjoin the
intermediate passage 50, which enables the pressure to be better
equalized across the outlets 52. The inlets 48 each provide an
entrance 54 at the interior cooling channel 42. The extended
intermediate passage 50 provide exits 56 arranged at the end of the
airfoil structure near the tip 32. The cooling passage 46 has a
generally S-shaped cross-section. The flow path from the entrance
54 to the exit 56 can replace the straight, drilled holes
previously used. Trip strips 58, schematically shown in FIG. 4, are
arranged in the cooling passage 46 as desired, for example, along
portions of the outlets 52 to improve cooling. Cross-section of the
trip strips can be any shapes such as block (as shown),
semi-circular, triangular, semi-elliptic, and alike. Pedestals may
also be provided.
[0020] In the example, the interior cooling channel 42 and cooling
passage 46 are provided by one or more ceramic cores arranged
within a mold. Referring to FIG. 5, a ceramic core 64 provides the
interior cooling channel 42. A refractory metal core (RMC) 66
provides the cooling passage 46. The ceramic core and the
refractory metal core are provided using different materials than
one another. One or more locating features 68, such as interlocking
protrusions and recesses, locate the RMC 66 relative to the ceramic
core 64. The cores 64, 66 are arranged within a cavity 62 of the
mold 60. The airfoil structure 30 is typically cast into the mold
60 to provide a structure, such as a single-crystal nickel alloy
structure.
[0021] The RMC 66 is formed to provide a desired core shape.
Typically, the RMC can be stamped out of a flat sheet metal.
Subsequently, this stamped RMC shape is bent to a desired shape to
provide a correspondingly shaped cooling passage 46, an example of
which is illustrated in FIG. 6. The RMC 66 includes a first and
second ends (generally, 70 and 72), which correspond to the inlets
and outlets 48 and 52, joined by a radially extending intermediate
portion 74. The first ends 70A, 70B respectively include a first
and second inlet area 76, 78 that can be different in shape and
size than one another. The outlets 72A, 72B, 72C include first,
second and third outlet areas 80, 82, 84 (shown in FIGS. 6A-6C and
respectively represented by cross-sectional lines A-A, B-B, C-C in
FIG. 6) that can be different than one another. Notches 86 are
provided in the RMC 66 to provide corresponding trip strips 58.
[0022] The RMC 66 can be configured provide different structural
flow characteristics with any desired geometry to produce holes of
any desired length, path and exit shape, for example. For example,
by utilizing different cross-sectional areas along the length of
the RMC 66 (for example in along the flow path from the entrance 54
to the exit 56), each hole may be designed to provide desired
pressure drop control across the radial length of the cooling
passage 46 rather than over pressurizing many of the drilled holes
with only a few holes optimized. The cooling passage 46 may include
any heat transfer augmentation features such as trip strips to
improve heat transfer characteristics and control pressure drops
through the holes. Diffuser features 90 may also be provided in the
cooling passage 46 and in the exits 56 (see, e.g., FIG. 4).
[0023] Although example embodiments have been disclosed, a worker
of ordinary skill in this art would recognize that certain
modifications would come within the scope of the claims. As another
example, the method disclosed above can be applied to manufacturing
blade outer air seals (BOAS). For that reason, the following claims
should be studied to determine their true scope and content.
* * * * *