U.S. patent application number 14/186218 was filed with the patent office on 2014-08-07 for turbine component cooling channel mesh with intersection chambers.
The applicant listed for this patent is Mikro Systems, Inc.. Invention is credited to Ching-Pang Lee, John J. Marra.
Application Number | 20140219818 14/186218 |
Document ID | / |
Family ID | 44533213 |
Filed Date | 2014-08-07 |
United States Patent
Application |
20140219818 |
Kind Code |
A1 |
Lee; Ching-Pang ; et
al. |
August 7, 2014 |
Turbine Component Cooling Channel Mesh with Intersection
Chambers
Abstract
A mesh (35) of cooling channels (35A, 35B) with an array of
cooling channel intersections (42) in a wall (21, 22) of a turbine
component. A mixing chamber (42A-C) at each intersection is wider
(W1, W2)) than a width (W) of each of the cooling channels
connected to the mixing chamber. The mixing chamber promotes swirl,
and slows the coolant for more efficient and uniform cooling. A
series of cooling meshes (M1, M2) may be separated by mixing
manifolds (44), which may have film cooling holes (46) and/or
coolant refresher holes (48).
Inventors: |
Lee; Ching-Pang;
(Cincinnati, OH) ; Marra; John J.; (Winter
Springs, FL) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Mikro Systems, Inc. |
Charlottesville |
VA |
US |
|
|
Family ID: |
44533213 |
Appl. No.: |
14/186218 |
Filed: |
February 21, 2014 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
12884486 |
Sep 17, 2010 |
8714926 |
|
|
14186218 |
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Current U.S.
Class: |
416/97R |
Current CPC
Class: |
F01D 5/18 20130101; F05D
2250/70 20130101; F05D 2260/2212 20130101; F01D 5/187 20130101;
F05D 2260/202 20130101 |
Class at
Publication: |
416/97.R |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Goverment Interests
STATEMENT REGARDING FEDERALLY SPONSORED DEVELOPMENT
[0001] Development for this invention was supported in part by
Contract No. DE-FC26-05NT42644, awarded by the United States
Department of Energy. Accordingly, the United States Government may
have certain rights in this invention.
Claims
1. A turbine component comprising: a mesh of cooling channels
comprising an array of cooling channel intersections located in a
wall of the turbine component; a mixing chamber located at each of
a plurality of the cooling channel intersections; wherein: each
mixing chamber comprises a width that is wider than a respective
width of each cooling channel connected thereto; each mixing
chamber comprises first and second widths that are perpendicular to
each other and equal to each other; and said two connected cooling
channels define respective longitudinal axes that intersect the
mixing chamber at an angle of 60 to 75 degrees with respect to each
other as viewed perpendicularly to a plane defined by at least one
of said longitudinal axes.
2. The turbine component of claim 1, wherein each mixing chamber
defines a shape that is not cylindrical or spherical.
3. The turbine component of claim 1, wherein the cooling channels
of the mesh are straight between the mixing chambers of the
mesh.
4. The turbine component of claim 1, wherein each mixing chamber
extends only within a depth range of said connected cooling
channels.
5. The turbine component of claim 1, wherein each mixing chamber
has a cylindrical or a spherical shape centered on the respective
intersection and a diameter that is greater than the respective
widths of the connected cooling channels.
6. The turbine component of claim 5, wherein each mixing chamber
comprises a spherical geometry that is truncated at opposite ends
thereof, limiting the mixing chamber to a depth range of said
connected channels.
7. The turbine component of claim 5, wherein the mixing chambers of
the mesh are separated by solid portions of the wall, each solid
portion comprising eight surfaces, alternating between straight
channel surfaces and spherical or cylindrical chamber surfaces.
8. The turbine component of claim 1, further comprising a coolant
inlet manifold along an inlet side of said interconnected mesh and
a coolant mixing manifold in the wall, wherein the coolant mixing
manifold extends along both an outlet side of said interconnected
mesh and along an inlet side of a second interconnected mesh
defined according to claim 1 within the wall.
9. The turbine component of claim 8, wherein the coolant mixing
manifold comprises coolant refresher holes that meter a coolant
into the coolant mixing manifold from a coolant supply channel in
the turbine component.
10. The turbine component of claim 8, wherein the coolant mixing
manifold comprises film cooling holes that meter a coolant from the
coolant mixing manifold to an outer surface of the wall.
11. The turbine component of claim 8, wherein the wall comprises
film cooling holes that meter a coolant from the coolant mixing
manifold to an outer surface of the wall and coolant refresher
holes that meter the coolant into the coolant mixing manifold from
a coolant supply channel in the turbine component, wherein the film
cooling holes are offset from the coolant refresher holes.
12. The turbine component of claim 1, further comprising a
refresher coolant inlet opening into each mixing chamber for
delivery of fresh coolant thereto.
13. A turbine component comprising: a first plurality of parallel
cooling channels located in a layer below a surface of a wall of
the component, each cooling channel from said first plurality of
parallel cooling channels defining a respective cooling channel
longitudinal axis; and a second plurality of parallel cooling
channels located in said layer, each cooling channel from said
second plurality of parallel cooling channels defining a respective
cooling channel longitudinal axis; wherein: viewed along an axis
substantially perpendicular to said surface, each cooling channel
longitudinal axis of the first plurality of parallel cooling
channels appears to intersect one or more cooling channel
longitudinal axes of the second plurality of parallel cooling
channels at an angle to define an interconnected mesh of the
cooling channels comprising an array of apparent intersections of
the cooling channels, each intersection comprising a mixing
chamber; each mixing chamber comprises a shape that defines an axis
that is substantially centered on the intersection and normal to
said surface; and each mixing chamber has a diameter greater than a
width of said each cooling channel of the intersection at a
mid-depth of the respective cooling channel.
14. The turbine component of claim 13, wherein a respective mixing
chamber extends only within a depth range of said each cooling
channel of the intersection.
15. The turbine component of claim 13, wherein the mixing chambers
of the mesh are separated by solid portions of the layer, each
solid portion comprising eight surfaces alternating between
straight channel surfaces and spherical or cylindrical chamber
surfaces.
16. The turbine component of claim 13, further comprising a coolant
inlet manifold along an inlet side of said interconnected mesh, and
a coolant mixing manifold in the wall, wherein the coolant mixing
manifold extends along an outlet side of said interconnected
mesh.
17. The turbine component of claim 16, wherein the coolant mixing
manifold comprises coolant refresher holes that meter a coolant
into the coolant mixing manifold from a coolant supply channel in
the turbine component.
18. The turbine component of claim 16, wherein the coolant mixing
manifold comprises film cooling holes that meter a coolant from the
coolant mixing manifold to an outer surface of the wall.
19. The turbine component of claim 16, wherein the wall comprises
film cooling holes that meter a coolant from the coolant mixing
manifold to an outer surface of the wall and coolant refresher
holes that meter coolant into the coolant mixing manifold from a
coolant supply channel in the turbine component, wherein the film
cooling holes are offset from the coolant refresher holes.
20. A turbine airfoil comprising: a first plurality of parallel
cooling channels located in a layer below a surface of an outer
wall of the airfoil, each cooling channel from said first plurality
of parallel cooling channels defining a respective cooling channel
longitudinal axis; a second plurality of parallel cooling channels
located in said layer; , each cooling channel from said second
plurality of parallel cooling channels defining a respective
cooling channel longitudinal axis wherein: viewed along an axis
substantially perpendicular to said surface, each cooling channel
longitudinal axis of the first plurality of parallel cooling
channels appears to intersect one or more cooling channel
longitudinal axes of the second plurality of parallel cooling
channels at an angle of 60 to 75 degrees in a first interconnected
mesh of the cooling channels comprising an array of intersections
of the cooling channels; each intersection comprising a mixing
chamber that is wider than each cooling channel of the intersection
at a mid-depth of said each cooling channel of the intersection;
the cooling channels of the mesh are straight between the mixing
chambers of the mesh; a coolant inlet manifold located along an
inlet side of said first interconnected mesh; a coolant mixing
manifold located in the wall along an outlet side of said first
interconnected mesh and along an inlet side of a second
interconnected cooling channel mesh within the layer; wherein: the
coolant mixing manifold comprises film cooling outlet holes or
coolant refresher inlet holes.
Description
FIELD OF THE INVENTION
[0002] This invention relates to cooling channels in turbine
components, and particularly to cooling channels intersecting to
form a cooling mesh in a turbine airfoil.
BACKGROUND OF THE INVENTION
[0003] Stationary guide vanes and rotating turbine blades in gas
turbines often have internal cooling channels. Cooling
effectiveness is important in order to minimize thermal stress on
these airfoils. Cooling efficiency is important in order to
minimize the volume of air diverted from the compressor for
cooling.
[0004] Film cooling provides a film of cooling air on outer
surfaces of an airfoil via holes in the airfoil outer surface from
internal cooling channels. Film cooling can be inefficient because
so many holes are needed that a high volume of cooling air is
required. Thus, film cooling is used selectively in combination
with other techniques.
[0005] Perforated cooling tubes may be inserted into span-wise
channels in an airfoil to create impingement jets against the inner
surfaces of the airfoil. A disadvantage is that heated
post-impingement air moves along the inner surfaces of the airfoil
and interferes with the impingement jets. Also, impingement tubes
require a nearly straight airfoil for insertion, but some turbine
airfoils have a curved span for aerodynamic efficiency.
[0006] Cooling channels may form an interconnected mesh that does
not require impingement tube inserts, and can be formed in curved
airfoils. The present invention improves efficiency and
effectiveness in a cooling channel mesh.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] The invention is explained in the following description in
view of the drawings that show:
[0008] FIG. 1 is a transverse sectional view of a prior art turbine
vane with impingement cooling inserts.
[0009] FIG. 2 is a side view of a prior art curved turbine vane
airfoil between radially inner and outer platforms.
[0010] FIG. 3 is a transverse sectional view of a prior art turbine
airfoil with mesh cooling channels.
[0011] FIG. 4 is a perspective view of the prior art turbine
airfoil of FIG. 3.
[0012] FIG. 5 is a sectional view of a cooling channel mesh per
aspects of the invention.
[0013] FIG. 6 is a transverse sectional view of an airfoil per
aspects of the invention.
[0014] FIG. 7 is a sectional view of a series of two cooling
meshes.
[0015] FIG. 8 is a perspective view of part of a casting core that
forms a spherical mixing chamber per aspects of the invention.
[0016] FIG. 9 is a perspective view of part of a casting core that
forms a truncated spherical mixing chamber per aspects of the
invention.
[0017] FIG. 10 is a perspective view of part of a casting core that
forms a cylindrical mixing chamber per aspects of the
invention.
DETAILED DESCRIPTION OF THE INVENTION
[0018] FIG. 1 is a transverse sectional view of a prior art turbine
airfoil 20A with a pressure side wall 21, a suction side wall 22, a
leading edge 23, a trailing edge 24, internal cooling channels 25,
26, impingement cooling baffles 27, 28, film cooling holes 29, and
coolant exit holes 30. The impingement cooling baffles are
thin-walled tubes inserted into the cooling channels 25, 26. They
are spaced apart from the channel walls. Cooling air enters an end
of each impingement baffle 27, 28, and flows span-wise within the
vane. It exits impingement holes 31, and impinges on the walls 21,
22.
[0019] FIG. 2 is a side view of a prior art curved turbine vane
airfoil 20B that spans between radially inner and outer platforms
32, 33. The platforms are mounted in a circular array of adjacent
platforms, forming an annular flow path for a working gas 34 that
passes over the vanes. This type of curved airfoil can make
insertion of impingement baffles 27, 28 impractical, so other
cooling means are needed.
[0020] FIG. 3 shows a prior art turbine airfoil 20C with a pressure
side wall 21 and a suction side wall 22 and a cooling channel mesh
35. A coolant supply channel 36 is separated from a coolant inlet
manifold 37 by a partition 38 with impingement holes 39. Coolant
jets 40 impinge on the inside surface of the leading edge 23, then
the coolant flows 41 into the mesh 35, and exits the trailing edge
exit holes 30.
[0021] FIG. 4 shows a perspective view of the prior art turbine
airfoil 20C of FIG. 3. The mesh 35 comprises a first plurality of
parallel cooling channels 35A, and a second plurality of parallel
cooling channels 35B, wherein the first and second plurality of
cooling channels intersect each other in a plane or level below a
surface of the airfoil, forming channel intersections 42. The
cross-sectional shape of the cooling channels may be either
circular or non-circular, including rectangular, square or
oval.
[0022] FIG. 5 shows a cooling mesh per aspects of the invention.
Each channel intersection has a mixing chamber 42A, which may be
spherical or cylindrical. The mixing chamber delays the coolant
flow, increasing heat transfer, and it provides a space and shape
for swirl, increasing uniformity and efficiency of cooling. The
mixing chambers 42A have a width W1 that is greater than a width W
of each of the channels opening into the chamber. Each cooling
channel 35A, 35B may have a width dimension W defined at mid-depth
of the channel as shown in FIG. 9. The mid-depth may be defined by
a geometric centerline 45 of the cooling channel as shown in FIGS.
8-10. The mixing chambers may have equal perpendicular widths W1,
W2, thus providing a chamber shape that promotes swirl. If the
mixing chambers are spherical or cylindrical, then each width W1,
W2 is a diameter thereof. The term "width" herein refers to a
transverse dimension measured at mid-depth 45 of the channels
connected to the mixing chamber.
[0023] Spherical and cylindrical mixing chambers have spherical or
cylindrical surfaces 43B between the four channel openings in the
chamber. Solid parts 43 of the wall 21, 22 separate adjacent mixing
chambers 42A and may have four channel surfaces 43A and four
chamber surfaces 43B. Thus, the solid parts 43 may have eight
surfaces alternating between straight channel surfaces 43A and
spherical or cylindrical surfaces 43B. This geometry maximizes the
surface area of the channels 35A, 35B for a given volume of the
mixing chambers 42A, and provides symmetrical mixing chambers for
swirl.
[0024] FIG. 6 is a sectional view of an airfoil per aspects of the
invention. The cooling channel mesh 35 is formed in a layer below
the surface of the walls 21, 22, as delineated by dashed lines. A
coolant supply channel 36 may be separated from a coolant inlet
manifold 37 by a partition 38 with impingement holes 39. Coolant
jets 40 may impinge on the inside surface of the leading edge 23.
Then the coolant flows 41 into the mesh 35, and exits the trailing
edge exit holes 30. The mesh 35 may follow the design of FIG. 5.
Periodic mixing manifolds 44 may be provided along the coolant flow
path in the walls 21, 22 for additional span-wise mixing. These
mixing manifolds 44 are closed off at the top and bottom. Film
cooling holes 46 may pass between a mixing manifold 44 and an outer
surface of the airfoil. Coolant refresher holes 48 may meter
coolant from the coolant supply channel 36 into the mixing manifold
44. The refreshment coolant flowing into the manifold 44 not only
reduces the temperature of the bulk fluid, but it also provides
momentum energy along a vector for additional mixing within the
manifold.
[0025] FIG. 7 is a sectional view of a series of two cooling meshes
M1, M2, separated by a mixing manifold 44. A coolant inlet manifold
37 receives coolant via one or more supply channels from the
turbine cooling system. The coolant inlet manifold 37 may be a
leading edge manifold as shown in FIG. 6. Or it may be at another
location, such as the locations of the mixing manifolds 44 shown in
FIG. 6. Coolant 41 flows through the first mesh M1, and then enters
a mixing manifold 44, which may include film cooling holes 46
and/or coolant refresher holes 48 as shown in FIG. 6. The coolant
then flows through the second cooling mesh M2. This sequence of
alternating meshes and mixing manifolds 44 may be repeated.
Finally, the coolant may exit through trailing edge exit holes 30
or it may be recycled in a closed-loop cooling system not
shown.
[0026] The intersection angle AA of the first and second cooling
channels 35A, 35B may be perpendicular, or not perpendicular, as
shown. Shallower intersection angles provide more direct coolant
flow between the manifolds 37, 44. An angle AA between 60.degree.
and 75.degree. provides a good combination of coolant throughput
and mixing, although other angles may be used.
[0027] The meshes M1, M2 and/or the mixing chambers 42A-C may vary
in size, density, or shape along a cooled wall depending on the
heating topography of the wall. The mixing manifolds 44 may vary in
spacing and type for the same reason. For example, coolant
refresher holes 48 may be spaced more closely on the leading half
of the pressure side wall 21 than in other areas. Likewise for film
cooling holes 46. Both film cooling holes and refresher holes may
be provided in the same mixing manifold 44 and they may offset from
each other to avoid immediate exit of refresher coolant.
[0028] FIG. 8 illustrates part of a casting core that forms a
spherical mixing chamber 42A by defining a volume that is
unavailable to molten metal during a casting process. FIG. 9
illustrates part of a casting core that forms a spherical mixing
chamber 42B that is truncated at opposite ends to the extent of
depth range D of the channels 35A, 35B connected thereto.
Truncation allows thinner component walls 21, 22. FIG. 10
illustrates part of a casting core that forms a cylindrical mixing
chamber 42C with an axis 50 centered on the intersection and normal
to the outer surface of the wall 21, 22. The cylindrical mixing
chamber may be truncated to the depth range D of the connected
channels 35A, 35B.
[0029] The mixing chambers may take shapes other than cylindrical
or spherical. However, a cylindrical or spherical shape of the
mixing chambers 42A-C beneficially guides the flow 41 into a
circular swirl that provides predictable mixing, and maximizes the
chamber volume while minimizing reduction of the channel
length.
[0030] Herein, the term "cooling air" is used to mean any cooling
fluid for internal cooling of turbine airfoils. In some cases,
steam may be used. The term "straight channel" or "straight span"
means a channel or segment thereof with a straight geometric
centerline and without flared or constricted walls.
[0031] While various embodiments of the present invention have been
shown and described herein, it will be obvious that such
embodiments are provided by way of example only. Numerous
variations, changes and substitutions may be made without departing
from the invention herein. Accordingly, it is intended that the
invention be limited only by the spirit and scope of the appended
claims.
* * * * *