U.S. patent application number 11/896157 was filed with the patent office on 2009-03-05 for multi-part cast turbine engine component having an internal cooling channel and method of forming a multi-part cast turbine engine component.
This patent application is currently assigned to General Electric Company. Invention is credited to Thomas Moors.
Application Number | 20090060714 11/896157 |
Document ID | / |
Family ID | 39717719 |
Filed Date | 2009-03-05 |
United States Patent
Application |
20090060714 |
Kind Code |
A1 |
Moors; Thomas |
March 5, 2009 |
Multi-part cast turbine engine component having an internal cooling
channel and method of forming a multi-part cast turbine engine
component
Abstract
A multi-part cast component for a turbine engine includes a
first component section having a main body portion including at
least one cooling flow passage section, and a second component
section having a main body including at least one cooling flow
passage section. The first and second component sections are joined
along a parting line to form a turbine engine component with the at
least one cooling flow passage section of the first component
section aligning with the at least one cooling flow passage of the
second component section to form a cooling flow channel.
Inventors: |
Moors; Thomas; (Greenville,
NC) |
Correspondence
Address: |
CANTOR COLBURN, LLP
20 Church Street, 22nd Floor
Hartford
CT
06103
US
|
Assignee: |
General Electric Company
Schenectady
NY
|
Family ID: |
39717719 |
Appl. No.: |
11/896157 |
Filed: |
August 30, 2007 |
Current U.S.
Class: |
415/115 ;
29/889.721 |
Current CPC
Class: |
F05D 2230/211 20130101;
F01D 5/187 20130101; Y10T 29/49341 20150115; F05D 2230/23
20130101 |
Class at
Publication: |
415/115 ;
29/889.721 |
International
Class: |
F01D 5/14 20060101
F01D005/14 |
Claims
1. A multi-part cast component for a turbine engine comprising: a
first component section having a main body portion including at
least one cooling flow passage section; and a second component
section having a main body portion including at least one cooling
flow passage section, said second component section being joined to
the first component section along a parting line to form a turbine
engine component with the at least one cooling flow passage section
of the first component section aligning with the at least one
cooling flow passage section of the second component section to
form a cooling flow channel.
2. The multi-part cast component according to claim 1, wherein the
first and second component sections are joined to form a two-part
turbine engine component.
3. The multi-part cast component according to claim 1, wherein the
first component section is a first rotor blade section and the
second component section is a second rotor blade section, said
first and second rotor blade sections being joined along the
parting line to form a turbine rotor blade.
4. The multi-part cast component according to claim 1, wherein the
at least one cooling flow passage section of the first component
section includes first and second cooling flow passage
sections.
5. The multi-part cast component according to claim 3, wherein the
at least one cooling flow passage section of the second component
section includes third and fourth cooling flow passage sections,
said first and third cooling flow passage sections being joined to
form a first cooling flow channel and said second and fourth
cooling flow passage sections being joined to form a second cooling
flow channel.
6. A method of forming a multi-part cast component for a turbine
engine comprising: forming a first component section; creating at
least one cooling flow passage section in the first component
section; forming a second component section; creating at least one
cooling flow passage section in the second component section; and
joining the first and second component section along a parting line
to form a turbine engine component with said at least one cooling
flow passage section of the first component section joining to the
at least one cooling flow passage section of the second component
section to form a cooling flow channel in the turbine
component.
7. The method of claim 6, wherein the turbine engine component is a
turbine rotor blade.
8. The method of claim 6, further comprising: forming first and
second cooling flow passage sections in the first component
section; and forming third and fourth cooling flow passage sections
in the second component section wherein, upon joining the first and
second component sections along the parting line, said first
cooling flow passage section registers with the third cooling flow
passage section to establish a first cooling flow channel and said
second cooling flow passage section registers with the fourth
cooling flow passage section to establish a second cooling flow
channel in the turbine component.
9. The method of claim 6, wherein the at least one cooling flow
passage section is machined from corresponding ones of the first
and second component sections.
10. The method of claim 6, wherein the at least one cooling flow
passage section is molded into corresponding ones of the first and
second component sections.
11. The method of claim 5, wherein the turbine engine component is
formed in two parts.
Description
BACKGROUND OF THE INVENTION
[0001] The present invention pertains to the art of turbine engines
and, more particularly, to a two part cast turbine engine component
having internally formed cooling cavities.
[0002] In general, gas turbine engines combust a fuel/air mixture
to release heat energy in the form of a high temperature gas stream
that is channeled to a turbine section via a hot gas path. More
specifically, a compressor compresses incoming air to a high
pressure. The high pressure air is delivered to a combustion
chamber to mix with fuel and form a combustible mixture. The
combustible mixture is then ignited to form a high pressure, high
velocity gas stream which is delivered to the turbine. The high
pressure air impacts upon rotor blades or buckets that form part of
a turbine rotor assembly. In this manner, the turbine converts
thermal energy from the high temperature, high velocity gas stream
to mechanical energy that rotates a turbine shaft.
[0003] In many cases, a cooling gas is delivered to internal
portions of each rotor blade in order to lower temperatures,
particularly in air foil portions of the rotor blade. The cooling
gas is delivered through internal passages integrally molded with
the blade. In some cases, the passages are formed using a lost wax
investment casting method. In other cases, the passages are formed
around a ceramic core. In either case, the passages are difficult
to manufacture, expensive, and limited in shape. In addition,
ceramic cores can react negatively with various alloys used in
creating the blades. Moreover, ceramic cores are fragile and prone
to shifting and breakage. Finally, internal surfaces of the
passages may contain residual core defects, bad grain orientation
or finning. The above manufacturing methods make visual inspection
of the passages difficult, if not impossible.
BRIEF DESCRIPTION OF THE INVENTION
[0004] In accordance with one aspect, the present invention
provides a multi-part cast component for a turbine engine. The
multi-part cast component includes a first component section having
a main body portion including a interior portion that defines a
first cooling flow passage and a second component section having a
main body portion including a interior portion that defines a
second cooling flow passage section. The first and second component
sections are joined along a parting line to form a turbine engine
component with the first and second cooling flow passage sections
aligning to form a cooling flow channel.
[0005] In accordance with another aspect, the present invention
provides a method of forming a multi-part cast component for a
turbine engine. The method includes casting a first component
section, forming a first cooling flow passage section in the first
component section, casting a second component section and forming a
second cooling flow passage section in the second turbine component
section. The method also requires joining the first and second
component sections along a parting line to form a turbine engine
component, with the first and second cooling flow passage sections
aligning to establish a cooling flow channel.
[0006] It should be appreciated that the present invention provides
a two part cast turbine engine component having an interior cooling
channel whose formation is readily formed and which can be easily
visually inspected while avoiding many of the drawbacks associated
with other casting methods. In any event, additional objects,
features and advantages of the various aspects of the present
invention will become more readily apparent from the following
detailed description when taken in conjunction with the drawings
wherein like reference numerals refer to corresponding parts in the
several views.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] FIG. 1 is a schematic illustration of a gas turbine engine
including a two-part cast engine component, shown in the form of a
rotor blade, constructed in accordance with an aspect of the
invention;
[0008] FIG. 2 is an enlarged perspective view of the rotor blade of
FIG. 1;
[0009] FIG. 3 is an end view of the rotor blade of FIG. 1; and
[0010] FIG. 4 is an elevational view showing two component sections
of the rotor blade of FIG. 1 prior to being joined.
DETAILED DESCRIPTION OF THE INVENTION
[0011] With initial reference to FIG. 1, a gas turbine engine
constructed in accordance with the present invention is generally
indicated at 10. Turbine engine 10 includes a compressor 12
operatively coupled to a turbine 14 and an electrical generator 16
via a shaft 18. Shaft 18 is illustrated as a single, monolithic
component, however, it should be readily understood that shaft 18
could also be formed in multiple segments with each segment being
coupled to an adjacent engine component.
[0012] In any event, engine 10 is further shown to include a
combustor 20 in which air 21 from compressor 12 and a fuel 22 are
mixed to form a combustible mixture. The combustible mixture is
ignited to form a high pressure, high temperature combustion
product or gas 28 that is used to drive turbine 14. More
specifically, high pressure, high temperature gas 28 enters into
turbine 14 and impinges upon a rotor assembly 35 having a plurality
of rotor blades, two of which are indicated at 40 and 41. Rotor
assembly 35 converts thermal energy from high pressure, high
temperature gas 28 into mechanical, rotational, energy.
[0013] When coupled to rotor assembly 35, rotor blades 40 and 41
are connected to a rotor disk (not shown) that is rotatably mounted
to a rotor shaft, such as shaft 18. In an alternative
configuration, rotor blades 40 and 41 are mounted within a rotor
spool (not shown). In any event, circumferentially adjacent rotor
blades 40 and 41 are identical such that a detailed description
will follow referring to rotor blade 40. However, it should be
understood that in the exemplary embodiment, each of the plurality
of rotor blades is similarly constructed.
[0014] As best shown in FIGS. 2 and 3, rotor blade 40 includes an
airfoil portion 60, a platform portion 62, a shank portion 64 and a
dovetail 66 which are collectively known as a bucket. Each airfoil
portion 60 includes a first sidewall 70 and a second sidewall 72.
In the embodiment shown, first sidewall 70 is convex and defines a
suction side of airfoil portion 60. Conversely, second sidewall 72
is concave and defines a pressure side of airfoil portion 60.
Sidewalls 70 and 72 collectively form a leading edge 74 and an
axially spaced trailing edge 76 of airfoil portion 60. More
specifically, trailing edge 76 is spaced chord-wise and downstream
from leading edge 74.
[0015] As will be discussed more fully below, leading edge 74 is
provided with a plurality of openings 77 that serve as ventilation
ducts allowing a cooling gas to pass through rotor blade 40.
Likewise, trailing edge 76 includes a plurality of openings 78 that
also serve as cooling gas ventilation ducts. First and second
sidewalls 70 and 72 extend outward in span from a bucket or blade
root 79 located adjacent to platform portion 62 to an airfoil tip
80. As best shown in FIG. 3, airfoil tip 80 includes an opening 82
defined by a recessed seat 84. A tip cap 85 is provided in recessed
seat 84 and closes opening 82. However, tip cap 85 may be provided
with small openings, such as indicated at 87 that allow a small
portion of the cooling gas to pass into a hot gas path (not
shown).
[0016] In the embodiment shown, shank portion 64 extends radially
inward from platform portion 62 to dovetail 66, and dovetail 66
extends radially inward from shank portion 64 to facilitate
securing rotor blade 40 to the rotor disk (not shown). Platform
portion 62 also includes an upstream side or skirt 107 and a
downstream side or skirt 109 that are connected by a pressure-side
edge 111 and a suction-side edge 112. Shank portion 64 is shown to
include a substantially concave side wall 120 and a substantially
convex sidewall (not shown) that are connected at an upstream
sidewall 124 and a downstream sidewall 126. In this manner, concave
sidewall 120 is recessed relative to upstream and downstream
sidewalls 124 and 126 respectively, such that when rotor blade 40
is coupled within rotor assembly 35, a shank cavity 128 is defined
between adjacent rotor blades 40 and 41.
[0017] As further shown in FIGS. 2 and 3, rotor blade 40 includes a
forward angel wing 130 and an aft angel wing 132 which extend
outward from upstream side wall 124 and downstream sidewall 126
respectively. Forward and aft angel wings 130 and 132 are
configured to seal corresponding forward and aft angel wing
cavities (not shown) defined within rotor assembly 35. In addition,
rotor blade 40 includes a forward, lower angel wing 134 which
extends outward from upstream sidewall 124 to facilitate sealing
between rotor blade 40 and the rotor disk (not shown).
[0018] In accordance with one aspect of the present invention,
rotor blade 40 is a multi-part, cast component of engine 10. More
specifically, rotor blade 40 includes a first component section 142
and a second component section 144 which are cast separately and
joined together along a parting line 150. In accordance with an
exemplary embodiment, first and second component sections 142 and
144 are cast from aluminum. However, component sections 142 and 144
can be formed from a variety of materials, such as super alloys,
and through a variety of known forming techniques.
[0019] Reference will now be made to FIG. 4 is describing first and
second component sections 142 and 144 of rotor blade 40. As shown,
first component section 142 includes a main body portion 160 having
a interior portion 162 that defines first and second cooling gas
passage sections 164 and 165. As shown, first cooling gas passage
section 164 includes an inlet section 168 which, in the exemplary
embodiment, is bifurcated and leads into a flow chamber 170. Flow
chamber 170 empties into a first flow section 171 that extends
longitudinally through interior portion 162 to a first flow
reversing section 172. From first flow reversing section 172, first
cooling flow passage section 164 leads to a flow return section 174
that extends longitudinally back through interior portion 162. Flow
return section 174 terminates in a second flow reversing section
176 that leads to an outlet flow section 178. Outlet flow section
178 delivers cooling gas through openings 77 in leading edge 74 as
well as openings 87 in airfoil tip 80.
[0020] In a manner similar to that described above, second cooling
flow passage section 165 includes an inlet section 188 which, in
the exemplary embodiment, is bifurcated and leads into a flow
chamber 190. Flow chamber 190 empties into a first flow section 191
that extends longitudinally through interior portion 162 to a first
flow reversing section 192. From first flow reversing section 192,
second cooling flow passage section 165 leads to a flow return
section 194 that extends longitudinally back through interior
portion 162. Flow return section 194 terminates in a second flow
reversing section 196 that leads to an outlet flow section 198.
Outlet flow section 198 delivers cooling gas through openings 78 in
trailing edge 76 as well as openings 82 in airfoil tip 80. At this
point it should be understood that while first and second cooling
flow passage sections 164 and 165 appear to be similar, the
particular path taken by each cooling passage section 164, 165
could differ depending upon the particular configuration or
geometry of rotor blade 40.
[0021] In a manner similar to that described above with respect to
first component section 142, second component section 144 includes
a main body portion 210 having a interior portion 212 that defines
third and fourth cooling flow passage sections 214 and 215. Second
component section 144 is actually a mirror image of first component
section 142. Accordingly, third cooling flow passage section 214 is
a mirror image of first cooling flow passage section 164 while
fourth cooling flow passage section 215 is a mirror image of second
cooling flow passage section 165. However, for the sake of
completeness, third cooling flow passage section 214 includes an
inlet section 218 which, in the exemplary embodiment, is bifurcated
and leads into a flow chamber 220. Flow chamber 220 empties into a
first flow section 221 that extends longitudinally through interior
portion 212 to a first flow reversing section 222. From first flow
reversing section 222, third cooling flow passage section 214 leads
to a flow return section 224 that extends longitudinally back
through interior portion 212. Flow return section 224 terminates in
a second flow reversing section 226 that leads to an outlet flow
section 228. Outlet flow section 228 delivers cooling gas through
openings 77 in leading edge 74 as well as openings 87 in airfoil
tip 80.
[0022] Also in a similar manner, fourth cooling flow passage
section 215 includes an inlet section 238 which, in the exemplary
embodiment, is bifurcated and leads into a flow chamber 240. Flow
chamber 240 empties into a first flow section 241 that extends
longitudinally through interior portion 212 to a first flow
reversing section 242. From first flow reversing section 242,
fourth cooling flow passage section 215 leads to a flow return
section 244 that extends longitudinally back through interior
portion 212. Flow return section 244 terminates in a second flow
reversing section 246 that leads to an outlet flow section 248.
Outlet flow section 248 delivers cooling gas through openings 78 in
trailing edge 76 as well as openings 82 in airfoil tip 80.
[0023] With this arrangement, cooling flow passages 164, 165 and
214, 215 are formed in respective first and second component
sections 142 and 144 using a variety of techniques such as
machining, molding and the like. Regardless of the technique
employed, once formed, cooling flow passage sections 164, 165 and
214, 215 can be readily visually inspected for irregularities,
which might detract from overall cooling efficiency. Next, first
and second component sections 142 and 144 are joined along parting
line 150 so that first cooling flow passage section 164 registers
with third cooling flow passage section 214 to form a first cooling
flow channel 300 in rotor blade 40. Similarly, second cooling flow
passage section 165 registers with fourth cooling flow passage
section 215 to form a second cooling flow channel 304 in rotor
blade 40. Component sections 142 and 144 can be joined by a variety
of known metal joining techniques such as welding, brazing and the
like. Of course, if a super alloy or a material other than metal is
used to form rotor blade 40, other joining techniques would be
employed. In any event, once formed, rotor blade 40 is incorporated
into rotor blade assembly 35 of a turbine engine 2.
[0024] At this point it should be understood that the various
aspects of the present invention lower manufacturing costs and
improve overall component quality. In addition, by forming the
component in multiple parts, new cooling channel shapes and designs
are now possible. That is, cooling channels can now be created that
include thick and thin portions or even alternating thick and thin
portions. Moreover, the present invention enables the formation of
extremely small channels that can advantageously deliver a cooling
gas to portions of the rotor blade heretofore unreachable by
present passage formation techniques. More intricate serpentine
shapes that carry the cooling medium through a larger portion of
the component are also now possible. Ceramic core molding also has
a tendency to react negatively with certain metals. The present
invention avoids these problems. Finally the present invention
enables close, visual, inspection of the component parts to check
for voids or other casting defects that were heretofore
undetectable. That is, existing ceramic core molding is not capable
of forming channels having differing dimensions or passages that
are extremely thin or possess other such characteristics. When
using ceramic core molding techniques, once the channels are
formed, the ceramic core must be carefully removed. Removing a
ceramic core from channels that are thick and thin or extremely
narrow is not possible without risking breakage. If the core
breaks, the rotor blade must be discarded.
[0025] Although described with reference to illustrated aspects of
the present invention, it should be readily understood that various
changes and/or modifications can be made to the invention without
departing from the scope thereof. For instance, while the first and
second cooling flow passages are shown to be similar, various other
configurations could also be employed without detracting from the
invention. Also, while the component is described as being cast
from aluminum, other metals including super alloys and non-metals
can also be used depending on the particular application of the
engine. It should also be understood that while described as a
rotor blade, various other engine components such as vanes,
buckets, nozzles and the like could also be formed by the present
invention. Finally, the component could be cast in any number of
parts and should not be seen as being limited to the two component
sections as shown. In general, the invention is only intended to be
limited by the scope of the following claims.
* * * * *