U.S. patent application number 12/372368 was filed with the patent office on 2009-10-22 for manufacture of cmc articles having small complex features.
This patent application is currently assigned to GENERAL ELECTRIC COMPANY. Invention is credited to Douglas M. CARPER, James D. STEIBEL, Suresh SUBRAMANIAN, Stephen M. WHITEKER.
Application Number | 20090261508 12/372368 |
Document ID | / |
Family ID | 37969616 |
Filed Date | 2009-10-22 |
United States Patent
Application |
20090261508 |
Kind Code |
A1 |
STEIBEL; James D. ; et
al. |
October 22, 2009 |
MANUFACTURE OF CMC ARTICLES HAVING SMALL COMPLEX FEATURES
Abstract
The present invention is ceramic matrix composite gas turbine
engine component comprising a plurality of cured ceramic matrix
composite plies, each ply comprising ceramic fiber tows, each
ceramic fiber tow comprising a plurality of ceramic fibers, the
tows in each ply lying adjacent to one another such that each ply
has a unidirectional orientation. The component further comprises a
layer of a coating on the ceramic fibers. The component further
comprises a ceramic matrix material lying in interstitial regions
between the fibers and tows of each ply and the interstitial region
between the plurality of plies, wherein at least a portion of the
component is no greater than about 0.021 inch thick. The present
invention is also a method for making such a ceramic matrix
composite component.
Inventors: |
STEIBEL; James D.; (Mason,
OH) ; WHITEKER; Stephen M.; (Covington, KY) ;
CARPER; Douglas M.; (Trenton, OH) ; SUBRAMANIAN;
Suresh; (Mason, OH) |
Correspondence
Address: |
MCNEES WALLACE & NURICK LLC
100 PINE STREET, P.O. BOX 1166
HARRISBURG
PA
17108-1166
US
|
Assignee: |
GENERAL ELECTRIC COMPANY
Schenectady
NY
|
Family ID: |
37969616 |
Appl. No.: |
12/372368 |
Filed: |
February 17, 2009 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
11359217 |
Feb 22, 2006 |
7507466 |
|
|
12372368 |
|
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Current U.S.
Class: |
264/319 |
Current CPC
Class: |
F05D 2300/611 20130101;
Y02T 50/60 20130101; F05C 2203/0839 20130101; F05D 2230/90
20130101; C04B 2237/704 20130101; C04B 2237/365 20130101; C04B
35/80 20130101; C04B 2235/5268 20130101; F05D 2250/312 20130101;
F01D 5/284 20130101; C04B 35/806 20130101; C04B 35/62868 20130101;
C04B 2237/32 20130101; Y10T 428/249929 20150401; C04B 2237/363
20130101; F01D 5/288 20130101; C04B 2235/5244 20130101; C04B
2237/61 20130101; Y10T 428/249928 20150401; F05D 2300/2261
20130101; F01D 5/282 20130101; Y02T 50/672 20130101; C04B 35/62863
20130101; C04B 2235/5256 20130101; F05D 2300/2283 20130101; C04B
35/573 20130101; Y02T 50/67 20130101; C04B 2237/38 20130101; B32B
18/00 20130101 |
Class at
Publication: |
264/319 |
International
Class: |
C04B 35/622 20060101
C04B035/622 |
Claims
1-13. (canceled)
14. A method for manufacturing an uncooled ceramic matrix composite
gas turbine engine component comprising the steps of: providing a
plurality of prepreg ceramic plies, the plies comprising prepreg
ceramic fiber tows, the tows in each ply lying adjacent to one
another in a planar arrangement such that each ply has a
unidirectional orientation; laying up the plurality of prepreg
ceramic plies in a preselected arrangement to form a component
shape; heating the component shape to form a ceramic preform; and
densifying the turbine blade preform with at least silicon to form
an uncooled ceramic matrix composite turbine blade, wherein at
least a portion of the component comprises cured ceramic plies
having a thickness in the range of about 0.005 inch to about 0.007
inch, inclusive of tows, coating, and interstitial ceramic matrix
material.
15. The method of claim 14, wherein the step of laying up comprises
laying up to form a component shape that comprises an airfoil.
16. The method of claim 14, wherein the fiber tows are silicon
carbide fiber tows.
17. A method for manufacturing a cooled ceramic matrix composite
gas turbine engine component comprising the steps of: providing a
plurality of prepreg ceramic plies, the plies comprising prepreg
ceramic fiber tows, the tows in each ply lying adjacent to one
another in a planar arrangement such that each ply has a
unidirectional orientation; laying up the plurality of prepreg
ceramic plies in a preselected arrangement around tooling to form a
component shape; heating the component shape to form a ceramic
preform; removing the tooling; densifying the preform with at least
silicon to form a ceramic matrix composite component wherein at
least a portion of the component comprises cured ceramic plies
having a thickness in the range of about 0.005 inch to about 0.007
inch, inclusive of tows, coating, and interstitial ceramic matrix
material; and drilling cooling holes in the ceramic matrix
composite component to form a cooled ceramic matrix composite
component.
18. The method of claim 17, wherein the step of laying up comprises
laying up to form a component shape that comprises an airfoil.
19. The method of claim 17, wherein the step of laying up comprises
laying up to form a component shape that is a turbine blade.
20. The method of claim 17, wherein the ceramic fiber tows are
silicon carbide fiber tows.
Description
FIELD OF THE INVENTION
[0001] The present invention relates generally to ceramic matrix
turbine engine components, and more particularly, to a ceramic
matrix composite gas turbine engine component have small complex
features.
BACKGROUND OF THE INVENTION
[0002] In order to increase the efficiency and the performance of
gas turbine engines so as to provide increased thrust-to-weight
ratios, lower emissions and improved specific fuel consumption,
engine turbines are tasked to operate at higher temperatures. As
the higher temperatures reach and surpass the limits of the
material comprising the components in the hot section of the engine
and in particular the turbine section of the engine, new materials
must be developed.
[0003] As the engine operating temperatures have increased, new
methods of cooling the high temperature alloys comprising the
combustors and the turbine airfoils have been developed. For
example, ceramic thermal barrier coatings (TBCs) were applied to
the surfaces of components in the stream of the hot effluent gases
of combustion to reduce the heat transfer rate and to provide
thermal protection to the underlying metal and allow the component
to withstand higher temperatures. These improvements helped to
reduce the peak temperatures and thermal gradients. Cooling holes
were also introduced to provide film cooling to improve thermal
capability or protection. Simultaneously, ceramic matrix composites
were developed as substitutes for the high temperature alloys. The
ceramic matrix composites (CMCs) in many cases provided an improved
temperature and density advantage over the metals, making them the
material of choice when higher operating temperatures were
desired.
[0004] A number of techniques have been used in the past to
manufacture turbine engine components, such as turbine blades using
ceramic matrix composites. However, such techniques have resulted
in difficulties related to the small features of gas turbine engine
components for helicopter engines.
[0005] A number of techniques have been used in the past to
manufacture turbine engine components, such as turbine blades using
ceramic matrix composites. One method of manufacturing CMC
components, set forth in U.S. Pat. Nos. 5,015,540; 5,330,854; and
5,336,350; incorporated herein by reference in their entirety and
assigned to the assignee of the present invention, relates to the
production of silicon carbide matrix composites containing fibrous
material that is infiltrated with molten silicon, herein referred
to as the Silcomp process. The fibers generally have diameters of
about 140 micrometers or greater, which prevents intricate, complex
shapes having features on the order of about 0.030 inches, such as
turbine blade components for helicopter gas turbine engines, to be
manufactured by the Silcomp process.
[0006] Other techniques, such as the prepreg melt infiltration
process have also been used, however, the smallest cured
thicknesses with sufficient structural integrity for such
components have been in the range of about 0.033 inch to about
0.036 inch, since they are manufactured with standard prepreg
plies, which normally have an uncured thickness in the range of
about 0.009 inch to about 0.011 inch. With standard matrix
composition percentages in the final manufactured component, the
use of such uncured thicknesses results in final cured thicknesses
in the range of about 0.033 inch to about 0.036 inch.
[0007] What is needed is a method of manufacturing CMC helicopter
turbine engine components that permits the manufacture of features
having a thickness in the range of about 0.015 inch to about 0.021
inch. In addition, a method of manufacturing CMC helicopter turbine
engine components having features with a thickness less than about
0.021 inch is also needed.
SUMMARY OF THE INVENTION
[0008] An embodiment of the present invention is ceramic matrix
composite gas turbine engine component comprising a plurality of
cured ceramic matrix composite plies, each ply comprising ceramic
fiber tows, each ceramic fiber tow comprising a plurality of
ceramic fibers, the tows in each ply lying adjacent to one another
such that each ply has a unidirectional orientation. The component
further comprises a coating on the ceramic fibers, the coating
selected from the group consisting of boron nitride, silicon
nitride, silicon carbide, and combinations thereof. The component
further comprises a ceramic matrix material lying in interstitial
regions between the fibers and tows of each ply and the
interstitial region between the plurality of plies, wherein at
least a portion of the component comprising three cured ceramic
matrix composite plies, wherein each ceramic ply in at least the
portion has a thickness in the range of about 0.005 inch to about
0.007 inch, inclusive of tows, coating, and interstitial ceramic
matrix material, wherein the total thickness in at least the
portion is no greater than about 0.021 inch.
[0009] Another embodiment of the present invention is a method for
manufacturing an uncooled ceramic matrix composite gas turbine
engine component. The method comprises providing a plurality of
prepreg ceramic plies, the plies comprising prepreg ceramic fiber
tows, the tows in each ply lying adjacent to one another in a
planar arrangement such that each ply has a unidirectional
orientation. The method further comprises laying up the plurality
of prepreg ceramic plies in a preselected arrangement to form a
component shape. The method further comprises heating the component
shape to form a ceramic preform. The method further comprises
densifying the turbine blade preform with at least silicon to form
an uncooled ceramic matrix composite turbine blade, wherein at
least a portion of the component comprises cured ceramic plies
having a thickness in the range of about 0.005 inch to about 0.007
inch, inclusive of tows, coating, and interstitial ceramic matrix
material.
[0010] Yet another embodiment of the present invention is another
method for manufacturing a cooled ceramic matrix composite gas
turbine engine component. The method comprises providing a
plurality of prepreg ceramic plies, the plies comprising prepreg
ceramic fiber tows, the tows in each ply lying adjacent to one
another in a planar arrangement such that each ply has a
unidirectional orientation. The method further comprises laying up
the plurality of prepreg ceramic plies in a preselected arrangement
around tooling to form a component shape. The method further
comprises heating the component shape to form a ceramic preform.
The method further comprises removing the tooling. The method
further comprises densifying the preform with at least silicon to
form a ceramic matrix composite component wherein at least a
portion of the component comprises cured ceramic plies having a
thickness in the range of about 0.005 inch to about 0.007 inch,
inclusive of tows, coating, and interstitial ceramic matrix
material. The method further comprises drilling cooling holes in
the ceramic matrix composite component to form a cooled ceramic
matrix composite component.
[0011] An advantage of the present invention is that the use of a
lower volume percentage of interstitial ceramic matrix material
permits the manufacture of CMC articles having small features.
[0012] Another advantage of the present invention is that the use
of thin prepreg plies permits the manufacture of CMC articles
having small features.
[0013] Other features and advantages of the present invention will
be apparent from the following more detailed description of the
preferred embodiment, taken in conjunction with the accompanying
drawings which illustrate, by way of example, the principles of the
invention.
BRIEF DESCRIPTION OF THE DRAWINGS
[0014] FIG. 1 is an example of an uncooled CMC LPT blade for a gas
turbine engine.
[0015] FIG. 2 is a cross-section of the uncooled CMC LPT blade of
FIG. 1, taken at the line 2-2 illustrating one embodiment of the
present invention.
[0016] FIG. 3 is a cross-section of the uncooled CMC LPT blade of
FIG. 1, taken at the line 2-2 illustrating another embodiment of
the present invention.
[0017] FIG. 4 is an example of a cooled CMC LPT blade for a gas
turbine engine.
[0018] FIG. 5 is a cross-section of the uncooled CMC LPT blade of
FIG. 4, taken at the line 5-5 illustrating one embodiment of the
present invention.
[0019] FIG. 6 is a cross-section of the uncooled CMC LPT blade of
FIG. 4, taken at the line 5-5 illustrating another embodiment of
the present invention.
DETAILED DESCRIPTION OF THE INVENTION
[0020] FIG. 1 depicts an exemplary uncoated helicopter gas turbine
engine LPT uncooled blade 10. In this illustration a turbine blade
10 comprises a ceramic matrix composite material. The turbine blade
10 includes an airfoil 12 against which the flow of hot exhaust gas
is directed. The turbine blade 10 is mounted to a turbine disk (not
shown) by a dovetail 14 that extends downwardly from the airfoil 12
and engages a slot of complimentary geometry on the turbine disk.
The LPT blade 10 of the present invention does not include an
integral platform. A separate platform is provided to minimize the
exposure of the dovetail 14 to hot gases of combustion. The airfoil
has a leading edge section 16 and a trailing edge section 18. The
blade 10 is a relatively small component, and as shown in FIG. 1,
the trailing edge section has a thickness in the range of about
0.015 inch to about 0.021 inch.
[0021] A CMC turbine blade according to an embodiment is preferably
a SiC/SiC composite material manufactured using the "prepreg" MI
method. Exemplary processes for making such SiC/SiC prepreg
material are described in U.S. Pat. Nos. 6,024,898 and 6,258,737,
which are assigned to the Assignee of the present invention and
which are hereby incorporated herein by reference in their
entirety. Such "prepregged" plies comprise silicon
carbide-containing fibers, where the fibers are bundled into tows
and the tows are all adjacent to one another such that all of the
fibers are oriented in the same direction. By "silicon
carbide-containing fiber" is meant a fiber having a composition
that includes silicon carbide, and preferably is substantially
silicon carbide. For instance, the fiber may have a silicon carbide
core surrounded with carbon, or in the reverse, the fiber may have
a carbon core surrounded by or encapsulated with silicon carbide.
These examples are given for demonstration of the term "silicon
carbide-containing fiber" and are not limited to this specific
combination. Other fiber compositions are contemplated, so long as
they include silicon carbide.
[0022] In order to manufacture the turbine blade 10 of the present
invention, a preselected number of SiC prepregged plies are laid up
in a preselected arrangement forming a turbine blade shape, where
the trailing edge section 18 has a cured thickness in the range of
about 0.015 inch to about 0.021 inch. In order to have a trailing
edge with sufficient structural integrity to withstand operation in
a gas turbine engine, it is preferred to have at least two
interface regions between the plies (i.e. three plies) to provide
sufficient structural support.
[0023] These three plies that comprise trailing edge are preferably
oriented such that all three plies are not oriented in the same
direction. By "0.degree. orientation" with respect to a prepreg
ply, it is meant that a ply is laid up such that the line of the
fiber tows is in the line of the long dimension or axis of the
turbine blade as known in the art. A 90.degree. orientation means
that the ply is laid up such that the line of the fiber tows is
perpendicular to the long dimension or axis of the turbine blade as
known in the art. All orientations other than 0.degree. and
90.degree. may be negative or positive depending on whether the ply
is rotated clockwise (positive) from a preselected plane in the
long dimension of the turbine blade or rotated counterclockwise
(negative) from the preselected plane in the long dimension of the
turbine blade as known in the art. Prepreg plies that are oriented
at 0.degree. have tensile strength in the final CMC product that is
about twenty times greater than prepreg plies that are oriented at
90.degree.. The remaining plies that do not pass through the small
feature may be arranged in any appropriate orientation as known in
the art. For example, the remaining plies could all be laid up in
an alternating formation such that the remaining plies are at about
a 0.degree. orientation, followed by about a 90.degree.
orientation, followed by about a 0.degree. orientation, followed by
a -90.degree. orientation, etc. as is known in the art, or such
that the remaining plies are at a -45.degree. orientation, followed
by a 0.degree. orientation, followed by a +45.degree. orientation,
followed by a 90.degree. orientation, or in any other mechanically
acceptable arrangement
[0024] The next step of the process is forming a ceramic preform by
heating the turbine blade shape by compression molding, bladder
molding, or autoclaving as known in the art. The final step of the
process is densifying the preform with at least silicon, and
preferably with boron-doped silicon to form an uncooled CMC turbine
blade as known in the art. In one embodiment of carrying out the
infiltration, the preform is contacted with molten silicon, which
is infiltrated into the preform. U.S. Pat. No. 4,737,328, which is
assigned to the assignee of the present invention, and which is
hereby incorporated by reference in its entirety, discloses an
infiltration technique.
[0025] In order to achieve a small trailing edge 18 thickness in
the range of about 0.015 inch to about 0.021 inch, one approach is
to use plies that comprise smaller ceramic fiber tows, where each
tow has a denier in the range of about 800 to about 1000 and
wherein the thickness of each tow is in range of about 0.002 inch
to about 0.003 inch.
[0026] As shown in the exemplary cross-section of the trailing edge
section 18 in FIG. 2, the edge section 18 comprises three cured
prepreg plies 44. Each ply 44 comprises prepreg tows 46, the tows
46 comprising a plurality of fibers 72 coated with a coating 48
selected from the group consisting of boron nitride, silicon
nitride, silicon carbide and combinations thereof, as such coating
48 is known in the art, wherein the denier of each tow 46 is in the
range of about 800 to about 1000, such that each ply 44 has a
thickness in the range of about 0.005 inch to about 0.007 inch. An
interstitial ceramic material 50 is present between and surrounding
the plies 44, tows 46, and fibers 72, such that the total thickness
of the three cured plies 44 is in the range of about 0.015 inch to
about 0.021 inch. The interstitial ceramic material 50 also forms
the two interface regions 56 between the plies. Preferably, the
three plies do not have the same orientation. In the exemplary
embodiment shown in FIG. 2, two outer plies 54 are oriented at
0.degree., while the middle ply 56 is oriented at 90.degree..
[0027] A second approach is to use a relatively small amount of
ceramic matrix material, such that the volume percentage of the
matrix material in the blade 10 is in the range of about 60 percent
to about 70 percent. Increasing the volume percent of the matrix
may be accomplished by decreasing carbon content in the matrix
and/or decreasing particle content to ensure ability to fully fill
with silicon. Such a relatively low percentage of matrix material
means that in order to achieve a thickness in the range of about
0.015 inch to about 0.021 inch, prepreg plies having a thickness in
the range of about 0.005 inch to about 0.007 inch can be used.
[0028] As shown in the second exemplary cross-section of the
trailing edge section 18 in FIG. 3, the edge section 18 comprises
three cured prepreg plies 58. Each ply 58 comprises prepreg tows
60, the tows 60 comprising a plurality of fibers 74 coated with a
coating 62 selected from the group consisting of boron nitride,
silicon nitride, silicon carbide and combinations thereof, as such
coating 62 is known in the art, wherein the denier of each tow 60
is in the range of about 1400 to about 1800, such that each ply 58
has a thickness in the range of about 0.006 inch to about 0.007
inch. An interstitial ceramic material 64 is present between and
surrounding the plies 58, tows 60, and fibers 74, such that the
total thickness of the three cured plies 58 is in the range of
about 0.018 inch to about 0.021 inch. The interstitial ceramic
material 64 also forms the two interface regions 70 between the
plies 58. Preferably, the three plies 58 do not have the same
orientation. In the exemplary embodiment shown in FIG. 2, two outer
plies 66 are oriented at 0.degree., while the middle ply 68 is
oriented at 90.degree..
[0029] FIG. 4 depicts an exemplary helicopter gas turbine engine
LPT cooled blade 110. In this illustration a turbine blade 110
comprises a ceramic matrix composite material. The turbine blade
110 includes an airfoil 112 against which the flow of hot exhaust
gas is directed. The turbine blade 110 is mounted to a turbine disk
(not shown) by a dovetail 114 that extends downwardly from the
airfoil 112 and engages a slot of complimentary geometry on the
turbine disk. The LPT blade 110 of the present invention does not
include an integral platform. A separate platform is provided to
minimize the exposure of the dovetail 114 to hot gases of
combustion. The airfoil has a leading edge section 116 and a
trailing edge section 118. The blade 110 is a relatively small
component, and as shown in FIG. 1, the trailing edge section has a
thickness in the range of about 0.015 inch to about 0.021 inch. The
blade 110 further comprises cooling holes 120.
[0030] A CMC cooled turbine blade according to an embodiment is
preferably a SiC/SiC composite material manufactured using the
"prepreg" MI method. Such "prepregged" plies comprise silicon
carbide-containing fibers, where the fibers are bundled into tows
and the tows are all adjacent to one another such that all of the
fibers are oriented in the same direction. By "silicon
carbide-containing fiber" is meant a fiber having a composition
that includes silicon carbide, and preferably is substantially
silicon carbide. For instance, the fiber may have a silicon carbide
core surrounded with carbon, or in the reverse, the fiber may have
a carbon core surrounded by or encapsulated with silicon carbide.
These examples are given for demonstration of the term "silicon
carbide-containing fiber" and are not limited to this specific
combination. Other fiber compositions are contemplated, so long as
they include silicon carbide.
[0031] In order to manufacture the cooled turbine blade 110 of the
present invention, a preselected number of SiC prepregged plies are
laid up in a preselected arrangement over a piece of tooling in a
preselected shape forming a turbine blade shape, where the trailing
edge section 18 has a cured thickness in the range of about 0.015
inch to about 0.021 inch. In order to have a trailing edge with
sufficient structural integrity to withstand operation in a gas
turbine engine, it is preferred to have at least two interface
regions between the plies (i.e. three plies) to provide sufficient
structural support.
[0032] These three plies that comprise trailing edge are preferably
oriented such that all three plies are not oriented in the same
direction. The remaining plies that do not pass through the small
feature the trailing edge section 118 may be arranged in any
appropriate orientation as known in the art. For example, the
remaining plies could all be laid up in an alternating formation
such that the remaining plies are at about a 0.degree. orientation,
followed by about a 90.degree. orientation, followed by about a
0.degree. orientation, followed by a -90.degree. orientation, etc.
as is known in the art, or such that the remaining plies are at a
-45.degree. orientation, followed by a 0.degree. orientation,
followed by a +45.degree. orientation, followed by a 90.degree.
orientation, or in any other mechanically acceptable
arrangement
[0033] The next step of the process is forming a ceramic preform by
heating the turbine blade shape by compression molding, bladder
molding, or autoclaving as known in the art. The next step is
removing the tooling by a physical, chemical, or thermal process.
For example, the tooling may be machined out of the interior, or
dissolved, etched away, or vaporized, depending on its material of
construction. The next step of the process is densifying the
preform with at least silicon, and preferably with boron-doped
silicon to form a CMC turbine blade as known in the art. In one
embodiment of carrying out the infiltration, the preform is
contacted with molten silicon, which is infiltrated into the
preform. U.S. Pat. No. 4,737,328, which is assigned to the assignee
of the present invention, and which is hereby incorporated by
reference in its entirety, discloses an infiltration technique. The
final step is drilling cooling holes 120 by any means known in the
art, for example by a Nd:YAG laser. Such use of an Nd:YAG laser is
discussed in U.S. Pat. No. 6,670,026, which is assigned to the
Assignee of the present invention, and which is hereby incorporated
by reference in its entirety.
[0034] As shown in the exemplary cross-section of the trailing edge
section 118 in FIG. 5, the edge section 118 comprises three cured
prepreg plies 144. Each ply comprises prepreg tows 146, the tows
146 comprising a plurality of fibers 172 coated with a coating 148
selected from the group consisting of boron nitride, silicon
carbide and combinations thereof, as such coating 148 is known in
the art, wherein the denier of each tow 146 is in the range of
about 800 to about 1000, such that each tow 146 has a thickness in
the range of about 0.005 inch to about 0.007 inch. An interstitial
ceramic material 150 is present between and surrounding the plies
144, tows 146, and fibers 172, such that the total thickness of the
three cured plies 144 is in the range of about 0.015 inch to about
0.021 inch. The interstitial ceramic material 150 also forms the
two interface regions 156 between the plies. Preferably, the three
plies 144 do not have the same orientation. In the exemplary
embodiment shown in FIG. 5, two outer plies 154 are oriented at
0.degree., while the middle ply 154 is oriented at 90.degree..
[0035] A second approach is to use a relatively small amount of
ceramic matrix material, such that the volume percentage of the
matrix material in the blade 110 is in the range of about 60
percent to about 70 percent. Increasing the volume percent of the
matrix may be accomplished by decreasing carbon content in the
matrix and/or decreasing particle content to ensure ability to
fully fill with silicon. Such a relatively low percentage of matrix
material means that in order to achieve a thickness in the range of
about 0.015 inch to about 0.020 inch, prepreg plies having a
thickness in the range of about 0.005 inch to about 0.008 inch can
be used.
[0036] As shown in the second exemplary cross-section of the
trailing edge section 118 in FIG. 6, the edge section 118 comprises
three cured prepreg plies 158. Each ply comprises prepreg tows 160
comprising a plurality of fibers 174 coated with a coating 162
selected from the group consisting of boron nitride, silicon
carbide and combinations thereof, as such coating 162 is known in
the art, wherein the denier of each tow 160 is in the range of
about 1400 to about 1800, such that each tow 160 has a thickness in
the range of about 0.006 inch to about 0.007 inch. An interstitial
ceramic material 164 is present between and surrounding the plies
158, tows 160, and fibers 174, such that the total thickness of the
three cured plies 158 is in the range of about 0.018 inch to about
0.021 inch. The interstitial ceramic material 164 also forms the
two interface regions 170 between the plies 160. Preferably, the
three plies 160 do not have the same orientation. In the exemplary
embodiment shown in FIG. 2, two outer plies 166 are oriented at
0.degree., while the middle ply 168 is oriented at 90.degree..
[0037] While exemplary embodiments of the invention have been
described with respect to a cooled and uncooled helicopter gas
turbine engine LPT blade, it should be appreciated that the
invention is not so limited and that the present invention also In
addition, the prepreg plies and methods of the present invention
may also be used to manufacture any gas turbine engine component
having small features, including any ceramic matrix composite
turbine engine component, such as any turbine blade, and uncooled
turbine nozzle, a cooled turbine nozzle, wherein the component is
initially laid up in a preselected arrangement using a preselected
number of prepreg CMC plies to manufacture an article having small
features requiring the use of cured plies having a thickness in the
range of about 0.005 inch to about 0.008 inch.
[0038] While the invention has been described with reference to a
preferred embodiment, it will be understood by those skilled in the
art that various changes may be made and equivalents may be
substituted for elements thereof without departing from the scope
of the invention. In addition, many modifications may be made to
adapt a particular situation or material to the teachings of the
invention without departing from the essential scope thereof.
Therefore, it is intended that the invention not be limited to the
particular embodiment disclosed as the best mode contemplated for
carrying out this invention, but that the invention will include
all embodiments falling within the scope of the appended
claims.
* * * * *