U.S. patent application number 10/758250 was filed with the patent office on 2005-07-21 for hybrid ceramic matrix composite turbine blades for improved processibility and performance.
This patent application is currently assigned to General Electric Company. Invention is credited to Carper, Douglas Melton, Jendrix, Richard William, Steibel, James Dale, Subramanian, Suresh.
Application Number | 20050158171 10/758250 |
Document ID | / |
Family ID | 34620692 |
Filed Date | 2005-07-21 |
United States Patent
Application |
20050158171 |
Kind Code |
A1 |
Carper, Douglas Melton ; et
al. |
July 21, 2005 |
Hybrid ceramic matrix composite turbine blades for improved
processibility and performance
Abstract
The present invention is a hybrid ceramic matrix composite
turbine engine component comprising an outer shell section(s) and
an inner core section(s), wherein the outer shell section(s) and
the inner core section(s) were bonded together using an MI process.
The outer shell section(s) comprises a SiC/SiC material that has
been manufactured using a process selected from the group
consisting of a slurry cast MI process and a prepreg MI process.
The inner core section(s) comprises a material selected from the
group consisting an Si/SiC composite material and a monolithic
ceramic material. The Si/SiC composite material may be manufactured
using the Silcomp process. The present invention may be a high
pressure turbine blade, a high pressure turbine vane, a low
pressure turbine blade, or a low pressure turbine vane. The present
invention is also a method of manufacturing a hybrid ceramic matrix
composite turbine engine component.
Inventors: |
Carper, Douglas Melton;
(Trenton, OH) ; Subramanian, Suresh; (Mason,
OH) ; Jendrix, Richard William; (Liberty Township,
OH) ; Steibel, James Dale; (Hamilton, 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: |
34620692 |
Appl. No.: |
10/758250 |
Filed: |
January 15, 2004 |
Current U.S.
Class: |
415/200 |
Current CPC
Class: |
F05D 2300/603 20130101;
C04B 2235/5244 20130101; F01D 5/284 20130101; C04B 35/62863
20130101; C04B 35/565 20130101; B32B 2315/02 20130101; C04B
2237/361 20130101; C04B 2237/61 20130101; Y02T 50/672 20130101;
B32B 18/00 20130101; C04B 35/62868 20130101; C04B 2235/5256
20130101; B32B 2311/22 20130101; Y02T 50/67 20130101; C04B 2237/368
20130101; C04B 2237/38 20130101; F01D 5/282 20130101; F05D
2300/2261 20130101; C04B 2235/616 20130101; C04B 35/62873 20130101;
C04B 2235/80 20130101; C04B 2235/5248 20130101; C04B 2237/363
20130101; C04B 2235/428 20130101; C04B 35/62871 20130101; C04B
2235/5268 20130101; C04B 2237/365 20130101; C04B 2235/422 20130101;
Y02T 50/60 20130101 |
Class at
Publication: |
415/200 |
International
Class: |
F01D 001/02 |
Claims
What is claimed is:
1. A ceramic matrix composite turbine engine component comprising:
a core insert section comprising a material selected from the group
consisting of a silicon-silicon carbide composite and a monolithic
ceramic; and an outer shell section comprising a silicon
carbide-silicon carbide composite material, wherein the at least
one outer shell section and the at least one core insert section
are bonded together using a silicon melt infiltration process.
2. The turbine engine component of claim 1, wherein the component
comprises a plurality of outer shell sections.
3. The turbine engine component of claim 1, wherein the component
comprises a plurality of core insert sections and a plurality of
outer shell sections.
4. The turbine engine component of claim 1, wherein the core insert
section is a silicon-silicon carbide composite.
5. The turbine engine component of claim 4, wherein the core insert
section is manufactured using a Silcomp process.
6. The turbine engine component of claim 1, wherein the core insert
section is a monolithic ceramic material.
7. The turbine engine component of 2, wherein the outer shell
section is silicon carbide-silicon carbide composite material.
8. The turbine engine component of claim 7, wherein the outer shell
section is manufactured using a slurry cast melt infiltration
process.
9. The turbine engine component of claim 7, wherein the outer shell
section is manufactured using a prepreg melt infiltration
process.
10. The turbine engine component of claim 1, wherein the component
is a turbine vane.
11. The turbine engine component of claim 1, wherein the component
is a turbine nozzle.
12. A method of manufacturing a ceramic matrix composite turbine
blade comprising the steps of: providing a core insert section
having a preselected geometry, the core insert section comprising a
material selected from the group consisting of silicon
carbide-silicon carbide composite preform having at least some
porosity, silicon-silicon carbide composite, silicon-silicon
carbide composite preform having at least some porosity,
silicon-silicon carbide composite, and a monolithic ceramic;
providing a plurality of plies of silicon carbide prepreg cloth;
laying up a preselected number of silicon carbide prepreg plies to
form an outer shell section; assembling the core insert section and
the outer shell section into a turbine blade form, the turbine
blade form comprising a dovetail section and an airfoil section,
wherein the core insert section is positioned in the dovetail
section of the turbine blade form; autoclaving the turbine blade
form; filling remaining porosity in the turbine blade form with at
least silicon using a silicon melt infiltration process, the
filling also forming a bond between the core insert section and the
outer shell preform.
13. The method of claim 12, wherein the core insert section is a
silicon-silicon carbide preform.
14. The method of claim 13, wherein the silicon-silicon carbide
preform includes carbon microspheres.
15. The method of claim 12, wherein the core insert section is a
silicon carbide-silicon carbide preform manufactured using a slurry
cast process.
16. The method of claim 12, wherein the core insert section is a
silicon carbide-silicon carbide preform manufactured using a
prepreg process.
17. A method of manufacturing a ceramic matrix composite turbine
blade comprising the steps of: providing a core insert section
having a preselected geometry, the core insert section comprising a
material selected from the group consisting of a silicon
carbide-silicon carbide composite preform having at least some
porosity, a silicon-silicon carbide composite, the silicon-silicon
carbide composite preform having at least some porosity, a
silicon-silicon carbide composite, and a monolithic ceramic;
providing an outer shell section preform, the outer shell preform
having at least some porosity; assembling the core insert section
and the outer shell preform into a turbine blade form, the turbine
blade form comprising a dovetail section and an airfoil section,
wherein the core insert section is positioned in the dovetail
section of the turbine blade form; and filling remaining porosity
in the turbine blade forms with at least silicon using the silicon
melt infiltration process, the filling also forming a bond between
the at least one core insert section and the at least one outer
shell preform.
18. The method of claim 17, wherein the core insert section is a
silicon-silicon carbide preform.
19. The method of claim 18, wherein the silicon-silicon carbide
preform includes carbon micro spheres.
20. The method of claim 19, wherein the core insert section is a
silicon carbide-silicon carbide preform manufactured using a slurry
cast process.
Description
FIELD OF THE INVENTION
[0001] The present invention relates generally to ceramic matrix
composite turbine engine components, and more particularly, to a
hybrid ceramic matrix composite turbine blade.
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 turbine 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.
However, their material properties, for example, fracture
resistance, particularly at low temperatures, has posed an
impediment to more widespread use in the engine.
[0004] A number of techniques have been used in the past to
manufacture thick dovetail sections of turbine engine components
using both polymeric and ceramic matrix composites. These
techniques include the use of insert plies and preform inserts in
the dovetail section to build up thicknesses. The local properties
of the composite are typically optimized by selecting the
appropriate ply stacking sequence. One of the main challenges in
manufacturing turbine blades using composite materials is the
building up of the thick dovetail section of such turbine blades.
Due to the relatively brittle nature of the ceramic matrix
composites (CMCs), particularly at low temperatures, and the unique
processing step used to make these materials such as melt
infiltration and chemical vapor infiltration, the concepts that
have been used in the past for polymeric matrix composites (PMCs)
are not directly useful here, since such concepts do not fully
densify thicker portions of CMCs, which are required for turbine
blade dovetail sections. High compressive strength is required in
the centermost regions of the dovetail section of turbine blades.
Such high compressive strength is not achieved when the thick
dovetail section of CMC turbine blades is not fully densified. In
addition, the material density in the dovetail section of CMC
turbine blades is not predictable because of the level of
densification that does occur is unpredictable. What is needed is a
method that fully densifies thick dovetail sections of CMC turbine
blades in a predictable manner to produce a CMC turbine blade
having high temperature capability.
SUMMARY OF THE INVENTION
[0005] Improvements in manufacturing technology and materials are
the keys to increased performance and reduced costs for many
articles. As an example, continuing and often interrelated
improvements in processes and materials have resulted in major
increases in the performance of aircraft gas turbine engines, such
as the improvements of the present invention. A novel method for
manufacturing a turbine blade made from a ceramic matrix composite
(CMC) produces a dovetail section having increased strength and
toughness thereby improving the functionality of the component. The
high strength of the dovetail section of the turbine blades
produced by the method of the present invention overcomes the
problem of weak dovetail sections caused by failure to attain full
densification within the dovetail sections of previously
manufactured CMC turbine blades.
[0006] The turbine section of the engine is aft of the combustor
section and converts the hot gases of combustion from the
combustion of the fuel into mechanical energy. The turbine section
includes a high pressure turbine section (HPT), positioned
immediately aft of the combustor section, and a low pressure
turbine section (LPT) positioned immediately behind the HPT.
Combustion gases moving past the HPT are very energetic, and the
HPT blades are small. As the gases expand and move rearward, they
become less energetic. The LPT blades, in order to extract the
maximum amount of energy from the less energetic fluid, are longer
and extend further outward into the expanding annuli of the turbine
section. Although generally applicable to all turbine blades, the
present invention is particularly directed to the longer, larger
LPT blades. The present invention enables manufacture of larger
blades having even thicker dovetails that otherwise could not be
fully densified to acceptable levels of compressive strength. In
addition, the present invention provides a predictable material
density, which is not possible with prior art methods of
manufacture.
[0007] Turbine blades generally include an airfoil against which
the flow of hot exhaust gas is directed and are mounted to turbine
disks by a dovetail that extends downwardly from the airfoil and
engages a slot on the turbine disk. A platform extends
longitudinally outward from the area where the airfoil is joined to
the dovetail being substantially perpendicular to the airfoil
extending in a first direction away from the platform and the
dovetail extending in a second direction away from the platform and
opposite the first direction. The airfoil may be described as
having a root end adjacent to the platform, and an oppositely
disposed tip end remote from the platform that extends into the hot
air stream from the combustor section of the engine. The dovetail
may be described as having a first end adjacent to the platform and
an oppositely disposed tip end remote from the dovetail.
[0008] The dovetail, named for its shape, is designed to allow
assembly into a correspondingly shaped slot in the turbine disk.
The dovetail experiences very high shear stresses, as it must
counteract the rotational forces that would otherwise eject the
blade from the disk. These forces increase as the weight of the
blade increases. Thus, these forces are higher for LPT blades than
for HPT blades. In order to achieve the required thickness build up
in the dovetail section of CMC turbine blades, a number of
techniques can be used. In the present invention, a novel technique
includes manufacturing the turbine blade from a pre-fabricated core
insert section(s) and an outer shell section(s). The core insert
section(s) is made from either the same material as the outer shell
section(s) or monolithic ceramic material. The core insert
section(s) will have a cross-section varying over the length of the
blade and is manufactured separately from the outer shell
section(s), forming at least a portion of the dovetail section. The
pre-fabricated core insert section(s) may form an inner portion of
the turbine airfoil. The outer shell section(s) forms at least an
outer portion of the turbine blade and at least an outer portion of
the dovetail section. The thickness of the outer shell section(s)
is selected based on the shear stress distribution of the part and
may be in the thickness range that can be densified with silicon
through melt infiltration easily without introducing significant
porosity. The thickness of the outer shell section(s) is
preselected based on the known shear stress distribution in the
specific turbine blade. The interface between the outer shell
section(s) and the core insert section(s) is selected so as to be
located in a low shear stress region of the turbine blade.
[0009] Without wishing to be bound by theory, it is believed that a
more complete densification can be obtained using the melt
infiltration densification process than with other densification
processes. Melt infiltration processes are the preferred method of
manufacturing CMCs. Melt infiltration processes produce CMCs with
the high thermal conductivity, as a result of substantially
complete densification relative to other processes for a given
configuration, that is required for CMCs to handle thermal
stresses. In addition, melt infiltration provides high matrix
cracking strength.
[0010] The properties of the core insert section(s) located in the
low shear stress region typically differ from the outer shell
section(s), which is located in high shear stress regions. The core
insert section(s) may be made stiffer than the outer shell
section(s) by manufacturing the core insert section(s) with less
residual silicon content. CMCs produced using melt infiltration
techniques typically contain silicon volume percent in the range of
about 20 volume percent to about 30 volume percent of silicon in
the SiC matrix. The core insert section(s) of the present invention
can initially be made with less than about 5 volume percent silicon
if a SilComp process is used and can be made with 0 volume percent
if a sintered SiC core insert section(s) is used. Such a reduction
would help to improve the vibration performance of the turbine
blade as the change in stiffness from the outer shell to the core
insert will reduce vibrations as the less rigid core insert
section(s) absorbs the vibrational energy.
[0011] The density of the core insert section(s) may also be
tailored by introducing microspheres comprising materials selected
from the group consisting of carbon, glass, ceramic, and
combinations thereof, into the core insert section(s) prior to
silicon melt infiltration. Such use of carbon microspheres will
change the modulus, while reducing the density of the core insert
section(s) with as little reduction in stiffness as possible. The
range of the carbon microsphere in the initial core insert
section(s) is in the range of about 0 volume percent to about 50
volume percent.
[0012] Once the core insert section(s) is manufactured, the core
insert section(s) and the outer shell section(s) are assembled, the
core insert section(s) having been laid up in a preferred stacking
sequence scheme using a slurry cast or a prepreg method. The outer
shell sections and the inner core section(s) are assembled into a
turbine blade form on a lay-up tool and cured in an autoclave in a
conventional manner. The cured blade is densified using a silicon
melt infiltration process that can fill the remaining porosity in
the outer shell sections and possibly the inner core section(s),
depending on the process used to manufacture the inner core
section(s). The use of the silicon melt infiltration process yields
a good bond at the interface between the inner core section(s) and
outer shell section(s). This is particularly true when SilComp
material is used for the core insert section(s), because residual
silicon in the SilComp material softens during the melt
infiltration of the outer shell section forming an integral bond
with the outer shell through the mechanism of diffusion bonding.
The dimensions of the outer shell section(s) and the inner core
section(s) can be selected such that the high stress regions uses a
high tensile strength and tough CMC and the inner core insert
section(s) uses a high compressive strength material. Also, the
skin core interface can be located in a low shear stress region of
the part. The present invention is also a turbine engine component
comprising at least one core insert section comprising a material
selected from the group consisting of a silicon-silicon carbide
composite, a silicon carbide-silicon carbide composite, and a
monolithic ceramic; and at least one outer shell section comprising
a silicon carbide-silicon carbide composite material, wherein the
at least one outer shell section and the at least one core insert
section were bonded together using a silicon melt infiltration
process.
[0013] The present invention is also a turbine engine component
comprising a core insert section comprising a material selected
from the group consisting of a silicon-silicon carbide composite
and a monolithic ceramic and an outer shell section comprising a
silicon carbide-silicon carbide composite material, wherein the
outer shell section and the core insert section were bonded
together using a silicon melt infiltration process.
[0014] The present invention is also a method of manufacturing a
ceramic matrix composite turbine component. The method of the
present invention includes providing at least one core insert
section having a preselected geometry, the at least one core insert
section comprising a material selected from the group consisting of
silicon carbide-silicon carbide composite preform having at least
some porosity, silicon-silicon carbide composite, silicon-silicon
carbide composite preform having at least some porosity,
silicon-silicon carbide composite, and a monolithic ceramic. The
process also includes providing a plurality of plies of silicon
carbide prepreg cloth and laying up a preselected number of silicon
carbide prepreg plies to form at least one outer shell section. At
least one at least one core insert section is assembled with the at
least one outer shell section into a turbine blade form, wherein
the turbine blade form comprises a dovetail section and an airfoil
section, and wherein the at least one core insert section is
positioned in the dovetail section of the turbine blade form. The
turbine blade form is then autoclaved. Remaining porosity in the
turbine blade form is filled with at least silicon using a silicon
melt infiltration process. The filler material also forms a bond
between the at least one core insert section and the at least one
outer shell preform.
[0015] The present is also another method of manufacturing a
ceramic matrix composite turbine blade component. The method of the
present invention includes providing a core insert section having a
preselected geometry, the core insert section comprising a material
selected from the group consisting of silicon carbide-silicon
carbide composite preform having at least some porosity,
silicon-silicon carbide composite, silicon-silicon carbide
composite preform having at least some porosity, silicon-silicon
carbide composite, and a monolithic ceramic. An outer shell section
preform having at least some porosity is provided. The at least one
core insert section and the at least one outer shell preform are
laid up into a turbine blade form, the turbine blade form
comprising a dovetail section and an airfoil section, wherein the
at least one core insert section is positioned in the dovetail
section of the turbine blade form. Remaining porosity in the
turbine blade form is filled with at least silicon using a silicon
melt infiltration process. The filler material also forms a bond
between the core insert section and the outer shell preform.
[0016] An advantage of the present invention is that a thick CMC
dovetail section for a turbine blade may be readily built up in
which the properties of portions of the dovetail are tailored to
meet the stresses to which they will be subjected.
[0017] Another advantage of the present invention is that it
permits the fabrication of complex composite articles that are
difficult or impossible to fabricate by conventional fabrication
technology in which properties are substantially isotropic through
the article.
[0018] Another advantage of the present invention is that different
materials having different physical properties may be used within
the dovetail section of the turbine blade to modify the physical
properties of an otherwise unitary blade.
[0019] Another advantage of the present invention is that the
ability of the composite to alternate vibrations improves its crack
arresting capabilities as well as its ability to resist crack
formation, thereby improving back fracture toughness and fatigue
over the design life of the article.
[0020] Yet another advantage of the present invention is the
ability to manufacturing larger turbine blades having thicker
dovetails.
[0021] 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
[0022] FIG. 1 is an exemplary example of a LPT blade in an aircraft
engine.
[0023] FIG. 2 is a flow chart illustrating a method of manufacture
of a CMC turbine blade of the present invention using a prepreg MI
process.
[0024] FIG. 3 a flow chart illustrating an alternative method of
manufacture of a CMC turbine blade of the present invention using a
slurry cast MI process.
[0025] FIG. 4 is a cross-sectional view of an embodiment of a CMC
turbine blade dovetail of the present invention.
[0026] FIG. 5 is a cross-sectional view of an embodiment of a CMC
turbine blade dovetail of the present invention.
DETAILED DESCRIPTION OF THE INVENTION
[0027] FIG. 1 depicts an exemplary aircraft engine LPT blade 20. In
this illustration a turbine blade 20 comprises a ceramic matrix
composite material. The turbine blade 20 includes an airfoil 22
against which the flow of hot exhaust gas is directed. The turbine
blade 20 is mounted to a turbine disk (not shown) by a dovetail 24
that extends downwardly from the airfoil 22 and engages a slot of
similar geometry on the turbine disk. The LPT blade of the present
invnetion does not have an integral platform. A separate platform
is provided to minimize the exposure of the dovetail to hot gases
of combustion. The airfoil may be described as having a root end 40
and an oppositely disposed tip end 32.
[0028] FIG. 2 is a flow chart illustrating a method of manufacture
of a CMC turbine blade of the present invention using a method of
silicon carbide-silicon carbide CMC manufacture known as a
"prepreg" melt infiltration process (prepreg MI). Initially, outer
shell SiC preforms are manufactured using plies of "prepregged"
ceramic cloth. Such "prepregged" plies typically comprise
silicon-carbide containing fibers wound onto a drum in a
uni-directional manner to form a sheet. The fibers were previously
coated with a ceramic layer such as boron nitride (BN) and silicon
nitride (Si.sub.3N.sub.4) and that has been previously infiltrated
with high char materials. 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. 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 incorporated herein by reference.
[0029] The first step of the process 100 is the manufacture of a
core insert section. The number of such core insert sections that
are assembled into a blade and their geometry will depend upon the
number of outer shell sections preselected for the turbine blade.
For example, as shown in FIG. 4, if two outer shell sections 36 are
used for the turbine blade, then one core insert section 34 is
required to go between the two outer shell sections. In an
alternate embodiment, as shown in FIG. 5, where five outer shell
sections 34 are manufactured, a total of five core insert sections
36 are required to build up the dovetail section 24 of the turbine
blade.
[0030] Both the core insert section(s) and outer shell section(s)
of the CMC turbine blade may be manufactured using the prepreg MI
process. Alternatively the core insert section may comprise
silicon-silicon carbide (Si/SiC) composite materials, made by a
process known as "SilComp" may be used as the monolithic core
insert section(s) of the present invention in conjunction with
prepreg MI outer shell section(s). A technical description of such
an Si/SiC composite material is described in detail in U.S. Pat.
No. 5,015,540, U.S. Pat. No. 5,330,854, and U.S. Pat. No.
5,336,350, which are all assigned to the Assignee of the present
invention and which are incorporated by reference herein. The core
insert section(s) may also be any other monolithic ceramic as known
in the art. The size and composition of the core insert section(s)
may be selectively tailored to provide the appropriate stiffness
and density through such techniques as the addition of microspheres
during manufacture of the core insert section(s), which causes the
formation of porosity within the matrix. Additionally, the
stiffness of the core insert section(s) could be made higher than
the outer shell by reducing the residual silicon content in the
core insert section(s) if desired. The SilComp core insert
section(s) may or may not be densified through MI prior to the step
120 of positioning the core insert section(s) into a preselected
position with respect to the laid-up outer shell section(s).
[0031] When the core insert section(s) and outer shell section(s)
of the CMC turbine blades are both manufactured using the prepreg
process, the manufacture of the core insert section(s) of the
dovetail section of the turbine blade is accomplished by laying-up
a preselected number of plies in a preselected orientation scheme
on a lay-up tool to provide the appropriate mechanical properties
for the turbine blade. Prepreg plies are comprised of bundled fiber
tows that are positioned adjacent to one another such that all of
the tows are aligned in the same direction. In terms of prepreg ply
orientation, by "0.degree. orientation," it is meant that the
prepreg plies are laid up so that the tows are oriented in a
preselected long dimension plane of the turbine blade lying
substantially in the radial direction of the engine as known in the
art. By "90.degree. orientation," it is means that the plies are
laid up such that the orientations of the tows are perpendicular to
the preselected long dimension plane. All orientations of then
0.degree. or 90.degree. may be negative or positive depending on
whether the ply is rotated clockwise (positive) from the
preselected long dimension plane or rotated counterclockwise
(negative) from the preselected long dimension plane as known in
the art. In one embodiment a core insert section is laid up with
2-D prepreg plies cut to form the geometry of the core insert
section when stacked. The 2-D prepreg plies are laid up in an
alternative formation, such that the plies are at a 45.degree.
orientation, followed by a -45.degree. orientation, followed by a
45.degree. orientation, followed by a -45.degree. orientation, etc.
as known in the art. Such an orientation would maximize tensional
rigidity of the core insert section. To complete the manufacture of
the core insert section(s), the uncured core insert section(s) is
then cured in an autoclave, as known in the art, and machined to
the required geometry for the core insert section. The precursor
material is referred to as "green" since it is still relatively
flexible, as compared to the final ceramic matrix composite
material, which is rather brittle.
[0032] The next step 110, which may occur in parallel with the
prior steps for the core insert section(s), is laying up a
preselected number of prepreg plies of silicon carbide cloth, which
will form the outer shell section(s), in a preselected orientation
scheme in a preselected stacking sequence. Alternatively, the order
of steps 100 and 110 may be reversed. The next step 120 is
positioning the core insert section(s) into a preselected position
with respect to the laid-up outer shell section(s). The core insert
section(s) is positioned in a manner to build up the dovetail
section(s) of the turbine blade into the shape of the final
dovetail section(s) of the turbine blade, producing a turbine blade
form. The next step 130 is autoclaving the turbine blade form, as
known in the art, to cure the uncured the outer shell section(s).
The next step 140 densities the turbine blade form using a silicon
melt infiltration process as known in the art, forming the CMC
turbine blade of the present invention. Such densification will
create a strong bond at the interface between the outer shell
section(s) and the core insert section(s). If the prepreg MI
process was used to form the core insert section(s), the step 140
of densifying will further densify the core insert section(s). If
the SilComp process was used to form the core insert section(s),
and the core insert section(s) was not previously densified through
MI, the step 140 of densifying also densify the outer shell
section(s). If the SilComp process was used to form the core insert
section(s), and the core insert section(s) was previously densified
through MI, the step 140 of densifying will simply form a bond
between the silicon of the core insert section(s) and the outer
shell section(s). The resulting article comprises a unitary
composite material structure with at least two separate sections
with at least two of the sections having different materials
properties. The article may be sized to be thicker or thinner as
appropriate as known in the art.
[0033] FIG. 3 is a flow chart illustrating a method of manufacture
of a CMC turbine blade of the present invention using a method of
silicon carbide-silicon carbide CMC manufacture known as a slurry
cast melt infiltration process (slurry cast MI). Both the core
insert section(s) and outer shell section(s) of the CMC turbine
blade may be manufactured using the slurry cast MI process. A
technical description of such a slurry cast MI method is described
in detail in U.S. Pat. No. 6,280,550 B1, which is assigned to the
Assignee of the present invention and which is incorporated herein
by reference. The core insert section(s) may also be manufactured
with the SilComp process.
[0034] The optional first step of the process 200 is the
manufacture of a core insert section(s) preform using the slurry
cast process, wherein some porosity remains in the core insert
section(s). In a preferred embodiment, the porosity of the core
insert section(s) will be in the range of about 20 volume percent
to about 40 volume percent. The number of such core insert sections
and their preselected geometry will depend upon the number of
layers of outer shell sections preselected for the turbine blade
and the mechanical properties required of the turbine blade. For
example, as shown in FIG. 4, if two outer shell sections 36 are
used for the turbine blade, and then one core insert section 34 is
required to go between the two outer shell sections 36. In an
alternate embodiment, as shown in FIG. 5, where five outer shell
sections 36 are manufactured, a total of five core insert sections
34 are required to build up the dovetail section 24 of the turbine
blade. Both the core insert section(s) and outer shell section(s)
of the CMC turbine blade may be manufactured using the slurry cast
MI process. The second optional first step of the process 210 is
the manufacture of a SilComp Si/SiC core insert section(s).
Alternatively the core insert section(s) may comprise
silicon-silicon carbide (Si/SiC) composite materials, made by a
process known as "SilComp" may be used as the monolithic core
insert section(s) of the present invention in conjunction with
prepreg MI outer shell section(s). The core insert section(s) may
also be any other monolithic ceramic as known in the art. The size
and composition of the core insert section(s) may be selectively
tailored to provide the appropriate stiffness and density through
such techniques as the addition of carbon microspheres during
manufacture of the core insert section(s), which causes the
formation of porosity within the matrix. Additionally, the
stiffness of the core insert section(s) could be made higher than
the shell by reducing the residual silicon content in the core
insert section(s).
[0035] The next step 220 is the manufacture of an outer shell
preform(s) using the slurry cast method, such a slurry cast method
leaving the preform(s) with some porosity remaining. Alternatively,
the order of steps 200 and 220 or 200 and 210 may be reversed or
performed at the same time. The next step 230 is the positioning of
the outer shell preform(s) and the core insert section(s) to form a
turbine blade preform. The core insert section(s) is positioned
within at least the dovetail section of the turbine blade form. The
next step 240 is densifying the turbine blade form using a silicon
melt infiltration process as known in the art, forming the CMC
turbine blade of the present invention. Such densification will
create a strong bond between the outer shell section(s) and the
core insert section(s). If the slurry cast MI process was used for
the core insert section(s), the step 240 of densifying will also
densify the core insert section(s). If the SilComp process was used
for the core insert section(s), and the core insert section(s) was
not previously densified through MI, the step 140 of densifying
will densify the core insert section(s) and the outer shell
sections. If the SilComp process was used for the core insert
section(s), and the core insert section(s) was previously densified
through MI, the step 140 of densifying will simply form a bond
between the silicon of the core insert section(s) and the outer
shell sections. The resulting article comprises a composite
material structure with at least two separate sections having
different material properties. The article may be sized to be
thicker or thinner as appropriate as known in the art.
[0036] In an additional, alternative embodiment, the outer shell
section(s) are manufactured using a slurry cast MI process, while
the core insert section(s) are manufactured using a prepreg MI
process. In another additional alternative embodiment, the outer
shell section(s) are manufactured using a prepreg MI process, while
the core insert section(s) are manufactured using a slurry cast MI
process.
[0037] The process of the present invention may be used to
manufacture any turbine engine airfoil, including, but not limited
to high pressure turbine blades, high pressure turbine vanes, low
pressure turbine blades, and low pressure turbine vanes.
[0038] FIG. 4 is a cross-sectional view of a turbine blade dovetail
24 of an embodiment of a CMC turbine blade 20 of the present
invention. The turbine blade 20 includes an airfoil (not shown)
against which the flow of hot exhaust gas is directed. The turbine
blade 20 is mounted to a turbine disk (not shown) by a dovetail 24
that extends downwardly from the airfoil and engages a slot on the
turbine disk where it is secured in position. In the embodiment
shown in FIG. 4, there is one silicon-silicon carbide ceramic
matrix composite core insert section 34 that is manufactured using
a SilComp method. Alternatively, the core insert section may be a
monolithic ceramic material. There are also two outer silicon
carbide-silicon carbide ceramic matrix composite shell sections 36
that have been manufactured using a slurry cast MI method. The bond
between the core insert section 34 and the outer shell sections 36
is created during the final processing step of silicon melt
infiltration.
[0039] FIG. 5 is a cross-sectional view of a turbine blade dovetail
24 of an embodiment of a CMC turbine blade 20 of the present
invention. The turbine blade 20 includes an airfoil (not shown)
against which the flow of hot exhaust gas is directed. The turbine
blade 20 is mounted to a turbine disk (not shown) by a dovetail 24
that extends downwardly from the airfoil and engages a slot on the
turbine disk where it is secured in position. In the embodiment
shown in FIG. 5, there are four silicon-silicon carbide ceramic
matrix composite core insert sections 34 that are manufactured
using the Silcomp method. Alternatively, the core insert section
may be a monolithic ceramic material. There are also two silicon
carbide-silicon carbide outer shell sections 36 that are
manufactured using a slurry cast MI. The bond between the core
insert section 34 and the outer shell sections 36 are created
during the final processing step of silicon melt infiltration.
[0040] FIG. 6 is a cross-sectional view of a turbine blade dovetail
24 of an embodiment of a CMC turbine blade 20 of the present
invention. The turbine blade 20 includes an airfoil (not shown)
against which the flow of hot exhaust gas is directed. The turbine
blade 20 is mounted to a turbine disk (not shown) by a dovetail 24
that extends downwardly from the airfoil and engages a slot on the
turbine disk where it is secured in position. In the embodiment
shown in FIG. 6, there is one silicon-silicon carbide ceramic
matrix composite core insert section 34 that is manufactured using
a SilComp method. Alternatively, the core insert section may be a
monolithic ceramic material. There are also two outer silicon
carbide-silicon carbide ceramic matrix composite shell sections 36
that have been manufactured using a prepreg MI method. The bond
between the core insert section 34 and the outer shell sections 36
is created during the final processing step of silicon melt
infiltration.
[0041] FIG. 7 is a cross-sectional view of a turbine blade dovetail
24 of an embodiment of a CMC turbine blade 20 of the present
invention. The turbine blade 20 includes an airfoil (not shown)
against which the flow of hot exhaust gas is directed. The turbine
blade 20 is mounted to a turbine disk (not shown) by a dovetail 24
that extends downwardly from the airfoil and engages a slot on the
turbine disk where it is secured in position. In the embodiment
shown in FIG. 7, there are four silicon-silicon carbide ceramic
matrix composite core insert sections 34 that are manufactured
using the Silcomp method. Alternatively, the core insert section
may be a monolithic ceramic material. There are also two silicon
carbide-silicon carbide outer shell sections 36 that are
manufactured using a prepreg MI. The bond between the core insert
section 34 and the outer shell sections 36 are created during the
final processing step of silicon melt infiltration.
[0042] The present invention is also a turbine engine airfoil
comprising an outer shell section(s) and an inner core section(s),
wherein the outer shell section(s) and the inner core section(s)
are bonded together using an MI process. The outer shell section(s)
comprises a SiC/SiC material that is manufactured using a process
selected from the group consisting of a slurry cast MI process and
a prepreg MI process. The inner core section(s) comprises a
material selected from the group consisting of a SiC/SiC composite
material manufactured using a slurry cast MI process or a prepreg
MI process, an Si/SiC composite material, and a monolithic ceramic
material. The Si/SiC composite material may be manufactured using
the Silcomp process. The present invention may be a high pressure
turbine blade, a high pressure turbine vane, a low pressure turbine
blade, or a low pressure turbine vane.
[0043] 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.
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