U.S. patent application number 12/144804 was filed with the patent office on 2009-12-24 for single crystal nickel-based superalloy compositions, components, and manufacturing methods therefor.
This patent application is currently assigned to HONEYWELL INTERNATIONAL INC.. Invention is credited to Yiping Hu.
Application Number | 20090317287 12/144804 |
Document ID | / |
Family ID | 41211225 |
Filed Date | 2009-12-24 |
United States Patent
Application |
20090317287 |
Kind Code |
A1 |
Hu; Yiping |
December 24, 2009 |
SINGLE CRYSTAL NICKEL-BASED SUPERALLOY COMPOSITIONS, COMPONENTS,
AND MANUFACTURING METHODS THEREFOR
Abstract
Single crystal nickel-based superalloy compositions, single
crystal nickel-based superalloy components, and methods of
fabricating such components are provided. In an embodiment, by way
of example only, a single crystal nickel-based superalloy
composition consists essentially of, in weight percent from about
3.8 to about 4.2 percent chromium, from about 10.0 to about 10.5
percent cobalt, from about 1.8 to about 2.2 percent molybdenum,
from about 4.8 to about 5.2 percent tungsten, from about 5.8 to
about 6.2 percent rhenium, from about 5.5 to about 5.8 percent
aluminum, from about 5.8 to about 6.2 percent tantalum, from about
3.8 to about 4.2 percent ruthenium, from about 0.13 to about 0.17
percent hafnium, and a balance of nickel.
Inventors: |
Hu; Yiping; (Greer,
SC) |
Correspondence
Address: |
HONEYWELL INTERNATIONAL INC.;PATENT SERVICES
101 COLUMBIA ROAD, P O BOX 2245
MORRISTOWN
NJ
07962-2245
US
|
Assignee: |
HONEYWELL INTERNATIONAL
INC.
Morristown
NJ
|
Family ID: |
41211225 |
Appl. No.: |
12/144804 |
Filed: |
June 24, 2008 |
Current U.S.
Class: |
420/443 ;
420/444 |
Current CPC
Class: |
C30B 29/52 20130101;
F05D 2300/607 20130101; C30B 11/003 20130101; Y02T 50/672 20130101;
F01D 5/28 20130101; Y02T 50/60 20130101; Y02T 50/67 20130101 |
Class at
Publication: |
420/443 ;
420/444 |
International
Class: |
C22C 19/05 20060101
C22C019/05 |
Claims
1. A single crystal nickel-based superalloy composition consisting
essentially of, in weight percent: from about 3.8 to about 4.2
percent chromium; from about 10.0 to about 10.5 percent cobalt;
from about 1.8 to about 2.2 percent molybdenum; from about 4.8 to
about 5.2 percent tungsten; from about 5.8 to about 6.2 percent
rhenium; from about 5.5 to about 5.8 percent aluminum; from about
5.8 to about 6.2 percent tantalum; from about 3.8 to about 4.2
percent ruthenium; from about 0.13 to about 0.17 percent hafnium;
and a balance of nickel.
2. The single crystal nickel-based superalloy composition of claim
1, further consisting essentially of from about 0.001 to about
0.015 weight percent of a material selected from a group consisting
of yttrium, lanthanum, and a combination thereof.
3. The single crystal nickel-based superalloy composition of claim
2, further consisting essentially of from about 0.02 to about 0.06
weight percent carbon and from about 0.003 to about 0.006 weight
percent boron.
4. The single crystal nickel-based superalloy composition of claim
1, wherein a sum of the molybdenum, tungsten, ruthenium, tantalum,
and rhenium is from about 20.0 to about 25.0 weight percent.
5. A single crystal nickel-based superalloy component fabricated of
a single crystal composition consisting essentially of, in weight
percent: from about 3.8 to about 4.2 percent chromium; from about
10.0 to about 10.5 percent cobalt; from about 1.8 to about 2.2
percent molybdenum; from about 4.8 to about 5.2 percent tungsten;
from about 5.8 to about 6.2 percent rhenium; from about 5.5 to
about 5.8 percent aluminum; from about 5.8 to about 6.2 percent
tantalum; from about 3.8 to about 4.2 percent ruthenium; from about
0.13 to about 0.17 percent hafnium; and a balance of nickel.
6. The single crystal nickel-based superalloy component of claim 5,
wherein the single crystal nickel-based superalloy component is a
turbine blade.
7. The single crystal nickel-based superalloy component of claim 5,
wherein the single crystal nickel-based superalloy component is a
vane.
8. The single crystal nickel-based superalloy component of claim 5,
further consisting essentially of from about 0.001 to about 0.015
weight percent of one selected from a group consisting of yttrium,
lanthanum, and a combination thereof.
9. The single crystal nickel-based superalloy component of claim 8,
further consisting essentially of from about 0.02 to about 0.06
weight percent carbon and from about 0.003 to about 0.006 weight
percent boron.
10. The single crystal nickel-based superalloy component of claim
5, wherein a sum of the molybdenum, tungsten, ruthenium, tantalum,
and rhenium is from about 20.0 to about 25.0 weight percent.
11. A method for fabricating a single crystal nickel-based
superalloy component, the method comprising the steps of: providing
an alloy comprising, in weight percent: from about 3.8 to about 4.2
percent chromium; from about 10.0 to about 10.5 percent cobalt;
from about 1.8 to about 2.2 percent molybdenum; from about 4.8 to
about 5.2 percent tungsten; from about 5.8 to about 6.2 percent
rhenium; from about 5.5 to about 5.8 percent aluminum; from about
5.8 to about 6.2 percent tantalum; from about 3.8 to about 4.2
percent ruthenium; from about 0.13 to about 0.17 percent hafnium;
and a balance of nickel; and fabricating a single crystal component
from the alloy.
12. The method of claim 11, wherein the step of providing an alloy
comprises the step of providing an alloy further comprising from
about 0.001 to about 0.015 weight percent of a material selected
from a group consisting of yttrium, lanthanum, and a combination
thereof.
13. The method of claim 12, wherein the step of providing an alloy
comprises the step of providing an alloy further comprising from
about 0.02 to about 0.06 weight percent carbon and from about 0.003
to about 0.006 weight percent boron.
14. The method of claim 11, wherein the step of fabricating a
single crystal component from the alloy comprises the step of
fabricating a turbine blade and a vane from the alloy.
Description
TECHNICAL FIELD
[0001] The inventive subject matter generally relates to materials
for gas turbine engine applications, and more particularly relates
to single crystal nickel-based superalloy compositions and
components of gas turbine engines made therefrom.
BACKGROUND
[0002] Turbine engines are used as the primary power source for
various kinds of aircrafts. The engines may also serve as auxiliary
power sources that drive air compressors, hydraulic pumps, and
industrial electrical power generators. Most turbine engines
generally follow the same basic power generation procedure.
Compressed air is mixed with fuel and burned, and the expanding hot
combustion gases are directed against stationary turbine vanes in
the engine. The vanes turn the high velocity gas flow partially
sideways to impinge onto turbine blades mounted on a rotatable
turbine disk. The force of the impinging gas causes the turbine
disk to spin at high speed. Jet propulsion engines use the power
created by the rotating turbine disk to draw more air into the
engine, and the high velocity combustion gas is passed out of the
gas turbine aft end to create forward thrust. Other engines use
this power to turn one or more propellers, electrical generators,
or other devices. Because fuel efficiency increases as engine
operating temperatures increase, turbine blades and vanes are often
fabricated from high-temperature materials, such as
precipitation-strengthening superalloys. When in a cast form, these
superalloys, which include nickel- and cobalt-based alloys, possess
many properties that may be desirable during exposure to the engine
operating temperatures. For example, these superalloys may have
better elevated-temperature strength and environmental resistance.
However, although known precipitation-strengthening superalloys
exhibit superior mechanical properties under high temperature and
high pressure conditions, they may be subject to oxidation and
corrosion attack when exposed to high-temperature and
highly-pressurized gases in the turbine engine.
[0003] To protect the superalloys from oxidation and corrosion
attacks, the turbine engine airfoils may include a coating system
thereon. In some cases, the coating system may include a thermal
barrier coating (TBC). Typical TBCs may be formed from yttria
stabilized zirconia (YSZ) and/or yttria stabilized zirconia doped
with other oxides such as Gd.sub.2O.sub.3, TiO.sub.2, and the like.
The coating system may further include an environment-resistant
bond coat underneath the TBC to extend its service life. Bond coats
may be overlay coatings that may include, for example, MCrAlX,
where M is a metallic element like nickel, cobalt, and/or a
combination of both nickel and cobalt, and X is yttrium or other
reactive and metallic elements. A bond coat can also be a diffusion
coating such as a simple aluminide coating, a reactive
element-modified aluminide coating, a platinum-modified aluminide
coating or a reactive element-modified platinum aluminide coating.
Thus, during exposure to high temperature, such as during ordinary
service use thereof, the bond coating oxidizes first to form a
thermally grown oxide (TGO) to protect underlying bond coat from
further oxidation. However, if the TGO layer grows too quickly
and/or becomes too thick, adherence of the TBC to the bond coat may
be weakened, and cracks between the TBC and the TGO as well as
between the TGO and the bond coat may form. In such case, the TBC
may prematurely spall off, thus decreasing the service life of the
superalloy component.
[0004] Accordingly, it is desirable to provide an improved
superalloy composition for gas turbine engine applications, wherein
the superalloy has improved elevated-temperature mechanical
properties and environment-resistant performance. In addition, it
is desirable for the improved superalloy composition to be
relatively inexpensive to produce and to incorporate into engine
systems. Furthermore, other desirable features and characteristics
of the inventive subject matter will become apparent from the
subsequent detailed description of the inventive subject matter and
the appended claims, taken in conjunction with the accompanying
drawings and this background of the inventive subject matter.
BRIEF SUMMARY
[0005] Single crystal nickel-based superalloy compositions, single
crystal nickel-based superalloy components, and methods of
fabricating such components are provided.
[0006] In an embodiment, by way of example only, a single crystal
nickel-based superalloy composition consists essentially of, in
weight percent from about 3.8 to about 4.2 percent chromium, from
about 10.0 to about 10.5 percent cobalt, from about 1.8 to about
2.2 percent molybdenum, from about 4.8 to about 5.2 percent
tungsten, from about 5.8 to about 6.2 percent rhenium, from about
5.5 to about 5.8 percent aluminum, from about 5.8 to about 6.2
percent tantalum, from about 3.8 to about 4.2 percent ruthenium,
from about 0.13 to about 0.17 percent hafnium, and a balance of
nickel.
[0007] In another embodiment, by way of example only, a single
crystal nickel-based superalloy component is fabricated from a
single crystal superalloy consisting essentially of, in weight
percent from about 3.8 to about 4.2 percent chromium, from about
10.0 to about 10.5 percent cobalt, from about 1.8 to about 2.2
percent molybdenum, from about 4.8 to about 5.2 percent tungsten,
from about 5.8 to about 6.2 percent rhenium, from about 5.5 to
about 5.8 percent aluminum, from about 5.8 to about 6.2 percent
tantalum, from about 3.8 to about 4.2 percent ruthenium, from about
0.13 to about 0.17 percent hafnium, and a balance of nickel.
[0008] In still another embodiment, by way of example only, a
method for fabricating a single crystal nickel-based superalloy
component includes the steps of providing an alloy comprising, in
weight percent from about 3.8 to about 4.2 percent chromium, from
about 10.0 to about 10.5 percent cobalt, from about 1.8 to about
2.2 percent molybdenum, from about 4.8 to about 5.2 percent
tungsten, from about 5.8 to about 6.2 percent rhenium, from about
5.5 to about 5.8 percent aluminum, from about 5.8 to about 6.2
percent tantalum, from about 3.8 to about 4.2 percent ruthenium,
from about 0.13 to about 0.17 percent hafnium, and a balance of
nickel and fabricating a single crystal component from the
alloy.
BRIEF DESCRIPTION OF THE DRAWINGS
[0009] The inventive subject matter will hereinafter be described
in conjunction with the following drawing figures, wherein like
numerals denote like elements, and wherein:
[0010] FIG. 1 is a perspective view of a component, according to an
embodiment;
[0011] FIG. 2 is a cross-sectional view of a portion of a
component, according to an embodiment; and
[0012] FIG. 3 is a flow diagram of a process for fabricating a
single crystal nickel-based superalloy component, according to an
embodiment.
DETAILED DESCRIPTION
[0013] The following detailed description of the inventive subject
matter is merely exemplary in nature and is not intended to limit
the inventive subject matter or the application and uses of the
inventive subject matter. Furthermore, there is no intention to be
bound by any theory presented in the preceding background of the
inventive subject matter or the following detailed description of
the inventive subject matter.
[0014] A superalloy composition is provided that may have improved
elevated-temperature properties compared to conventional superalloy
compositions. In an embodiment, the inclusion of certain elements,
such as molybdenum and rhenium, may improve stress rupture strength
and creep resistance, when the superalloy is exposed to engine
operating temperatures, such as temperatures greater than about
2200.degree. F. (1205.degree. C.). In another embodiment, when
element ruthenium is included in the superalloy composition, phase
stability of the superalloy may be greatly improved.
[0015] The superalloy composition may be embodied in a single
crystal nickel-based superalloy component. FIG. 1 is a perspective
view of a component 150, according to an embodiment. Here, the
component 150 is shown as turbine blade. However, in other
embodiments, the component 150 may be a turbine vane or other
component that may be implemented in a gas turbine engine, or other
high-temperature system. In an embodiment, the component 150
includes an airfoil 152 including a pressure side surface 153, an
attachment portion 154, a leading edge 158 including a blade tip
155, and a platform 156. In accordance with an embodiment, the
component 150 may be formed with a non-illustrated outer shroud
attached to the tip 155. The component 150 may have non-illustrated
internal air-cooling passages that remove heat from the turbine
airfoil. After the internal air has absorbed heat from the
superalloy blade, the air is discharged into a combustion gas flow
path through passages 159 in the airfoil wall. Although the turbine
component 150 is illustrated as including certain parts and having
a particular shape and dimension, different shapes, dimensions and
sizes may be alternatively employed depending on particular gas
turbine engine models and particular applications.
[0016] FIG. 2 is a cross-sectional view of a portion of a component
200, according to an embodiment. The component 200 may be, for
example, a turbine airfoil such as the turbine blade shown in FIG.
1 and may include a protective coating system 202 disposed over a
substrate 201. In an embodiment, the protective coating system 202
may include a bond coating 204, a thermal barrier coating 208, and
one or more intermediate layers therebetween, such as a thermally
grown oxide (TGO) 210. In one embodiment, the bond coating 204 is a
diffusion aluminide coating. The diffusion aluminide coating may be
formed by depositing an aluminum layer over the component 200, and
subsequently interdiffusing the aluminum layer with the substrate.
According to one embodiment, the diffusion aluminide coating is a
simple diffusion aluminide, including a single layer made up of
aluminum diffused into the substrate 201. In another embodiment,
the diffusion aluminide coating may have a more complex structure
and may include one or more additional metallic layers that are
diffused into the aluminum and/or the substrate 201. For example,
an additional metallic layer may include a platinum layer, a
hafnium and/or a zirconium layer, or a co-deposited hafnium,
zirconium, and platinum layer. In another embodiment, the bond
coating 204 may be an overlay coating comprising MCrAlX, wherein M
is an element selected from cobalt, nickel, or combinations
thereof, and X is an element selected from hafnium, zirconium,
yttrium, tantalum, rhenium, ruthenium, palladium, platinum,
silicon, or combinations thereof. Some examples of MCrAlX
compositions include NiCoCrAlY and CoNiCrAlY. In another exemplary
embodiment, the bond coating 204 may include a combination of two
types of bond coatings, such as a diffusion aluminide coating
formed on an MCrAlX coating. In any case, the bond coating 204 may
have a thickness in a range of from about 25 microns (.mu.m) to
about 150 .mu.m, according to an embodiment. In other embodiments,
the thickness of the bond coating 204 may be greater or less.
[0017] The thermal barrier coating 208 may be formed over the bond
coating 204 and may comprise, for example, a ceramic. In one
example, the thermal barrier coating 208 may comprise a partially
stabilized zirconia-based thermal barrier coating, such as yttria
stabilized zirconia (YSZ). In an embodiment, the thermal barrier
coating may comprise yttria stabilized zirconia doped with other
oxides, such as Gd.sub.2O.sub.3, TiO.sub.2, and the like. In
another embodiment, the thermal barrier coating 208 may have a
thickness that may vary and may be, for example, in a range from
about 50 .mu.m to about 300 .mu.m. In other embodiments, the
thickness of the thermal barrier coating 208 may be in a range of
from about 100 .mu.m to about 250 .mu.m. In still other
embodiments, the thermal barrier coating 208 may be thicker or
thinner than the aforementioned ranges.
[0018] The thermally-grown oxide layer 210 may be located between
the bond coating 204 and the thermal barrier coating 208. In an
embodiment, the thermally-grown oxide layer 210 may be grown from
aluminum in the above-mentioned materials that comprise the bond
coating 204. For example, during bond coating 204 exposed to high
temperature in air or oxygen-containing atmosphere, oxidation may
occur thereon to result in the formation of the oxide layer 210. In
one embodiment, the thermally-grown oxide layer 210 may be
relatively thin, and may be less than 2 .mu.m thick.
[0019] The substrate 201 over which the coating system 202 is
disposed may be fabricated from a single crystal superalloy
material. A "single crystal superalloy material" may be defined as
a superalloy material formed to have a single crystallographic
orientation throughout its entirety and being substantially free of
high angle boundaries. In some embodiments, an incidental amount of
low angle boundaries, which are commonly defined as the boundaries
between adjacent grains whose crystallographic orientation differs
by less than about 5 degrees, such as tilt or twist boundaries, may
be present within the single crystal superalloy material after
solidification and formation of the single crystal superalloy
material, or after some deformation of the component during creep
or other deformation process. However, preferably, low angle
boundaries are not present in the single crystal superalloy
component.
[0020] The single crystal superalloy material may be comprised of a
material having a microstructure that, after heat treatment, forms
an array of gamma prime precipitates in a matrix. In an embodiment,
the matrix may comprise nickel, which has been strengthened by the
addition of various alloying elements. To achieve a cooperative
optimization of physical, chemical, and mechanical properties of a
completed component and to optimize the retention of such
properties during the operating lifetime of the component, the
alloying elements are selected, which may, either alone or in
combination, exhibit desired properties. Additionally, the
selection of the alloying elements may also depend on an ability to
provide creep strength, phase stability, and environmental
resistance, to the single crystal component.
[0021] In an embodiment, one or more of the alloying elements may
be selected from chromium, cobalt, molybdenum, tungsten, rhenium,
aluminum, tantalum, ruthenium, hafnium, yttrium, lanthanum, carbon,
and boron. According to an embodiment, these elements may be
included to form a superalloy composition. In an embodiment, the
superalloy composition may include, in weight percent, from about
3.8 to about 4.2 percent chromium, from about 10.0 to about 10.5
percent cobalt, from about 1.8 to about 2.2 percent molybdenum,
from about 5.8 to about 6.2 percent rhenium, from about 3.8 to
about 4.2 percent ruthenium, from about 5.5 to about 5.8 percent
aluminum, from about 5.8 to about 6.2 percent tantalum, from about
4.8 to about 5.2 percent tungsten, from about 0.13 to about 0.17
percent hafnium, and a balance of nickel. In accordance with
another embodiment, the superalloy composition may further include,
in weight percent, from about 0.001 to about 0.015 weight percent
of a material selected from the group consisting of yttrium,
lanthanum, and a combination thereof. In still another embodiment,
the superalloy composition may still further include, in weight
percent, from about 0.02 to about 0.06 percent carbon and from
about 0.003 to about 0.006 percent boron. In any case, in an
embodiment, a sum of molybdenum plus tungsten plus rhenium plus
tantalum plus ruthenium present in the superalloy composition may
be from about 20.0 to about 25.0 weight percent.
[0022] It has been found that the inclusion of molybdenum and
rhenium in the composition improves creep strength and stress
rupture property of the alloy by increasing lattice misfit between
the gamma matrix and the gamma-prime precipitates. In this regard,
in an embodiment, an amount of from about 1.8 to about 2.2 percent
molybdenum, by weight, and an amount of from about 5.8 to about 6.2
percent of rhenium, by weight may be included. Additionally,
rhenium may refine the size of the gamma-prime precipitates, which
may contribute to the improved strength of the gamma matrix.
Rhenium may also improve the creep strength of the alloy by
decreasing a coarsening rate of the gamma-prime precipitates,
during extended elevated temperature exposure. It has also been
found that the addition of ruthenium to the superalloy composition
promotes phase stability. Moreover, by balancing elements cobalt
and chromium and strengthening elements such as the molybdenum,
tungsten, tantalum, rhenium, and ruthenium in the range of from
about 20.0 to about 25.0 weight percent, the formation of
topologically close-packed phases (TCP) such as P and R, and sigma
phases may be avoided.
[0023] FIG. 3 is a flow diagram of a process 300 for fabricating a
single crystal superalloy component, according to an embodiment.
First, an alloy of the above-described superalloy composition is
initially provided, step 302. In an embodiment, the alloy includes
the aforementioned superalloy composition. Next, the single crystal
superalloy component then may be fabricated, step 304. In
accordance with an embodiment, the alloy of step 302 may be formed
into a single crystal turbine hardware. Without further processing,
the single crystal hardware would contain gamma-prime precipitates
(referred to as cooling gamma-prime) having a variety of sizes. If
a further solution and aging heat treatment procedure is performed,
the gamma-prime precipitate phase dissolves into the gamma matrix
and is re-precipitated out during an aging treatment conducted at a
lower temperature.
[0024] To dissolve the gamma-prime phase into the gamma matrix, the
single crystal piece is heated to a temperature which is greater
than the solvus temperature of the gamma-prime phase, but less than
the melting temperature of the alloy. The melting temperature,
termed the solidus temperature for an alloy which melts over a
temperature range, should be sufficiently greater than the solvus
temperature so that the single crystal piece may be heated and
maintained within the temperature range between the solvus and the
solidus for a time sufficiently long to dissolve the gamma-prime
precipitation phase into the gamma matrix. The solidus temperature
is about 2,400.degree. F. (about 1315.56.degree. C.) to
2,450.degree. F. (about 1343.33.degree. C.), and accurate control
to within a few degrees in commercial heat treating equipment is
not available. In an embodiment, the solidus may be at least about
15.degree. F. (about 8.3.degree. C.) greater than the solvus
temperature for the gamma-prime precipitates.
[0025] The "heat treatment window" or difference between the
gamma-prime solvus and the alloy solidus temperatures preferably is
at least at 15.degree. F. (about 8.3.degree. C.), and more
preferably is greater than about 50.degree. F. (about 27.8.degree.
C.). In an embodiment of heat treatment of the cast single crystal
pieces, the pieces are solution heat treated at a temperature of
about 2,400.degree. F. (about 1315.6.degree. C.) for a period of
about three hours, to dissolve the gamma-prime precipitation phase
formed during solidification, into the gamma matrix. The solution
heat treatment may be accomplished at any temperature within the
heat treatment window between the gamma-prime solvus and the
solidus temperatures. Greater temperatures allow shorter heat
treatment times. However, the heat treatment temperature is not
raised to a maximum level, to allow a margin of error in the heat
treatment equipment. After the heat treating process is completed,
the solution heat-treated single crystal pieces are rapidly cooled
to supersaturate the matrix with the gamma-prime forming elements.
A fast inert gas fan cool to a temperature of less than about
1,000.degree. F. (about 537.78.degree. C.) has been found
sufficient to achieve the necessary supersaturation.
[0026] Following the solution heat treatment and supersaturation
cooling, the solution heat-treated single crystal pieces are aged
to precipitate the gamma-prime particles from the gamma matrix. The
aging heat treatment can be separated from coating treatment or
combined with the coating treatment. In an embodiment, the aging
heat treatment is conducted at 1150.degree. C. for 4 hours,
followed by 1080.degree. C. for 4 hours and 899.degree. C. for 4 to
20 hours. As noted previously, gas turbine components are typically
coated with a corrosion- and oxidation-resistant coating and
thermal barrier coating prior to use. During the coating procedure,
the component being coated is heated to elevated temperatures. In
an embodiment, the component is heated to a temperature of about
1,950.degree. F. (about 1065.56.degree. C.) for about four hours.
This heat treatment causes some precipitation of the gamma-prime
phase from the gamma matrix, thus partially accomplishing the aging
heat treatment. The aging heat treatment may be completed by a
further elevated temperature exposure, separate from the coating
procedure. A sufficient additional aging heat treatment is
accomplished at a temperature of about 1,650.degree. F. (about
899.degree. C.) for a time of from about four to about twenty
hours, following the heat treatment at 1,950.degree. F. (about
1065.56.degree. C.) for four hours. The aging heat treatment is not
limited to the aforementioned heat treatment sequence, but instead
may be accomplished by any acceptable approach which precipitates
the desired volume fraction of gamma-prime particles, morphology
and size within the gamma matrix, wherein the precipitation may
occur from the supersaturated heat treated gamma matrix.
[0027] In some cases, the microstructure of the as-solidified
single crystals may include irregular gamma-prime particles and
regions of gamma-prime eutectic phases. The solution heat treatment
may dissolve the irregular gamma-prime particles and most or all
the gamma-prime eutectic constituent into the gamma matrix. The
subsequent aging treatment precipitates an array of gamma-prime
precipitates having a generally cuboidal shape and relatively
uniform size. The gamma-prime precipitates may vary from about 0.3
to about 0.5 microns in size.
[0028] To fabricate the single crystal superalloy component from
the single crystal superalloy mentioned above, vacuum-induction
melting and casting processes may be used. In an embodiment, a
thermal gradient solidification method may be employed. Here,
molten metal of the superalloy composition is poured into a heat
resistant ceramic mold having essentially the desired shape of the
final fabricated component. The mold and molten metal contained
therein are placed within a furnace, induction heating coil, or
other heating device to melt the metal, and the mold and molten
metal are gradually cooled in a temperature gradient. In this
process, metal adjacent the cooler end of the mold solidifies
first, and the interface between the solidified and liquid metal
gradually moves through the metal as cooling continues. Such
thermal gradient solidification can be accomplished by placing a
chill block adjacent one end of the mold and then turning off the
heat source, allowing the mold and molten metal to cool and
solidify in a controlled desirable temperature gradient.
Alternatively, the mold and molten metal can be gradually withdrawn
from the heat source.
[0029] It is found that certain crystallographic orientations such
as <001> grow to the exclusion of others during such a
thermal gradient solidification process, so that a single grain
becomes dominant throughout the article. Techniques have been
employed to promote the formation of the single crystal orientation
rapidly, so that substantially all the article has the same single
crystal orientation. Such techniques include seeding whereby an
oriented single crystal starting material is positioned adjacent
the metal first solidified, so that the metal initially develops
that orientation. Another technique is a geometrical selection
process. However, other techniques for forming a single crystal
alternatively may be used. For example, a liquid metal cooling
process or cast may be used to fabricate single crystal turbine
components. In this process, a nickel-based superalloy is melted
and poured into a ceramic mold placed inside a multi-zone heater.
For solidification, the cast components are immersed at a constant
rate into a liquid tin bath.
[0030] Accordingly, single crystal nickel-based superalloy
compositions for hot-section components of gas turbine engines,
such as gas turbine blades and vanes, have been provided. The alloy
compositions provide such components with mechanical and
environment-resistant properties that may be superior to those of
traditional superalloy materials.
[0031] While at least one exemplary embodiment has been presented
in the foregoing detailed description of the inventive subject
matter, it should be appreciated that a vast number of variations
exist. It should also be appreciated that the exemplary embodiment
or exemplary embodiments are only examples, and are not intended to
limit the scope, applicability, or configuration of the inventive
subject matter in any way. Rather, the foregoing detailed
description will provide those skilled in the art with a convenient
road map for implementing an exemplary embodiment of the inventive
subject matter, it being understood that various changes may be
made in the function and arrangement of elements described in an
exemplary embodiment without departing from the scope of the
inventive subject matter as set forth in the appended claims and
their legal equivalents.
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