U.S. patent application number 12/341486 was filed with the patent office on 2009-06-25 for method for fabricating a thick ti64 alloy article to have a higher surface yield and tensile strengths and a lower centerline yield and tensile strengths.
This patent application is currently assigned to GENERAL ELECTRIC COMPANY. Invention is credited to Ming Cheng LI, Peter WAYTE.
Application Number | 20090159161 12/341486 |
Document ID | / |
Family ID | 40787182 |
Filed Date | 2009-06-25 |
United States Patent
Application |
20090159161 |
Kind Code |
A1 |
WAYTE; Peter ; et
al. |
June 25, 2009 |
METHOD FOR FABRICATING A THICK Ti64 ALLOY ARTICLE TO HAVE A HIGHER
SURFACE YIELD AND TENSILE STRENGTHS AND A LOWER CENTERLINE YIELD
AND TENSILE STRENGTHS
Abstract
A Ti-6Al-4V-0.2O (Ti64) forged article is fabricated by forging
a workpiece to make a forged gas turbine engine component having a
thick portion thereof with a section thickness greater than 21/4
inches. The forged article is heat treated by solution heat
treating at a temperature of from about 50.degree. F. to about
85.degree. F. below the beta-transus temperature of the alloy,
thereafter water quenching the gas turbine engine component to room
temperature, and thereafter aging the gas turbine engine component
at a temperature of from about 900.degree. F. to about 1350.degree.
F.
Inventors: |
WAYTE; Peter; (Maineville,
OH) ; LI; Ming Cheng; (Cincinnati, OH) |
Correspondence
Address: |
MCNEES, WALLACE & NURICK LLC
100 PINE STREET, PO BOX 1166
HARRISBURG
PA
17108-1166
US
|
Assignee: |
GENERAL ELECTRIC COMPANY
Schenectady
NY
|
Family ID: |
40787182 |
Appl. No.: |
12/341486 |
Filed: |
December 22, 2008 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
10692985 |
Oct 24, 2003 |
7481898 |
|
|
12341486 |
|
|
|
|
Current U.S.
Class: |
148/557 ;
148/671 |
Current CPC
Class: |
C22F 1/183 20130101;
B21K 3/04 20130101; C21D 9/00 20130101; B21K 3/00 20130101; C22C
14/00 20130101 |
Class at
Publication: |
148/557 ;
148/671 |
International
Class: |
C22F 1/18 20060101
C22F001/18 |
Claims
1. A method for fabricating a forged titanium-alloy article,
comprising the steps of providing a workpiece made of an alpha-beta
titanium alloy having a nominal composition in weight percent of 6
percent aluminum, 4 percent vanadium, 0.2 percent oxygen, balance
titanium and impurities, wherein the titanium alloy has a
beta-transus temperature; thereafter forging the workpiece at a
temperature below the beta-transus temperature to make an
alpha-beta forged gas turbine engine component, wherein the forged
gas turbine engine component has a thick portion thereof with a
section thickness greater than 21/4 inches; thereafter rough
machining the forged gas turbine engine component; thereafter heat
treating the machined forged gas turbine engine component by the
steps consisting essentially of solution heat treating the machined
forged gas turbine engine component at a temperature of from about
50.degree. F. to about 85.degree. F. below the beta-transus
temperature, thereafter water quenching the gas turbine engine
component to room temperature, and thereafter aging the gas turbine
engine component at a temperature of from about 900.degree. F. to
about 1350.degree. F.; and thereafter final machining the forged
gas turbine engine component.
2. The method of claim 1, wherein the step of providing the
workpiece includes the steps of preparing a melt of the titanium
alloy, thereafter casting the melt of the titanium alloy to form an
ingot, thereafter converting the ingot to a billet by hot working,
and thereafter cutting the billet transversely to form a mult that
serves as the workpiece.
3. The method of claim 1, wherein the step of forging the workpiece
includes the step of forging the workpiece to make the forged gas
turbine engine component selected from the group consisting of a
compressor disk, a fan disk, and a gas turbine engine mount.
4. The method of claim 1, wherein the step of forging the workpiece
includes the step of forging the workpiece to make a forged
compressor disk or a forged fan disk.
5. The method of claim 1, wherein the step of aging comprises aging
the gas turbine engine component at a temperature of about
1300.degree. F.
6. The method of claim 1, wherein the step of solution heat
treating includes the step of solution heat treating the machined
forged gas turbine engine component at a temperature of from about
55.degree. F. to about 85.degree. F. below the beta-transus
temperature
7. The method of claim 1, wherein the step of forging includes the
step of forging at a temperature within about 60.degree. F. of the
beta transus temperature.
8. The method of claim 1, wherein the step of solution heat
treating includes the step of solution heat treating the forged gas
turbine engine component for a time of from about 45 minutes to
about 75 minutes.
9. The method of claim 1, wherein the step of water quenching is
initiated within about 20 seconds of completing the step of
solution heat treating.
10. The method of claim 1, wherein the step of aging includes the
step of aging the forged gas turbine engine component for a time of
at least about 4 hours.
11. The method of claim 1, including an additional step, after the
step of forging the workpiece and before the step of heat treating,
of ultrasonically inspecting the forged gas turbine engine
component.
12. The method of claim 1, including an additional step, after the
step of forging the workpiece and before the step of final
machining, of ultrasonically inspecting the forged gas turbine
engine component.
13. A method for fabricating a forged titanium-alloy article
comprising the steps of providing a workpiece made of an alpha-beta
titanium alloy having a nominal composition in weight percent of 6
percent aluminum, 4 percent vanadium, 0.2 percent oxygen, balance
titanium and impurities, wherein the titanium alloy has a
beta-transus temperature; thereafter forging the workpiece at a
temperature within about 60.degree. F. of the beta transus
temperature to make an alpha-beta forged gas turbine engine
component, wherein the forged gas turbine engine component has a
thick portion thereof with a section thickness greater than 21/4
inches; thereafter rough machining the forged gas turbine engine
component; thereafter heat treating the machined forged gas turbine
engine component by the steps consisting essentially of solution
heat treating the machined forged gas turbine engine component at a
temperature of from about 55.degree. F. to about 85.degree. F.
below the beta-transus temperature, thereafter water quenching the
gas turbine engine component to room temperature, and thereafter
aging the gas turbine engine component at a temperature of about
1300.degree. F.; and thereafter final machining the gas turbine
engine component.
14. The method of claim 13, wherein the step of providing the
workpiece includes the steps of preparing a melt of the titanium
alloy, thereafter casting the melt of the titanium alloy to form an
ingot, thereafter converting the ingot to a billet by hot working,
and thereafter cutting the billet transversely to form a mult that
serves as the workpiece.
15. The method of claim 13, wherein the step of forging the
workpiece includes the step of forging the workpiece to make the
forged gas turbine engine component selected from the group
consisting of a compressor disk, a fan disk, and a gas turbine
engine mount.
16. The method of claim 13, wherein the step of forging the
workpiece includes the step of forging the workpiece to make a
forged compressor disk or a forged fan disk.
17. The method of claim 13, wherein the step of solution heat
treating includes the step of solution heat treating the forged gas
turbine engine component for a time of from about 45 minutes to
about 75 minutes.
18. The method of claim 13, wherein the step of water quenching is
initiated within about 20 seconds of completing the step of
solution heat treating.
19. The method of claim 13, wherein the step of aging includes the
step of aging the forged gas turbine engine component for a time of
at least about 4 hours.
20. The method of claim 13, wherein the step of final machining
includes the step of removing alpha-case at a surface of the gas
turbine engine component.
Description
RELATED APPLICATIONS
[0001] This application is a continuation-in-part of allowed U.S.
application Ser. No. 10/692,985 filed Oct. 23, 2003, which is
hereby incorporated by reference.
FIELD OF THE INVENTION
[0002] This invention relates to the fabrication of thick articles
of Ti64 alloy and, more particularly, to the fabrication of such
articles with a controllable difference in the near-surface and
centerline mechanical properties.
BACKGROUND OF THE INVENTION
[0003] Ti64 alloy, having a nominal composition in weight percent
of 6 percent aluminum, 4 percent vanadium, 0.2 percent oxygen,
balance titanium and impurities, is one of the most widely used
titanium-base alloys. The Ti64 alloy is an alpha-beta titanium
alloy that may be heat treated to have a range of properties that
are useful in aerospace applications. Ti64 alloy is used in both
thin-section and thick-section applications, and heat treated
according to the section thickness. In an example of interest, Ti64
alloy is used to make thick-section forged parts of aircraft gas
turbine engines, such as compressor disks, fan disks, and engine
mounts, which have at least some locations with a section thickness
of greater than 21/4 inches. The present approach is concerned with
such thick-section articles.
[0004] In the current best practice to achieve the optimal
combination of strength and other properties, after forging the
thick-section Ti64 articles are typically heat treated at a
temperature of 1750.degree. F., followed by an anneal heat
treatment at 1300.degree. F. The result is a 0.2 percent yield
strength throughout the article of from about 120 ksi ("ksi" is an
abbreviation for "thousands of pounds per square inch") to about
140 ksi. This strength has been satisfactory for many thick-section
applications.
[0005] To achieve higher yield strengths in the article, a more
heavily alloyed, heavier forgeable alloy such as Ti 17, having a
nominal composition in weight percent of 5 percent aluminum, 4
percent molybdenum, 4 percent chromium, 2 percent tin, and 2
percent zirconium, is used. The Ti 17 alloy uses a higher
percentage of expensive alloying elements than does Ti64 alloy,
with the result that a large, thick-section part made of Ti 17
alloy is significantly more expensive than the same part made of
Ti64 alloy.
[0006] There is a need for an improved approach to achieving
excellent mechanical properties in forgeable titanium alloys. The
present invention fulfills this need, and further provides related
advantages.
SUMMARY OF THE INVENTION
[0007] The present invention provides a fabrication approach for
thick-section parts made of Ti64 alloy. This approach achieves
significantly improved properties where needed for the surface and
near-surface regions of the thick-section parts made of this
well-proven alloy. The ability to use an established alloy is an
important advantage, as new procedures for melting, casting, and
forging a new alloy are not required. Nor is it necessary to employ
a more heavily alloyed composition such as Ti17.
[0008] A method for fabricating a forged titanium-alloy article
comprises the steps of providing a workpiece made of a titanium
alloy having a nominal composition in weight percent of 6 percent
aluminum, 4 percent vanadium, 0.2 percent oxygen, balance titanium
and impurities. The titanium alloy has a beta-transus temperature.
The workpiece is thereafter forged at a temperature below the beta
transus temperature to make an alpha-beta forged gas turbine engine
component, such as a compressor disk, a fan disk, or a gas turbine
engine mount. The forged article, which is preferably a gas turbine
engine component, has a thick portion thereof with a section
thickness greater than 21/4 inches.
[0009] The forged gas turbine engine component is thereafter heat
treated by solution heat treating the forged gas turbine engine
component at a temperature of from about 50.degree. F. to about
85.degree. F. below the beta-transus temperature, preferably for a
time of from about 45 minutes to about 75 minutes. The gas turbine
engine component is thereafter quenched to room temperature and
thereafter aged for a minimum of 4 hours at a temperature between
900.degree. F. and 1300.degree. F. Desirably, the water quenching
is initiated within about 20 seconds of completing the step of
solution heat treating by removal of the component from the
solution-treating furnace.
[0010] The forged gas turbine engine component is thereafter final
machined. The final machining is typically performed both to remove
the high-oxygen, less ductile alpha-case at the surface and to
produce the final features of the gas turbine engine component.
[0011] In the usual practice, the forged gas turbine engine
component is ultrasonically inspected in a rough-machined shape
generated by rough machining the forging either prior to the
solution heat treat or following all heat treatment. The ultrasonic
inspection is performed either after the step of forging the
workpiece and before the step of heat treating, or after the step
of heat treating and before the step of final machining. Where the
forged gas turbine engine component is a compressor or fan disk,
and where the ultrasonic inspection is performed after the step of
forging and before the step of heat treating, after the ultrasonic
inspection rough slots may be machined into the periphery of the
disk so that the subsequent heat treatment imparts the improved
properties to the bottoms of the slots.
[0012] The thick section of the gas turbine engine component given
this heat treatment procedure desirably has a 0.2 percent yield
strength of from about 120 ksi to about 140 ksi at its centerline,
and a higher 0.2 percent yield strength of from about 160 ksi to
about 175 ksi at a location nearer a surface thereof. The higher
yield strength region of about 160-175 ksi typically extends
downwardly from the surface of the gas turbine engine component to
a depth of from about 3/4 to about 1 inch below the surface. There
is additionally an increase in the tensile strength associated with
the increased yield strength. At greater depths, the gas turbine
engine component has the lower yield strength range of about
120-140 ksi.
[0013] In the work leading to the present invention, it was
recognized that the near-surface regions of the thick gas turbine
engine components are subjected to the highest stresses in service
at locations about 1/2 inch below the final machined finished part
surface. The present heat treatment procedure produces the highest
yield strength and tensile strength material in the near surface
regions of the thick article, where the tensile strength is most
needed. The near surface regions thus perform mechanically as
though they are made of a stronger material than the conventionally
heat treated Ti64 material that is found toward the center regions
of the thick article. The result is that the Ti64 material may be
used in applications for which it would otherwise not have
sufficient mechanical properties.
[0014] 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. The scope of the invention is not, however, limited
to this preferred embodiment.
BRIEF DESCRIPTION OF THE DRAWINGS
[0015] FIG. 1 is a block flow diagram of a preferred embodiment of
an approach for fabricating a forged titanium-alloy article;
[0016] FIG. 2 is a perspective view of a disk such as a compressor
disk or a fan disk;
[0017] FIG. 3 is a perspective view of a gas turbine engine mount;
and
[0018] FIG. 4 is a schematic sectional view through the disk of
FIG. 2, taken on line 4-4.
DETAILED DESCRIPTION OF THE INVENTION
[0019] FIG. 1 depicts in block diagram form a method for practicing
a preferred approach for fabricating a forged titanium-alloy
article. The method comprises the steps of providing a workpiece
made of the titanium alloy, known as Ti64, having a nominal
composition in weight percent of 6 percent aluminum, 4 percent
vanadium, 0.2 percent oxygen, balance titanium and impurities, step
20. The Ti64 titanium alloy has a nominal beta-transus temperature
of about 1820.degree. F., although the beta-transus temperature
varies with compositional variations from the nominal composition.
In the preferred practice, the titanium alloy is melted and cast as
an ingot, and converted by hot working to billet form. The billet
is sliced transversely to form a workpiece termed a "mult".
[0020] In the preferred embodiment, the workpiece is forged to make
a forged gas turbine engine component, step 22. (As used herein,
"forged gas turbine engine component" includes both the final
forged gas turbine engine component and also the precursors of the
final article resulting from the forging step 22.) The forged gas
turbine engine component has a thick portion thereof with a section
thickness greater than 21/4 inches, termed a "thick-section"
article. The entire forged gas turbine engine component need not
have a section thickness greater than 21/4 inches, as long as at
least some portion of the forged gas turbine engine component has
the section thickness of greater than 21/4 inches. FIGS. 2-3
illustrate the final form (after all of the processing is complete)
of two forged gas turbine engine components of particular interest,
a compressor or fan disk 50 (FIG. 2) and a gas turbine engine mount
60 (FIG. 3).
[0021] The step 20 of providing the workpiece is performed by
conventional techniques known in the art. The step of forging may
also occur by known techniques, but is preferably conducted below a
beta transus temperature of the workpiece, and more preferably
within about 60.degree. F. of the beta transus temperature to
produce an alpha-beta workpiece.
[0022] After the forging step 22, the forged gas turbine engine
component is optionally ultrasonically inspected, step 24, by known
techniques. In the usual practice where step 24 is performed, the
forged gas turbine engine component is first annealed at
1300.degree. F. for 1 hour and cooled to room temperature. It is
then rough machined into a rough-machined shape with at least some
flat sides to facilitate the ultrasonic inspection of step 24. The
rough-machined shape is larger than the final machined shape of the
article, so that at least some material may be machined away in the
subsequent final-machining step. In one embodiment, the rough
machining is within about one inch of the final dimensions of the
article to be formed. In the case where the forged gas turbine
engine component is a compressor or fan disk, after the ultrasonic
inspection is performed rough slots 52 may be machined into the
periphery of the disk so that the subsequent heat treatment imparts
the improved properties to the surface and near-surface regions
near the bottoms of the slots.
[0023] The forged gas turbine engine component is heat treated,
step 26. The heat treatment 26 includes three substeps, performed
sequentially one after the other as illustrated. The first substep
28 is solution heat treating the forged gas turbine engine
component at a solution-heat-treatment temperature of from about
50.degree. F. to about 85.degree. F. below the beta-transus
temperature, preferably in the range of about 55.degree. F. to
about 85.degree. F. below the beta-transus temperature. The nominal
beta-transus temperature for Ti64 alloy is about 1820.degree. F.,
and the solution heat treating step 28 is performed at a
temperature of from about 1770.degree. F. to about 1735.degree. F.
for the nominal-composition Ti64 alloy. This
solution-heat-treatment temperature range may be adjusted somewhat
for variations in the exact composition of the Ti64 alloy being
employed, as long as the solution-heat-treatment temperature is
from about 50.degree. F. to about 85.degree. F. below the
beta-transus temperature. The preferred time for solution heat
treating of the forged gas turbine engine component is from about
45 minutes to about 75 minutes, most preferably about 60 minutes,
at the solution heat treating temperature of from about 50.degree.
F. to about 85.degree. F. below the beta-transus temperature. The
solution heat treating 28 is preferably accomplished in air and in
a furnace held at the solution heat treatment temperature.
[0024] The second substep of the heat treatment 26 is water
quenching the gas turbine engine component to room temperature,
step 30. The gas turbine engine component is transferred from the
solution heat treating furnace to a water quench bath as quickly as
possible at the conclusion of step 28. Desirably, the water
quenching 30 is initiated within about 20 seconds of removing the
gas turbine engine component from the solution-heat-treating
furnace, which removal completes the solution heat treating step
28.
[0025] The third substep of the heat treatment 26 is aging the gas
turbine engine component at a temperature in the range of from
about 900.degree. F. to about 1350.degree. F., step 32, after the
step 30 is complete. Preferably, the aging is at a temperature of
about 1300.degree. F. The aging step 32 is preferably continued for
a time of at least about 4 hours after all of the gas turbine
engine component reaches the aging temperature. The aging heat
treating 32 is preferably accomplished in air and in a furnace held
at the aging heat treatment temperature.
[0026] After the heat treating step 26, the forged-and-heat-treated
gas turbine engine component is optionally ultrasonically
inspected, step 34, by known techniques. If the gas turbine engine
component has not previously been rough machined in the manner
discussed in relation to step 24, that rough machining is performed
as part of step 34, before the ultrasonic inspection. Although
steps 24 and 34 are each optional, it is desirable that at least
one of them be performed.
[0027] The gas turbine engine component is thereafter final
machined to the finished shape and dimensions, step 36. The final
machining removes the high-oxygen, less ductile alpha-case on the
surface of the forging, typically a thickness of about 0.020 inches
of material, and also produces the final features of the gas
turbine engine component, such as the final form of the dovetail
slots 52 on the rim of the compressor or fan disk 50 of FIG. 2.
[0028] FIG. 4 is a schematic sectional view of the disk 50,
illustrating the structure resulting from the present approach.
There is a section centerline 54 and two surfaces 56 of the disk
50. The section has a local section thickness t.sub.s that may be
constant or, as illustrated, variable. At least some portion of the
section thickness t.sub.s is greater than 21/4 inches, so that the
disk 50 may be considered a "thick" section. There is a hardened
depth d.sub.H of a hardened zone 58 extending below each of the
surfaces 56. The hardened depth d.sub.H typically extends from the
surface 56 to a depth of from about 3/4 inch to about 1 inch below
the surface 56, the "near-surface" region. The 0.2 percent yield
strength of the material in the hardened zone 58, such as at a
depth of about 1/2 inch below the surface, is from about 160 ksi
("ksi" is a standard abbreviation for "thousands of pounds per
square inch", so that 160 ksi is 160,000 pounds per square inch) to
about 175 ksi in the hardened zone 58. The remaining central zone
59, which can have a variable thickness as illustrated, has a lower
yield strength. The 0.2 percent yield strength is from about 120
ksi to about 140 ksi measured at the centerline 54.
[0029] This variation in yield strength is produced by the heat
treatment of step 26 of FIG. 1. The different yield strengths
within the two zones 58 and 59 is a desirable feature, so that the
greatest yield strength is provided where it is needed during the
service of the gas turbine engine component, near its surface.
[0030] It has been known in the art to heat treat thin pieces of
Ti64 material, less than about 2 inches thick, by solution heat
treating at a temperature of from about 50.degree. F. to about
75.degree. F. below the beta-transus temperature, thereafter water
quenching to a temperature of less than about 850.degree. F., and
thereafter aging at a temperature of from about 900.degree. F. to
about 1000.degree. F. However, the benefits could not be extended
to thicknesses greater than about 2 inches. In the present
approach, it is recognized that a harder zone near the surface of
the article and a softer zone in the center of the article is
beneficial to the resulting properties. This approach permits the
Ti64 alloy to be used to higher performance levels, and avoids the
need to utilize more-expensive alloys to make thick-section
articles.
[0031] Although a particular embodiment of the invention has been
described in detail for purposes of illustration, various
modifications and enhancements may be made without departing from
the spirit and scope of the invention. Accordingly, the invention
is not to be limited except as by the appended claims.
* * * * *