U.S. patent application number 09/957844 was filed with the patent office on 2003-03-27 for pre-service oxidation of gas turbine disks and seals.
Invention is credited to Heaney, Joseph Aloysius III, Nagaraj, Bangalore Aswatha, Schaeffer, Jon Conrad, Weimer, Michael James.
Application Number | 20030056855 09/957844 |
Document ID | / |
Family ID | 25500223 |
Filed Date | 2003-03-27 |
United States Patent
Application |
20030056855 |
Kind Code |
A1 |
Weimer, Michael James ; et
al. |
March 27, 2003 |
PRE-SERVICE OXIDATION OF GAS TURBINE DISKS AND SEALS
Abstract
A gas turbine component, such as a turbine disk or a rotating
seal, is fabricated by furnishing a substrate shaped as a gas
turbine component made of a nickel-base superalloy, and oxidizing
the substrate to produce an oxidized substrate having thereon a
layer comprising an oxide and having a thickness of at least about
500 Angstroms. The step of oxidizing is performed prior to entry of
the component into service and in an atmosphere that does not
contain combustion gas. The oxidized gas turbine component is
thereafter placed into service.
Inventors: |
Weimer, Michael James;
(Loveland, OH) ; Nagaraj, Bangalore Aswatha; (West
Chester, OH) ; Schaeffer, Jon Conrad; (Simpsonville,
SC) ; Heaney, Joseph Aloysius III; (Middletown,
OH) |
Correspondence
Address: |
GREGORY GARMONG
P.O. BOX 12460
ZEPHYR COVE
NV
89448
US
|
Family ID: |
25500223 |
Appl. No.: |
09/957844 |
Filed: |
September 21, 2001 |
Current U.S.
Class: |
148/276 ;
148/277; 29/889; 29/889.2 |
Current CPC
Class: |
Y10T 29/49316 20150115;
Y10T 29/49336 20150115; F01D 5/288 20130101; C23C 28/044 20130101;
C23C 8/80 20130101; Y10T 29/4932 20150115; F05D 2300/21 20130101;
C23C 8/14 20130101; C23C 8/02 20130101 |
Class at
Publication: |
148/276 ;
29/889.2; 29/889; 148/277 |
International
Class: |
C23C 008/10 |
Claims
What is claimed is:
1. A method for fabricating a gas turbine component, comprising the
steps of: furnishing a substrate shaped as a gas turbine component
made of a nickel-base superalloy; thereafter oxidizing the
substrate to produce an oxidized substrate having thereon an oxide
layer with a thickness of at least about 500 Angstroms, the step of
oxidizing being performed in an atmosphere that does not contain
combustion gas: and thereafter placing the oxidized substrate into
service.
2. The method of claim 1, wherein the step of furnishing the
substrate includes a step of mechanically processing the substrate
using a process selected from the group consisting of machining,
peening, and grit blasting.
3. The method of claim 1, wherein the step of furnishing the
substrate includes a step of furnishing a component selected from
the group consisting of a gas turbine disk and a gas turbine
seal.
4. The method of claim 1, wherein the step of furnishing the
substrate includes a step of depositing a protective coating on the
substrate, and wherein the step of oxidizing produces an oxidized
protective coating.
5. The method of claim 4, wherein the step of depositing the
protective coating includes the step of depositing the protective
coating comprising an element selected from the group consisting of
aluminum and chromium, and mixtures thereof.
6. The method of claim 1, wherein the step of oxidizing the
substrate includes the step of heating the substrate in air.
7. The method of claim 1, wherein the step of oxidizing the
substrate includes the step of heating the substrate in an
atmosphere comprising from about 0.2 to about 1000 parts per
million of oxygen.
8. The method of claim 1, wherein the step of oxidizing the
substrate includes the step of heating the substrate to a
temperature of at least about 1200.degree. F., for a time of at
least about 2 hours.
9. The method of claim 1, including an additional step, after the
step of oxidizing the substrate and prior to the step of placing
the oxidized substrate into service, of depositing a top coating on
the oxidized substrate.
10. The method of claim 1, wherein the step of oxidizing the
substrate includes the step of oxidizing the substrate to produce
the oxide layer with the thickness of from about 500 Angstroms to
about 6000 Angstroms.
11. A method for fabricating a gas turbine component, comprising
the steps of: furnishing a substrate shaped as a gas turbine
component made of a nickel-base superalloy; thereafter oxidizing
the substrate to produce an oxidized substrate, the step of
oxidizing being conducted at a temperature of at least about
1200.degree. F., for a time of at least about 2 hours, and in an
oxygen-containing atmosphere that does not contain combustion gas;
and thereafter placing the oxidized substrate into service.
12. The method of claim 11, wherein the step of furnishing the
substrate includes a step of mechanically processing the substrate
using a process selected from the group consisting of machining,
peening, and grit blasting.
13. The method of claim 11, wherein the step of furnishing the
substrate includes a step of depositing a protective coating on the
substrate, and wherein the step of oxidizing produces an oxidized
coating.
14. The method of claim 13, wherein the step of depositing the
protective coating includes the step of depositing the protective
coating comprising an element selected from the group consisting of
aluminum, chromium, and mixtures thereof.
15. The method of claim 1, wherein the step of oxidizing the
substrate includes the step of heating the substrate in air.
16. The method of claim 11, wherein the step of oxidizing the
substrate includes the step of heating the substrate in an
atmosphere comprising from about 0.2 to about 1000 parts per
million of oxygen.
17. The method of claim 11, including an additional step, after the
step of oxidizing the substrate and prior to the step of placing
the oxidized substrate into service, of depositing a top coating on
the oxidized substrate.
18. The method of claim 11, wherein the step of oxidizing the
substrate includes the step of oxidizing the substrate to produce
an oxide layer having a thickness of from about 500 Angstroms to
about 6000 Angstroms.
19. The method of claim 11, wherein the step of furnishing the
substrate includes a step of furnishing a component selected from
the group consisting of a gas turbine disk and a gas turbine seal.
Description
[0001] This invention relates to gas turbine engines and, more
particularly, to the fabrication of the turbine disks and seals,
and their protection against oxidation and corrosion.
BACKGROUND OF THE INVENTION
[0002] In an aircraft gas turbine (jet) engine, air is drawn into
the front of the engine, compressed by a shaft-mounted compressor,
and mixed with fuel. The mixture is burned, and the hot combustion
gases are passed through a turbine mounted on the same shaft. The
flow of combustion gas turns the turbine, which turns the shaft and
provides power to the compressor and to the fan. In a more complex
version of the gas-turbine engine, the compressor and a
high-pressure turbine are mounted on one shaft having a first set
of turbines, and the fan and a low-pressure turbine are mounted on
a separate shaft having a second set of turbines. The hot exhaust
gases and the air propelled by the fan flow from the back of the
engine, driving it and the aircraft forward. The hotter the
combustion and exhaust gases, the more efficient is the operation
of the jet engine. There is thus an incentive to raise the
combustion-gas temperature.
[0003] The turbine (sometimes termed a "turbine rotor") includes
one or more turbine disks, a number of turbine blades mounted to
the turbine disks and extending radially outwardly therefrom into
the combustion-gas flow path, and rotating seals that prevent the
hot combustion gases from contacting the turbine shaft and related
components. The maximum operating temperature of the combustion gas
is limited by the materials used in the turbine. Great efforts have
been made to increase the temperature capabilities of the turbine
blades, resulting in increasing combustion gas operating
temperatures and increased engine efficiency.
[0004] As the maximum operating temperature of the combustion gas
increases, the turbine disk and seals are subjected to higher
temperatures in the combustion-gas environment. As a result,
oxidation and corrosion of the turbine disk and seals have become
of greater concern. Alkaline sulfate deposits resulting from the
ingested dirt and the sulfur in the combustion gas are a major
source of the corrosion, but other elements in the aggressive
combustion-and bleed gas environment may also accelerate the
corrosion. The oxidation and corrosion damage may lead to premature
removal and replacement of the turbine disk and seals unless the
damage is reduced or repaired.
[0005] The turbine disks and seals for use at the highest operating
temperatures are made of nickel-base superalloys selected for good
toughness and fatigue resistance. These superalloys are selected
for their mechanical properties. They have some resistance to
oxidation and corrosion damage, but that resistance is not
sufficient to protect them at the operating temperatures that are
now being reached.
[0006] The current state of the art is to operate the turbine disks
and seals without any coatings to protect them against oxidation
and corrosion. At the same time, a number of oxidation-resistant
and corrosion-resistance coatings have been considered for use on
the turbine blades. These available turbine-blade coatings are
generally too thick and heavy for use on the turbine disks and
seals and also may adversely affect the fatigue life of the turbine
disks and seals. There remains a need for an approach for
protecting turbine disks and seals against oxidation and corrosion
as the operating-temperature requirements of the turbine disks and
seals increase. This need extends to other components of the gas
turbine engine as well. The present invention fulfills this need,
and further provides related advantages.
BRIEF SUMMARY OF THE INVENTION
[0007] The present approach provides an approach for fabricating a
nickel-base superalloy component of a gas turbine engine, such as a
turbine disk or a seal, and components made thereby. The gas
turbine component has improved oxidation and corrosion resistance
as compared with conventional gas turbine components. There is very
little increased weight and added dimension to the turbine
component as a result of utilizing the present approach. The
present fabrication approach is economically applied and is
environmentally friendly. It is not limited by line-of-sight
application procedures, so that otherwise-inaccessible portions of
the component may be treated. The protection extends over the
entire processed surface area of the component, so that protection
is provided even in areas where there may be cracks or
discontinuities in other applied coatings.
[0008] A method for fabricating a gas turbine component comprises
the steps of furnishing a substrate shaped as a gas turbine
component, such as a gas turbine disk or a seal, and made of a
nickel-base superalloy, and oxidizing the substrate to produce an
oxidized substrate having thereon a layer comprising an oxide and
having a thickness of at least about 500 Angstroms. The step of
oxidizing is performed in an atmosphere that does not contain
combustion gas. The oxidized substrate is thereafter placed into
service.
[0009] This approach may be used in conjunction with a number of
additional processing steps. The step of furnishing the substrate
may include a step of preprocessing the substrate by machining,
peening, and grit blasting. A protective coating may be deposited
on the substrate, so that the step of oxidizing produces an
oxidized coating. The protective coating may include an element
such as aluminum, chromium, silicon, phosphorus, or mixtures
thereof.
[0010] The oxidizing step may be performed in an air atmosphere, so
that there is some formation of nitrides as well. The oxidizing
step may be performed in an oxygen-only atmosphere, such as from
about 0.2 to about 1000 parts per million of oxygen. In a typical
case, the step of oxidizing the substrate includes heating the
substrate to a temperature of from about 1200.degree. F. to about
1550.degree. F., for a time of at least about 2 hours.
[0011] A top coating may optionally be deposited on the oxidized
substrate after the oxidation but before the oxidized substrate is
placed into service.
[0012] The present invention involves in-situ formation of an oxide
layer, not deposition of a coating or a layer from a separate
source. The approach does not involve line-of-sight deposition, so
that the entire component is protected without regard to position
relative to a source. The oxidation is performed after all forging
and other mechanical surface processing of the component to its
final shape and surface condition are completed, although
subsequent coating that does not disrupt the oxide is permitted.
Further mechanical operations after oxidation would disrupt the
oxide and render it ineffective.
[0013] The oxide layer typically has a thickness of from about 1000
Angstroms to about 6000 Angstroms, so that it adds very little
weight or dimension to the component. This thin oxide layer
improves the oxidation and corrosion resistance of the component by
at least 50 percent as compared with an unprotected component,
without adversely affecting the mechanical properties such as
strength, toughness, and fatigue resistance. The oxide layer
includes oxides of the components of the superalloy, such as
chromium, titanium, nickel, cobalt, aluminum, and tantalum, and may
also include titanium and other nitrides if the oxidation is
performed in air.
[0014] An important feature of the present processing is that the
oxidation treatment is performed prior to the component entering
service, and without combustion gas or other gas containing
corrosive agents present. Prior turbine components are oxidized
when they enter service and are heated to their operating
temperatures, but that oxidation is performed in an environment
that includes the combustion products which inhibit the formation
of a protective oxide and include compounds such as the sulfides
and carbides that contribute to corrosion damage. In that prior
approach, the corrosive agents are incorporated into the surface of
the turbine component before the oxide has a chance to form in the
manner of the present approach.
[0015] 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
[0016] FIG. 1 is a schematic elevational view of a
turbine-disk-and-seal structure;
[0017] FIG. 2 is a block flow diagram of an approach for practicing
the invention;
[0018] FIG. 3 is an enlarged schematic sectional view through the
turbine disk of FIG. 1 along line 3-3, but prior to its entering
service, illustrating a first embodiment of the invention;
[0019] FIG. 4 is an enlarged schematic sectional view through the
turbine disk of FIG. 1 along line 3-3, but prior to its entering
service, illustrating a second embodiment of the invention;
[0020] FIG. 5 is an enlarged schematic sectional view through the
turbine disk of FIG. 1 along line 3-3, but prior to its entering
service, illustrating a third embodiment of the invention; and
[0021] FIG. 6 is a chart comparing the cycles to failure for
specimens given various treatments.
DETAILED DESCRIPTION OF THE INVENTION
[0022] FIG. 1 schematically depicts a turbine-disk-and-seal
structure 20 including a stage 1 turbine disk 22 and a stage 2
turbine disk 24 mounted to a shaft 26. Seals 28 are mounted to the
shaft 26 and rotate with the shaft 26 to protect the shaft 26 from
the flow 30 of hot combustion gases. Sets of turbine blades 32 and
34 extend from the turbine disk 22 and the turbine disk 24,
respectively. The turbine blades 32 and 34 are protected by their
own protective systems, and, are not the subject of the present
invention. The present invention is concerned with damage to and
protection of the turbine disks 22 and 24 and the seals 28. The
present approach may be applied as appropriate to other components
of the gas turbine engine, such as the stationary shroud seals (not
shown).
[0023] The seals 28 include a CDP seal 36, a forward seal 38, an
interstage seal 40, and an aft seal 42. The forward seal 38 and a
forward-facing side 44 of the stage 1 turbine disk 22 are
particularly subject to corrosion and oxidation damage due to the
combination of heat and corrosive/oxidative effects of the
contaminants in the bleed gas cooling. The preferred embodiment of
the present invention is concerned with protecting these areas,
although it is applicable to the protection of other areas and
components as well.
[0024] FIG. 2 depicts an approach for practicing the present
invention. FIGS. 3-5 illustrate structures produced by this
approach. An article and thence a substrate 70 with a surface 72 is
provided, numeral 50. The article is preferably a component of a
gas turbine engine, preferably the turbine disk 22 or the forward
seal 38. The article is typically a polycrystal made of a
nickel-base superalloy. As used herein, "nickel-base" means that
the composition has more nickel present than any other element. The
nickel-base superalloys are typically of a composition that is
strengthened by the precipitation of gamma-prime phase or a related
phase. The nickel-base superalloy alloy typically has a
composition, in weight percent, of from about 4 to about 25 percent
cobalt, from about 10 to about 20 percent chromium, from about 0 to
about 7 percent aluminum, from 0 to about 12 percent molybdenum,
from about 1 to about 5 percent tungsten, from about 0 to about 3
percent tantalum, from 0 to about 6 percent titanium, from 0 to
about 6 percent niobium, from 0 to about 0.3 percent carbon, from 0
to about 0.02 percent boron, from 0 to about 1.5 percent hafnium,
balance nickel and incidental impurities. Specific examples of
nickel-base superalloys with which the present invention is
operable are Rene 88DT, having a nominal composition in weight
percent of 13 percent cobalt, 16 percent chromium, 2.1 percent
aluminum, 3.7 percent titanium, 4 percent tungsten, 0.7 percent
niobium, 4 percent molybdenum, 0.03 percent zirconium, balance
nickel and minor elements; and ME3, having a nominal composition in
weight percent of 20.6 percent cobalt, 13 percent chromium, 3.4
percent aluminum, 3.7 percent titanium, 2.1 percent tungsten, 2.4
percent tantalum, 0.9 percent niobium, 3.8 percent molybdenum,
balance nickel and minor elements. The present approach is operable
with other alloys as well
[0025] The substrate 70 is typically worked to its desired shape
and size, as by forging or rolling, as part of step 50. It may
optionally thereafter be mechanically processed, numeral 52, using
a metalworking technique such as machining, peening, or grit
blasting. In machining, material is removed from the surface 72 of
the substrate 70 in relatively large cuttings or other pieces or
amounts. In grit blasting, a relatively small amount of material is
removed from the surface 72 by contact with an abrasive grit
propelled toward the surface. In peening, material is not removed
from the surface 72 but instead the surface 72 is worked by the
impingement of shot. Machining, grit blasting, and peening are all
known metalworking techniques for use in other contexts.
[0026] Optionally, a protective coating 74 may be deposited on the
surface 72, numeral 54. FIG. 3 illustrates the case where such a
protective coating 74 is deposited, while FIGS. 4 and 5 illustrate
cases where no such protective coating 74 is applied. Such a
protective coating 74 may be an aluminide or chromide coating
deposited by a process such as chemical vapor deposition, slurry,
or pack cementation.
[0027] The substrate 70, with or without practicing the optional
steps 52 and/or 54, is oxidized, numeral 56. If step 54 is not
employed, the surface 72 of the substrate 70 is oxidized (FIG. 4
and FIG. 5). If step 54 is employed, a surface 76 of the protective
coating 74 is oxidized (FIG. 3). The oxidation 56 is performed
after steps 52 and 54, if any, and before placing the component
into service. The step 56 is therefore performed in the absence of
combustion gas, bleed gas, and the corrosive species that are
present in the vicinity of the gas turbine disk during service. The
oxidation 56 produces a layer 78 comprising oxides of the elements
present at the exposed surface 72 or 76. This layer 78 is termed
herein an "oxide layer", although it may also contain non-oxide
species such as nitrides and specifically titanium nitride, if the
oxidation step 56 is performed in an atmosphere that contains
nitrogen.
[0028] The oxide layer 78 is formed by heating the substrate 70 and
the protective coating 74, where present, in an oxygen-containing
atmosphere. In one embodiment, the oxygen-containing atmosphere has
from about 0.2 to about 1000 parts per million of oxygen,
preferably from about 0.2 to about 100 parts per million of oxygen.
This atmosphere may be a partial vacuum, or a mix of oxygen and an
inert gas such as argon. In another embodiment, the
oxygen-containing atmosphere is air at atmospheric pressure, which
contains about 21 percent by volume of oxygen and about 78 percent
by volume nitrogen.
[0029] The oxidation 56 is preferably performed at temperature of
at least about 1200.degree. F. to about 1550.degree. F., for a time
of at least about 2 hours, and in the oxidizing environment.
Preferably, the oxidation 56 is performed at a temperature of about
1300.degree. F. for a time of from about 8 to about 36 hours.
[0030] The result of the oxidation treatment 56 is the oxide layer
78. The oxide layer 78 is preferably at least about 500 Angstroms
thick, is preferably from about 500 Angstroms to about 6000
Angstroms thick, and is most preferably from about 1000 Angstroms
to about 3000 Angstroms thick. If the oxide layer 78 is thinner
than about 500 Angstroms, there is the possibility of incomplete
coverage and defects extending through the oxide layer 78. If the
oxide layer 78 is thicker than about 6000 Angstroms, there is an
increasing likelihood of spallation of the oxide layer 78 during
the thermal cycling that is associated with service of the turbine
component, with an associated shortening of the life of the
component. In the preferred form of the invention, that of FIG. 4
where steps 54 and 58 are not employed, this thin oxide layer 78
adds virtually no thickness or weight to the article that is
oxidized.
[0031] The oxide layer 78 predominantly comprises aluminum oxide
and/or chromium oxide, but it may also include other constituents
such as titanium oxide, nickel oxide, and cobalt oxide. It may also
contain nitrides such as titanium nitride, if the oxidation 56 is
performed in air. As noted earlier, the composition of the oxide
layer 78 depends upon the elements that are found at the surface 72
or 76 that is exposed during the oxidation. A feature of the
present approach is that the surface oxidation is not a
line-of-sight process, so that all portions of the surface 72 or 76
are covered and protected.
[0032] A top coating 80 (FIG. 5) may optionally thereafter be
applied over the oxide layer 78, numeral 58. The top coating 80 may
be of any operable type, such as aluminum oxide, tantalum oxide,
titanium oxide, silicon oxide, or chromium oxide. The top coating
80 may be applied by any operable technique, such as chemical vapor
deposition. The cop coating step 58 may be used with the approach
of FIG. 3, 4, or 5.
[0033] After the oxidation 56 and any of the optional steps 52, 54,
and 58, the component is placed into service, numeral 60. Only
then, during service, is the component exposed to the hot gases and
ingested dirt containing corrosive species such as sulfides and
sulfates. The placing into service is performed only after the
oxidation step 56 is complete and the substrate 70 is protected by
the oxide layer 78.
[0034] The present invention was reduced to practice using a Rene'
88DT substrate 70 in flat-panel tests and the embodiment of FIG. 4.
Specimens of the substrate 70 were given controlled oxidation
treatments (step 56) at 1300.degree. F. for times of 8, 12, 16, 24,
and 48 hours in air. For comparison, other specimens were not
oxidized in this manner at all, and other specimens were given a
simulated first engine service cycle in an engine environment. The
specimens were then tested in an accelerated corrosion test at
1300.degree. F. for 2 hours in an environment of sodium sulfite,
calcium sulfate, and carbon. The surfaces of the specimens were
inspected after every cycle. Failure was determined as the number
of cycles required to cause base metal pitting.
[0035] FIG. 6 presents the comparative corrosion test results. (The
nomenclature 1300/8 means a controllable oxidation at 1300.degree.
F. for 8 hours.) The specimen that was not controllably oxidized
and the specimen that was engine oxidized failed after about 3
cycles. Specimens given the controlled oxidation had corrosion test
lives ranging from about 6 to about 9 cycles, a significant
improvement.
[0036] Although particular embodiments of the invention have 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.
* * * * *