U.S. patent application number 12/870315 was filed with the patent office on 2011-11-17 for gas turbine engine compressor components comprising thermal barriers, thermal barrier systems, and methods of using the same.
Invention is credited to Brian Thomas Hazel, Eric Scott Huron, DOUGLAS GERARD KONITZER.
Application Number | 20110280716 12/870315 |
Document ID | / |
Family ID | 44487201 |
Filed Date | 2011-11-17 |
United States Patent
Application |
20110280716 |
Kind Code |
A1 |
KONITZER; DOUGLAS GERARD ;
et al. |
November 17, 2011 |
GAS TURBINE ENGINE COMPRESSOR COMPONENTS COMPRISING THERMAL
BARRIERS, THERMAL BARRIER SYSTEMS, AND METHODS OF USING THE
SAME
Abstract
Gas turbine engine compressor disks having a hot flowpath side;
a shaft having a first surface positioned in the hot flowpath side;
and a thermal barrier applied to at least the first surface of the
shaft where the thermal barrier is operable to maintain the
temperature of the shaft below about 700.degree. C. (1300.degree.
F.) when the hot flowpath side experiences a service operating
temperature of from about 700.degree. C. (1300.degree. F.) to about
788.degree. C. (1450.degree. F.).
Inventors: |
KONITZER; DOUGLAS GERARD;
(Cincinnati, OH) ; Huron; Eric Scott; (Cincinnati,
OH) ; Hazel; Brian Thomas; (Cincinnati, OH) |
Family ID: |
44487201 |
Appl. No.: |
12/870315 |
Filed: |
August 27, 2010 |
Related U.S. Patent Documents
|
|
|
|
|
|
Application
Number |
Filing Date |
Patent Number |
|
|
61345327 |
May 17, 2010 |
|
|
|
Current U.S.
Class: |
415/177 |
Current CPC
Class: |
F05D 2260/95 20130101;
F05C 2201/0463 20130101; F01D 5/06 20130101; F05D 2230/90 20130101;
Y02T 50/67 20130101; F05D 2300/611 20130101; Y02T 50/60 20130101;
Y02T 50/6765 20180501; F05C 2201/0466 20130101; F05D 2300/11
20130101 |
Class at
Publication: |
415/177 |
International
Class: |
F04D 29/58 20060101
F04D029/58 |
Claims
1. A gas turbine engine compressor disk comprising: a hot flowpath
side; a shaft having a first surface positioned in the hot flowpath
side; and a thermal barrier applied to at least the first surface
of the shaft wherein the thermal barrier is operable to maintain
the temperature of the shaft below about 700.degree. C.
(1300.degree. F.) when the hot flowpath side experiences a service
operating temperature of from about 700.degree. C. (1300.degree.
F.) to about 788.degree. C. (1450.degree. F.).
2. The disk of claim 1 comprising a thermal barrier selected from
the group consisting of thermal barrier coatings, metal heat
shields, and thermal blankets.
3. The disk of claim 2 wherein the thermal barrier coating
comprises yttria stabilized zirconia.
4. The disk of claim 2 wherein the metal heat shield comprises a
material selected from the group consisting of metals, metal
alloys, and metal superalloys based on nickel, cobalt, iron, and
combinations thereof.
5. The disk of claim 2 wherein the thermal blanket comprises a low
conductivity material.
6. A thermal barrier system for a hot flowpath side of a gas
turbine engine compressor shaft comprising: a thermal barrier
selected from the group consisting of thermal barrier coatings,
metal heat shields, and thermal blankets applied to at least a
first surface of the compressor shaft wherein the thermal barrier
is operable to maintain the temperature of the shaft below about
700.degree. C. (1300.degree. F.) when the hot flowpath side
experiences a service operating temperature of from about
700.degree. C. (1300.degree. F.) to about 788.degree. C.
(1450.degree. F.).
7. The system of claim 6 wherein the thermal barrier comprises a
thermal expansion of greater than about 8.times.10.sup.-6/.degree.
C.
8. The system of claim 7 wherein the thermal barrier comprises an
in-plane modulus of less than 5 Msi.
9. The system of claim 8 wherein the thermal barrier is capable of
providing a reduction in the service operating temperature of the
shaft of at least about 5.degree. C./mm of thermal barrier
thickness.
10. The system of claim 9 wherein the thermal barrier coating
comprises yttria stabilized zirconia.
11. The system of claim 9 wherein the metal heat shield comprises a
material selected from the group consisting of metals, metal
alloys, and metal superalloys based on nickel, cobalt, iron, and
combinations thereof.
12. The system of claim 9 wherein the thermal blanket comprises a
low conductivity material.
13. A method for reducing service temperature of operation of a gas
turbine engine compressor shaft comprising: providing a compressor
shaft having a first surface in a hot flowpath side and a second
surface in a cool cavity side; and applying a thermal barrier to at
least the first surface of the compressor shaft wherein the thermal
barrier is operable to maintain the temperature of the shaft below
about 700.degree. C. (1300.degree. F.) when the hot flowpath side
experiences a service operating temperature of from about
700.degree. C. (1300.degree. F.) to about 788.degree. C.
(1450.degree. F.).
14. The method of claim 13 comprising providing a compressor shaft
comprising a polycrystalline substrate selected from the group
consisting of metals, metal alloys, and metal superalloys based on
nickel, cobalt, iron, and combinations thereof.
15. The method of claim 14 comprising applying a thermal barrier
capable of providing a reduction in the service operating
temperature of the shaft of at least about 5.degree. C./mm of
thermal barrier thickness.
16. The method of claim 14 comprising applying a thermal barrier
having a thermal expansion of greater than about
8.times.10.sup.-6/.degree. C.
17. The method of claim 14 comprising applying a thermal barrier
having an in-plane modulus of less than 5 Msi.
18. The method of claim 17 wherein the thermal barrier is selected
from the group consisting of thermal barrier coatings, metal heat
shields, and thermal blankets.
19. The method of claim 18 wherein the thermal barrier coating
comprises yttria stabilized zirconia.
20. The method of claim 18 wherein the metal heat shield comprises
a material selected from the group consisting of metals, metal
alloys, and metal superalloys based on nickel, cobalt, iron, and
combinations thereof.
Description
CROSS REFERENCE TO RELATED APPLICATIONS
[0001] This Application claims priority to U.S. Provisional
Application Ser. No. 61/345,327, filed May 17, 2010, which is
herein incorporated by reference in its entirety.
TECHNICAL FIELD
[0002] Embodiments described herein generally relate to gas turbine
engine compressor components comprising thermal barriers; thermal
barrier systems; and methods of using the same. More particularly,
embodiments described herein generally relate to gas turbine engine
high pressure compressor shafts and thermal barrier systems for use
therewith.
BACKGROUND OF THE INVENTION
[0003] Increasingly stringent demands are being imposed on the
efficacy of gas turbine engines employed in the aerospace and power
generation industries. This demand is driven by the requirement to
reduce the consumption of fossil fuels, and in turn, operating
costs. One way to improve turbine efficiency is to increase the
operating temperature of the engine. For example, the high pressure
compressor currently operates at about 700.degree. C. (about
1300.degree. F.), with the service operating temperature expected
to reach about 788.degree. C. (about 1450.degree. F.) in the
future.
[0004] With increased operating temperatures comes an increased
demand on materials, such as those used to make compressor disks.
Not only must these materials be able to withstand the higher
operating temperatures, but they must also endure increased
mechanical stresses, corrosion, erosion, and other severe operating
conditions, while continuing to fulfill lifetime requirements
expected by the industry.
[0005] One way to address these increased demands is through
modification of both the composition and structure of the material
used to make the compressor component. More specifically, the
composition of the material can be modified to make it capable of
operating at the higher temperatures, while the structure of the
material, and in particular the grain size, can be increased or
decreased depending on whether creep or fatigue, respectively, is
the more critical property. However, even with such modifications,
there can still be challenges to getting the material to perform
properly at increasing temperatures.
[0006] Accordingly, there remains a need for additional measures to
protect these materials from the high temperature, high stress
environments of the compressor section of a gas turbine engine.
BRIEF DESCRIPTION OF THE INVENTION
[0007] Embodiments herein generally relate to gas turbine engine
compressor disks comprising: a hot flowpath side; a shaft having a
first surface positioned in the hot flowpath side; and a thermal
barrier applied to at least the first surface of the shaft wherein
the thermal barrier is operable to maintain the temperature of the
shaft below about 700.degree. C. (1300.degree. F.) when the hot
flowpath side experiences a service operating temperature of from
about 700.degree. C. (1300.degree. F.) to about 788.degree. C.
(1450.degree. F.).
[0008] Embodiments herein also generally relate to thermal barrier
systems for a hot flowpath side of a gas turbine engine compressor
shaft comprising: a thermal barrier selected from the group
consisting of thermal barrier coatings, metal heat shields, and
thermal blankets applied to at least a first surface of the
compressor shaft wherein the thermal barrier is operable to
maintain the temperature of the shaft below about 700.degree. C.
(1300.degree. F.) when the hot flowpath side experiences a service
operating temperature of from about 700.degree. C. (1300.degree.
F.) to about 788.degree. C. (1450.degree. F.).
[0009] Embodiment herein also generally relate to methods for
reducing service temperature of operation of a gas turbine engine
compressor shaft comprising: providing a compressor shaft having a
first surface in a hot flowpath side and a second surface in a cool
cavity side; applying a thermal barrier to at least the first
surface of the compressor shaft wherein the thermal barrier is
operable to maintain the temperature of the shaft below about
700.degree. C. (1300.degree. F.) when the hot flowpath side
experiences a service operating temperature of from about
700.degree. C. (1300.degree. F.) to about 788.degree. C.
(1450.degree. F.).
[0010] These and other features, aspects and advantages will become
evident to those skilled in the art from the following
disclosure.
BRIEF DESCRIPTION OF THE DRAWINGS
[0011] While the specification concludes with claims particularly
pointing out and distinctly claiming the invention, it is believed
that the embodiments set forth herein will be better understood
from the following description in conjunction with the accompanying
figures, in which like reference numerals identify like
elements.
[0012] FIG. 1 is a schematic representation of one embodiment of a
high pressure compressor disk in accordance with the description
herein; and
[0013] FIG. 2 is an enlarged representation of the compressor disk
of FIG. 1 having a thermal barrier applied in accordance with the
description herein.
DETAILED DESCRIPTION OF THE INVENTION
[0014] Embodiments described herein generally relate to gas turbine
engine compressor components comprising thermal barriers; thermal
barrier systems; and methods of using the same. More particularly,
embodiments described herein generally relate to gas turbine engine
high pressure compressor shafts and thermal barrier systems for use
therewith.
[0015] For purposes of the description herein, the turbine engine
compressor component may generally be of any type, however, in one
embodiment, the component may be a rotating compressor component,
or portion thereof, that experiences a service operating
temperature of from about 537.degree. C. (about 1000.degree. F.) to
about 788.degree. C. (about 1450.degree. F.), and in one embodiment
from about 700.degree. C. (about 1300.degree. F.) to about
788.degree. C. (about 1450.degree. F.). One example of a rotating
compressor component that may benefit from the methods and systems
described herein can include, but should not be limited to,
compressor disks. While the entire disk may be protected, it may be
more advantageous and cost effective to protect a selected portion
of the disk, for example, the HPC shaft, as described herein
below.
[0016] Referring to FIG. 1, a representative embodiment of a
turbine engine compressor component 10 is provided. As described
previously, compressor component 10 may be of any operable type,
such as a disk 12. As shown in FIG. 1 compressor disk 12 may have a
shaft 14. Shaft 14 can comprise a first surface 13 located in the
hot flowpath side 16 of compressor disk 12, and a second surface
15, located in the cool cavity side 17 of compressor disk 12.
During operation, the temperature difference between the hot
flowpath side 16 and the cool cavity side 17 can be from about
230.degree. C. to about 430.degree. C. (about 500.degree. F. to
about 700.degree. F.). During operation, and absent a thermal
barrier as set forth herein below, hot air from hot flowpath side
16 can heat the entire shaft 14, including both first surface 13
and second surface 15, to temperatures upwards of 700.degree. C.
(about 1300.degree. F.). Therefore, it may be advantageous to
utilize the thermal barriers herein to protect the entire shaft 14,
or alternately, the first surface 13 thereof. Using a thermal
barrier to protect first surface 13 of shaft 14 can help to
maintain the temperature of shaft 14 at a temperature of about
700.degree. C. (about 1300.degree. F.) even though the service
operating temperature of the hot flowpath side 16 of compressor
disk 12 can reach up to about 788.degree. C. (about 1450.degree.
F.).
[0017] Referring to FIG. 2, compressor shaft 14 can comprise any of
a variety of polycrystalline substrates such as metals, metal
alloys, or metal superalloys, including those based on nickel,
cobalt, iron, or combinations thereof. In one embodiment, shaft 14
may comprise a nickel-based superalloy. As used herein throughout,
"based" indicates that the metal substrate comprises a greater
percentage of the listed element(s) than any other element. For
example, a nickel-based superalloy comprises more nickel than any
other element. By way of example and not limitation, one
nickel-based superalloy that may be used herein is Rene.RTM.08
(General Electric Co.) which is generally comprised, by weight, of
about 13% cobalt, about 16% chromium, about 4% molybdenum, about
3.7% titanium, about 2.1% aluminum, about 4% tungsten, about 0.70%
niobium, about 0.015% boron, about 0.03% zirconium, and about 0.03%
carbon, with the balance nickel and minor impurities.
[0018] Thermal barrier 20, as shown in FIG. 2, can be applied to at
least a portion of shaft 14, and in particular to first surface 13
of shaft 14. Thermal barrier can comprise any of a variety of
materials that comprise the following properties: barrier 20 can be
capable of providing a temperature reduction in the metal substrate
of at least about 5.degree. C./mm of thermal barrier thickness, and
in one embodiment a reduction of at least about 7.degree. C./mm of
thermal barrier thickness, as compared to the temperature of the
substrate absent the application of thermal barrier 20; barrier 20
can remain affixed to the substrate without decreasing the
performance thereof; and, barrier 20 can have a strain tolerance
and thermal expansion similar to that of the substrate.
Specifically, in one embodiment, thermal expansion can be greater
than about 8.times.10.sup.-6/.degree. C., and in another
embodiment, from about 10.times.10.sup.-6/.degree. C. To
accommodate the strain from the centrifugal loading, thermal
barrier 20 can have an in-plane modulus of less than 5 Msi, and in
one embodiment less than 2 Msi. Some examples of suitable thermal
barriers 20 are set forth below.
[0019] In one embodiment, thermal barrier 20 can comprise a thermal
barrier coating (TBC). The TBC may generally comprise any low
conductivity material, or combination of materials, currently
suitable for use as a TBC, for example, a ceramic such as 7% yttria
stabilized zirconia. As used herein, "low conductivity" indicates
that the conductivity of the TBC is less than that of the substrate
of the shaft 14. A layer of the TBC may be applied to HPC shaft 14
using conventional application methods such as, but not limited to,
plasma spray processes, chemical vapor deposition processes,
electron beam physical vapor deposition processes, sputtering
processes, and the like. TBC can provide the desired temperature
reduction, adhesion, thermal expansion, and strain tolerance
properties set forth previously. Those skilled in the art will
understand how to tailor the thickness of the TBC to achieve the
previously described properties, as well as how to apply the TBC
without damaging the underlying component.
[0020] In another embodiment, thermal barrier 20 can comprise a
metal heat shield affixed to first surface 13 of shaft 14 on hot
flowpath side 16 as shown in FIG. 2. Similar to the compressor
component, the metal heat shield can be made from any metal, metal
alloy, or metal superalloy including those based on nickel, cobalt,
iron, or combinations thereof as defined previously, that is
capable of withstanding the temperature and stress experienced by
shaft 14 during operation. Some examples of suitable materials for
the heat shield can include nickel alloys, commonly available under
the tradenames HS 188, or HASTX (available from Haynes
International, for example). The metal heat shield can be attached
to shaft 14 by a variety of mechanical or chemical attachment
methods. Such placement of the heat shield can help achieve the
previously described properties.
[0021] In yet another embodiment, thermal barrier 20 can comprise a
thermal blanket. Similar to the TBC, the thermal blanket may
comprise any "low conductivity" material wherein the conductivity
of the thermal blanket material is less than that of the substrate
of the shaft 14. In this embodiment, the thermal blanket can be
applied to insulate the substrate of shaft 14 from hot flow path
air. In this way, shaft 14 can be kept cooler by backside cooling
air. In one embodiment, the thermal blanket can reduce the
temperature of shaft 14 by between about 55.degree. C. (100.degree.
F.) and about 83.degree. C. (150.degree. F.). Those skilled in the
art will understand how to tailor the thickness of the thermal
blanket to achieve the previously described properties
[0022] The thermal barriers described herein for application to HPC
shaft can reduce the amount of heat that reaches the shaft. This
reduction in heat can result in both reduced thermal stress on the
HPC shaft, as well as a lower service operating temperature of the
shaft. The addition of the thermal barrier can allow the rotating
compressor component to be fabricated from materials having a lower
temperature capability than used currently, without sacrificing
performance.
[0023] This written description uses examples to disclose the
invention, including the best mode, and also to enable any person
skilled in the art to make and use the invention. The patentable
scope of the invention is defined by the claims, and may include
other examples that occur to those skilled in the art. Such other
examples are intended to be within the scope of the claims if they
have structural elements that do not differ from the literal
language of the claims, or if they include equivalent structural
elements with insubstantial differences from the literal language
of the claims.
* * * * *