U.S. patent application number 11/150929 was filed with the patent office on 2006-12-14 for turbine blade and method of fabricating same.
Invention is credited to Michael Beverley, Francis Bobie, John Peter Heyward.
Application Number | 20060280610 11/150929 |
Document ID | / |
Family ID | 36954681 |
Filed Date | 2006-12-14 |
United States Patent
Application |
20060280610 |
Kind Code |
A1 |
Heyward; John Peter ; et
al. |
December 14, 2006 |
Turbine blade and method of fabricating same
Abstract
A method for fabricating a turbine blade includes casting at
least one turbine blade that includes a blade tip shroud and at
least one seal tooth coupled to the blade tip shroud, and removing
at least a portion of the turbine blade such that a radially outer
surface of the blade tip shroud has at least two substantially flat
surfaces.
Inventors: |
Heyward; John Peter;
(Loveland, OH) ; Beverley; Michael; (West Chester,
OH) ; Bobie; Francis; (Hamilton, OH) |
Correspondence
Address: |
JOHN S. BEULICK (12729);C/O ARMSTRONG TEASDALE LLP
ONE METROPOLITAN SQUARE
SUITE 2600
ST. LOUIS
MO
63102-2740
US
|
Family ID: |
36954681 |
Appl. No.: |
11/150929 |
Filed: |
June 13, 2005 |
Current U.S.
Class: |
416/189 |
Current CPC
Class: |
C22C 14/00 20130101;
F05D 2240/80 20130101; F05D 2220/3215 20130101; Y02T 50/60
20130101; F05D 2250/29 20130101; F01D 5/225 20130101; F05D 2230/21
20130101; Y02T 50/671 20130101; F05D 2230/12 20130101; F05D 2250/19
20130101; F05D 2250/70 20130101 |
Class at
Publication: |
416/189 |
International
Class: |
B63H 1/16 20060101
B63H001/16 |
Claims
1. A method for fabricating a turbine blade for a gas turbine
engine, wherein the gas turbine engine includes a turbine including
a plurality of turbine blades, said method comprising: casting at
least one turbine blade that includes a blade tip shroud and at
least one seal tooth coupled to the blade tip shroud; and removing
at least a portion of the turbine blade such that a radially outer
surface of the blade tip shroud has at least two substantially flat
surfaces.
2. A method in accordance with claim 1 wherein removing at least a
portion of the turbine blade further comprises removing at least a
portion of the turbine blade such that a radially outer surface of
the blade tip shroud has substantially V-shaped cross-sectional
profile.
3. A method in accordance with claim 1 wherein casting a turbine
blade further comprises casting a turbine blade utilizing a
titanium aluminide material.
4. A method in accordance with claim 1 further comprising machining
the turbine blade such that an outer surface of the at least one
seal tooth has a substantially V-shaped cross-sectional
profile.
5. A method in accordance with claim 4 wherein the turbine blade
includes two seal teeth, said method further comprising machining
the turbine blade such that an outer surface of each seal tooth has
a substantially V-shaped cross-sectional profile.
6. A method in accordance with claim 1 wherein casting a turbine
blade further comprises casting a low pressure turbine blade
utilizing a titanium aluminide material.
7. A method in accordance with claim 1 wherein removing at least a
portion of the turbine blade further comprises removing at least a
portion of the turbine blade utilizing an electron discharge
machining (EDM) apparatus.
8. A turbine rotor blade comprising: a blade tip shroud comprising
a radially outer surface having at least two substantially flat
surfaces; and at least one seal tooth coupled to said blade tip
shroud.
9. A turbine rotor blade in accordance with claim 8 wherein said
blade tip shroud has a substantially V-shaped cross-sectional
profile.
10. A turbine rotor blade in accordance with claim 8 wherein each
blade tip shroud flat surface spans approximately half the
circumferential width of said blade tip shroud.
11. A turbine rotor blade in accordance with claim 8 further
comprising two seal teeth, wherein each seal tooth comprises a
radially outer surface having at least two substantially flat
surfaces.
12. A turbine rotor blade in accordance with claim 8 further
comprising two seal teeth, wherein each seal tooth has a
substantially V-shaped cross-sectional profile.
13. A turbine rotor blade in accordance with claim 8 wherein said
turbine blade comprises a titanium aluminide material.
14. A turbine rotor blade in accordance with claim 8 wherein said
turbine blade is coupled to a low pressure turbine.
15. A gas turbine engine rotor assembly comprising: a low pressure
turbine rotor; and a plurality of circumferentially-spaced rotor
blades coupled to said low pressure turbine rotor, each said rotor
blade comprising: a blade tip shroud comprising a radially outer
surface having at least two substantially flat surfaces; and at
least one seal tooth coupled to said blade tip shroud.
16. A gas turbine engine rotor assembly in accordance with claim 15
wherein said blade tip shroud has a substantially V-shaped
cross-sectional profile.
17. A gas turbine engine rotor assembly in accordance with claim 15
wherein each blade tip shroud flat surface spans approximately half
the circumferential width of said blade tip shroud.
18. A gas turbine engine rotor assembly in accordance with claim 15
wherein said turbine blade further comprises two seal teeth,
wherein each seal tooth comprises a radially outer surface having
at least two substantially flat surfaces.
19. A gas turbine engine rotor assembly in accordance with claim 15
wherein said turbine blade further comprises two seal teeth,
wherein each seal tooth has a substantially V-shaped
cross-sectional profile.
20. A gas turbine engine rotor assembly in accordance with claim 15
wherein said turbine blade comprises a titanium aluminide material.
Description
BACKGROUND OF THE INVENTION
[0001] This invention relates generally to turbine assemblies, and
more particularly, to methods and apparatus for fabricating gas
turbine engine rotor blades.
[0002] Gas turbine engines typically include a compressor, a
combustor, and at least one turbine. The compressor compresses air
which is mixed with fuel and channeled to the combustor. The
mixture is then ignited for generating hot combustion gases, and
the combustion gases are channeled to the turbine which extracts
energy from the combustion gases for powering the compressor, as
well as producing useful work to propel an aircraft in flight or to
power a load, such as an electrical generator.
[0003] The turbine includes a rotor assembly and a stator assembly.
The rotor assembly includes a plurality of rotor blades extending
radially outward from a disk. More specifically, each rotor blade
extends radially between a platform adjacent the disk and a blade
tip. A combustion gas flowpath through the rotor assembly is bound
radially inward by the rotor blade platforms, and radially outward
by a plurality of shrouds, wherein each shroud includes at least
one seal tooth.
[0004] At least one known gas turbine blade is fabricated utilizing
a nickel based alloy to produce a turbine blade that has a
substantially curved outer shroud. For example, the turbine blade
is cast with additional stock that is removed utilizing a simple
radial grind operation to produce the finished outer shroud and the
seal teeth. More specifically, the nickel based alloy is cast "to
size", i.e. the turbine blade does not require significant
machining to produce the finished turbine blade.
[0005] However, if the gas turbine blade is fabricated utilizing a
different material, for example, a titanium aluminide material,
casting a turbine blade that includes a substantially curved shroud
and at least one seal tooth, produces a turbine blade that has no
"as-cast" surfaces. More specifically, casting a turbine blade
utilizing titanium aluminide produces a turbine blade that has a
large amount of excess material relative to the final machined
configuration, shown in FIGS. 1 and 2. Therefore, at least two
machining operations are required to remove the excess material to
form the finished turbine blade having a curved outer shroud and
seal teeth, shown in FIG. 3. Accordingly, fabricating a gas turbine
blade having a curved outer shroud utilizing a material other than
nickel based alloy facilitates increasing the cost of fabricating
the turbine blade, thus increasing the cost of the blade to the
customer.
BRIEF SUMMARY OF THE INVENTION
[0006] In one aspect, a method for fabricating a turbine blade for
a gas turbine engine is provided. The gas turbine engine includes a
turbine including a plurality of turbine blades. The method
includes casting at least one turbine blade that includes a blade
tip shroud and at least one seal tooth coupled to the blade tip
shroud, and removing at least a portion of the turbine blade such
that a radially outer surface of the blade tip shroud has at least
two substantially flat surfaces.
[0007] In another aspect, a turbine rotor blade is provided. The
turbine rotor blade includes a blade tip shroud having a radially
outer surface including at least two substantially flat surfaces,
and at least one seal tooth coupled to said blade tip shroud.
[0008] In a further aspect, a gas turbine engine rotor assembly is
provided. The rotor assembly includes a low pressure turbine rotor,
and a plurality of circumferentially-spaced rotor blades coupled to
the low pressure turbine rotor. Each rotor blade includes having a
radially outer surface including at least two substantially flat
surfaces, and at least one seal tooth coupled to said blade tip
shroud.
BRIEF DESCRIPTION OF THE DRAWINGS
[0009] FIG. 1 is a side view of a known gas turbine blade during
initial fabrication;
[0010] FIG. 2 is a cross-sectional view of the known gas turbine
blade shown in FIG. 1 during initial fabrication;
[0011] FIG. 3 is a cross-sectional view of the known gas turbine
blade shown in FIGS. 1 and 2 after machining is completed;
[0012] FIG. 4 is a cross-sectional view of an exemplary gas turbine
engine;
[0013] FIG. 5 is a partial cut-away cross-sectional view of a rotor
assembly and a stator assembly that may be used with the gas
turbine engine shown in FIG. 4;
[0014] FIG. 6 is an exemplary method of fabricating a turbine blade
that can be utilized with the gas turbine shown in FIG. 4;
[0015] FIG. 7 is a side view of an exemplary gas turbine blade that
can be used with the gas turbine engine shown in FIG. 4 during an
initial fabrication stage;
[0016] FIG. 8 is a perspective view of exemplary gas turbine blade
shown in FIG. 7 after subsequent fabrication;
[0017] FIG. 9 is a side view of the gas turbine blade shown in FIG.
8; and
[0018] FIG. 10 is a cross-sectional view of the gas turbine blade
shown in FIG. 9.
DETAILED DESCRIPTION OF THE INVENTION
[0019] FIG. 4 is a schematic illustration of a gas turbine engine
10 including a fan assembly 12, a high pressure compressor 14, and
a combustor 16. Engine 10 also includes a high pressure turbine 18,
a low pressure turbine 20, and a booster 22. Fan assembly 12
includes an array of fan blades 24 extending radially outward from
a rotor disc 26. Engine 10 has an intake side 28 and an exhaust
side 30. In one embodiment, the gas turbine engine is a GE90
available from General Electric Company, Cincinnati, Ohio. In an
alternative embodiment, engine 10 includes a low pressure
compressor. Fan assembly 12, booster 22, and turbine 20 are coupled
by a first rotor shaft 31, and compressor 14 and turbine 18 are
coupled by a second rotor shaft 32.
[0020] In operation, air flows through fan assembly 12 and
compressed air is supplied to high pressure compressor 14 through
booster 22. The highly compressed air is delivered to combustor 16.
Hot products of combustion (not shown in FIG. 1) from combustor 16
drives turbines 18 and 20, and turbine 20 drives fan assembly 12
and booster 22 by way of shaft 31. Engine 10 is operable at a range
of operating conditions between design operating conditions and
off-design operating conditions.
[0021] FIG. 5 is a partial cut-away cross-sectional view of a
turbine rotor assembly 40 including a stator 42 that may be used
with gas turbine engine 10. In the exemplary embodiment, turbine
rotor assembly 40 can be utilized as a low pressure turbine with
gas turbine engine 10. In the exemplary embodiment, rotor assembly
40 includes a plurality of rotors 44 joined together by couplings
46 co-axially about an axial centerline axis (not shown ). Each
rotor 44 is formed by one or more blisks 48, and each blisk 48
includes an annular radially outer rim 50, a radially inner hub 52,
and an integral web 54 extending radially therebetween. Each blisk
48 also includes a plurality of blades 56 extending radially
outwardly from outer rim 50. In one embodiment, blades 56 are
integrally joined with respective rims 50. Alternatively, and for
at least one stage, each rotor blade 56 may be removably joined to
a respective rim 50 in a known manner using blade dovetails (not
shown) which mount in complementary slots (not shown) in a
respective rim 50.
[0022] Rotor blades 56 are configured for cooperating with a motive
or working fluid, such as air. In the exemplary embodiment, rotor
assembly 40 is a turbine, such as low pressure turbine 20 (shown in
FIG. 4), with rotor blades 56 configured for suitably directing the
motive fluid air in succeeding stages. Outer surfaces 58 of rotor
rims 50 define a radially inner flowpath surface of turbine 20 as
air flows from stage to stage.
[0023] Blades 56 rotate about the axial centerline axis up to a
specific maximum design rotational speed, and generate centrifugal
loads in rotating components. Centrifugal forces generated by
rotating blades 56 are carried by portions of rims 50 directly
below each rotor blade 56. Rotation of rotor assembly 40 and blades
56 extracts energy from the air which causes turbine 20 to rotate
and provide power to drive low pressure compressor 12 (shown in
FIG. 1). The radially inner flowpath is bound circumferentially by
adjacent rotor blades 56 and is bound radially with a shroud
80.
[0024] Rotor blades 56 each include a leading edge 60, a trailing
edge 62, and an airfoil 64 extending therebetween. Each airfoil 64
includes a suction side 76 and a circumferentially opposite
pressure side 78. Suction and pressure sides 76 and 78,
respectively, extend between axially spaced apart leading and
trailing edges 60 and 62, respectively and extend in radial span
between a rotor blade tip shroud 80 and a rotor blade root 82. A
blade chord is measured between rotor blade trailing and leading
edges 62 and 60, respectively. In the exemplary embodiment, rotor
blades 56 include rotor seal teeth 86 which rotate adjacent to a
stator shroud 88 and through a cavity 89 defined by stator shroud
88 and rotor blade tip shroud 80.
[0025] FIG. 6 is a flowchart illustrating an exemplary method 100
of fabricating a gas turbine rotor blade 110 that can be utilized
with a gas turbine engine such as, but not limited to gas turbine
10, shown in FIG. 4. In the exemplary embodiment, method 100
includes casting 102 at least one turbine blade 110 that includes a
blade tip shroud 112 and at least one seal tooth 114 coupled to
blade tip shroud 112, and removing 104 at least a portion 116 of
turbine blade 110 such that a radially outer surface 118 of blade
tip shroud 112 has a substantially V-shaped cross-sectional profile
120.
[0026] FIG. 7 is a side view of gas turbine blade 110 during an
initial fabrication stage. FIG. 8 is a perspective view of gas
turbine blade 110 shown in FIG. 7 after subsequent fabrication.
FIG. 9 is a side view of gas turbine blade 110 shown in FIG. 8.
FIG. 10 is a cross-sectional view of gas turbine blade 110 shown in
FIG. 9. In the exemplary embodiment, turbine blade 110 is a
low-pressure turbine rotor blade that is coupled to a low pressure
turbine such as, low pressure turbine 20, shown in FIG. 4. In
another embodiment, method 100 can be utilized to fabricate any
turbine blade utilized within gas turbine engine 10, shown in FIG.
4.
[0027] In the exemplary embodiment, gas turbine blade 110 is cast
utilizing a metallic material. In the exemplary embodiment, gas
turbine blade 110 is cast utilizing a titanium aluminide material,
for example. In the exemplary embodiment, turbine blade 110 is
"overcast", that is turbine blade 110 includes a portion 116 that
must be removed to produce a finished blade 110.
[0028] In the exemplary embodiment, after turbine blade 110 is
cast, excess portion 116 is removed utilizing an electron discharge
machining (EDM) apparatus 122. More specifically, EDM apparatus 122
is configured to remove excess portion 116 such that blade tip
shroud 112 and at least one seal tooth 114 are formed, and such
that radially outer surface 118 of blade tip shroud 112 has a
substantially V-shaped cross-sectional profile 120.
[0029] More specifically, blade tip shroud 112 if formed such that
blade tip shroud outer surface 118 includes a first non-arcuate
portion 130, i.e. a flat, and a second non-arcuate portion 132, or
flat, that are coupled together at an apex 134 such that outer
surface 118 has a substantially V-shaped cross-sectional profile
120. In the exemplary embodiment, first and second portions 130 and
132 are formed unitarily with turbine blade 110. More specifically,
EDM apparatus 122 is utilized to remove portion 116 such that outer
surface 118 has a substantially V-shaped cross-sectional
profile.
[0030] Moreover, each seal tooth 114 includes a first non-arcuate
portion 140, or flat, and a second non-arcuate portion 142, or
flat, that are coupled together at an apex 144 such each seal tooth
114 has a substantially V-shaped cross-sectional profile 120. In
the exemplary embodiment, first and second portions 140 and 142 are
formed unitarily with turbine blade 110. More specifically, EDM
apparatus 122 is utilized to remove portion 116 such that each seal
tooth 114 has a substantially V-shaped cross-sectional profile
120.
[0031] The methods described herein facilitate improving the form
of the seal teeth and the shroud non-flowpath surface of a gas
turbine blade such that the turbine blade can be fabricated
utilizing a wire cut EDM apparatus. More specifically, known EDM
machines can not produce curved surfaces in the circumferential
direction of the turbine blade. Accordingly, the turbine blade
described herein does not include the curved form of the shroud and
seal teeth of a known turbine blade, rather the arched or curved
radially outer surface of the outer shroud is replaced by two flat
surfaces, wherein each flat surface spans half the circumferential
width of the shroud. Fabricating a turbine blade, that includes an
outer shroud having a radially outer surface that is formed by two
flat surfaces, facilitates reducing leakage between the seal teeth
and the stator shroud. Moreover, although the exemplary embodiment,
illustrates an outer shroud and seal teeth each having two
substantially flat surfaces, it should be realized that the
exemplary turbine blade can include three or more flat surface that
define the radially outer surface of the outer shroud and seal
teeth.
[0032] Accordingly, the methods and apparatus described herein
facilitate reducing the amount of machining that is needed to
produce LPT Blades from TiAl material, thus reducing cost and cycle
time to fabricate a turbine blade. Moreover, the reduced cost and
cycle time to fabricate the turbine blade described herein also
facilitates reducing the overall weight of the gas turbine
engine.
[0033] The above-described methods and apparatus are cost-effective
and highly reliable. The turbine blade described herein includes a
radially outer surface and at least one seal tooth that each
include at least two substantially flat surfaces to facilitate
reducing the time and cost of fabricating a gas turbine engine.
[0034] While the invention has been described in terms of various
specific embodiments, those skilled in the art will recognize that
the invention can be practiced with modification within the spirit
and scope of the claims.
* * * * *