U.S. patent application number 12/146816 was filed with the patent office on 2008-10-23 for spar and shell constructed turbine blade.
This patent application is currently assigned to FLORIDA TURBINE TECHNOLOGIES, INC.. Invention is credited to Wesley D. Brown, Jack W. Wilson.
Application Number | 20080260538 12/146816 |
Document ID | / |
Family ID | 41717565 |
Filed Date | 2008-10-23 |
United States Patent
Application |
20080260538 |
Kind Code |
A1 |
Wilson; Jack W. ; et
al. |
October 23, 2008 |
Spar and shell constructed turbine blade
Abstract
A blade for a rotor of a gas turbine engine is constructed with
a spar and shell configuration. The spar is constructed in an
integral unit or multi-portions and includes a first wall adjacent
to the pressure side and a second wall adjacent to the suction
side, a tip portion extending in the spanwise direction and
extending beyond the first wall and the second wall and a root
portion extending longitudinally, an attachment portion having a
central opening for receiving the foot portion and a platform
portion. The root portion fits into the central opening and is
secured therein by a pin extending transversely through the
attachment and the foot portion. The shell fits over the spar and
is supported thereto by a plurality of complementary hooks
extending from the spar and, shell. The ends of the shell fit into
grooves formed on the tip .portion and the platform, The shell is
made from a high temperature resistant material, such as Molybdenum
or Niobium, and is formed from a wife EDM process.
Inventors: |
Wilson; Jack W.; (Palm Beach
Gardens, FL) ; Brown; Wesley D.; (Jupiter,
FL) |
Correspondence
Address: |
JOHN RYZNIC
FLORIDA TURBINE TECHNOLOGIES, INC., 1701 MILITARY TRAIL, SUITE 110
JUPITER
FL
33458-7887
US
|
Assignee: |
FLORIDA TURBINE TECHNOLOGIES,
INC.
Jupiter
FL
|
Family ID: |
41717565 |
Appl. No.: |
12/146816 |
Filed: |
June 26, 2008 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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11243308 |
Oct 4, 2005 |
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12146816 |
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|
10793641 |
Mar 4, 2004 |
7080971 |
|
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11243308 |
|
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60454120 |
Mar 12, 2003 |
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Current U.S.
Class: |
416/97R ;
416/226 |
Current CPC
Class: |
F01D 5/189 20130101;
Y10T 29/49327 20150115; Y10T 29/49341 20150115; Y10T 29/49339
20150115; F01D 5/20 20130101; F01D 5/147 20130101 |
Class at
Publication: |
416/97.R ;
416/226 |
International
Class: |
F01D 5/18 20060101
F01D005/18; F01D 5/14 20060101 F01D005/14 |
Claims
1. A shell for use in a turbine airfoil constructed from a spar and
shell; the shell comprising: The shell having an airfoil shape with
a leading edge and a trailing edge, and a pressure side and a
suction side extending between the two edges; The shell being
formed from an electric discharge machining process; and, The shell
being a thin shell wall to allow for near wall cooling of the inner
surface of the shell.
2. The shell of claim 1, and further comprising: The shell being
formed form a relatively high temperature resistant material that
cannot be cast or machined as a thin wall shell.
3. The shell of claim 2, and further comprising: The shell being
formed from Molybdenum or Niobium.
4. The shell of claim 1, and further comprising: The electric
discharge machining process is a wire electric discharge machining
process.
5. The shell of claim 1, and further comprising: The shell includes
at least one hook extending from the inner surface of the shell to
secure the shell to a spar of the turbine airfoil.
6. The shell of claim 5, and further comprising: The hook extends
along the spanwise length of the shell.
7. The shell of claim 5, and further comprising: The hook is also
formed from the electric discharge machining process and is formed
as a single piece along with the shell.
8. The shell of claim 1, and further comprising: The shell includes
a plurality of exit cooling holes in the trailing edge region of
the shell to discharge cooling air out from the shell.
9. The shell of claim 5, and further comprising: The; shell
includes a second hook in the same side of the shell as the first
hook; and, The two hooks face in opposite directions such that a
chordwise movement of the shell with respect to a spar is
limited.
10. The shell of claim 1, and further comprising: The shell has
substantially the same thickness from the platform end to the tip
end of the shell.
11. A turbine blade for use in a gas turbine engine the turbine
blade comprising: An attachment portion forming a platform and
having a central opening; A spar having a tip on one end; A shell
secured in place between the spar tip and the attachment portion;
and, The shell being formed form a relatively high temperature
resistant material that cannot be cast or machined as a thin wall
shell.
12. The turbine blade of claim 11, and further comprising: The
shell being a thin wall shell formed from an electric discharge
machining process.
13. The turbine blade of claim 12, and further comprising: The
electric discharge machining process is a wire electric discharge
machining process.
14. The turbine blade of claim 11, and further comprising: The
shell being formed from Molybdenum or Niobium.
15. The turbine blade of claim 11, and further comprising: The
shell and the spar both include hook means extending from the shell
and spar to secure the shell to a spar of the turbine airfoil.
16. The turbine blade of claim 11, and further comprising: The
shell has substantially the same thickness from the platform end to
the tip end of the shell.
17. The turbine blade of claim 11, and further comprising: The spar
having an internal cooling supply passage, a plurality of exit
cooling holes in the tip, and a plurality of near wall cooling
holes to discharge cooling air from the internal cooling supply
passage and onto the backside wall of the shell to provide near
wall cooling for the blade.
18. The turbine blade of claim 17, and further comprising: The
plurality of near wall cooling holes are located on the pressure
side and the suction side of the blade.
19. The turbine blade of claim 17, and further comprising: The
shell includes a row of trailing edge region exit cooling holes to
discharge cooling air from a space formed between the spar and the
shell.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application is claims the benefit to U.S. patent
application Ser. No. 11/243,308 filed on 10/04/2005 by Jack W.
Wilson, Jr. et al. and entitled TURBINE VANE WITH SPAR AND SHELL
CONSTRUCTION; which claims the benefit to U.S. patent application
Ser. No. 10/793,641 filed on Mar. 04, 2004 by Jack W, Wilson, Jr.
et al. and entitled COOLED TURBINE SPAR AND SHELL BLADE
CONSTRUCTION, now U.S. Pat. No. 7,080,971 B2 issued Jul. 25, 2006;
which claims the benefit to U.S. Provisional Patent Application
60/454,120 filed on Mar. 12, 2003, all of which are incorporated
herein by reference in their entirety.
BACKGROUND OF THE INVENTION
[0002] 1. Field of the Invention
[0003] The present invention relates generally to internally
cooled, turbine blades for a gas turbine engine, and more
specifically to a turbine blade made from a spar and shell
construction.
[0004] 2. Description of the Related Art including information
disclosed under 3.7 CFR 1.97 and 1.98
[0005] As one skilled in the gas turbine technology recognizes, the
efficiency of the engine is enhanced by operating the turbine at a
higher temperature and by increasing the turbine's pressure ratio.
Another feature that contributes to the efficacy of the engine is
the ability to cool the turbine with a lesser amount of cooling
air. The problem that prevents the turbine from being operated at
higher temperatures is the limitation of the structural integrity
of the turbine component parts that are jeopardized in its high,
temperature, hostile environment. Scientist and engineers have
attempted to combat the structural integrity problem by utilizing
internal cooling and selecting high temperature resistance
materials. The problem associated with internal cooling is twofold.
One, the cooling air that is utilized for the cooling comes from
the compressor that has already expended energy to pressurize this
air and the spent air in the turbine cooling process in essence is
a deficit in engine efficiency. The second problem is that the
cooling is through cooling passages and holes that are in the
turbine blade which, obviously, adversely affects the blade's
structural prowess. Because of the tortuous path that is presented
to the cooling air, the pressure drop that is a consequence thereof
requires higher pressure and more air to perform the cooling that
would otherwise take a lesser amount of air given the path becomes
less tortuous to the cooling air. While there are materials that
are available and can operate at a higher temperature that is
heretofore been used, the problem is how to harness these materials
so that they can be used efficaciously in the turbine
environment.
[0006] To better appreciate these problems it would be worthy of
note to recognize that traditional blade cooling approaches include
the use of cast nickel based alloys with load-bearing walls that
are cooled with radial flow channels and re-supply holes in
conjunction with film discharge cooling holes. Example of these
types of blades are exemplified by the following patents that are
incorporated herein by reference. U.S. Pat. No. 4,257,737 granted
to D. E. Andress et al on Mar. 24,1981 entitled "Cooled Rotor
Blade"; U.S. Pat. No. 4,753,575 granted to J. L. Levengood et al on
Jun. 28, 1988 entitled "Airfoil with Nested Cooling Channels"; U.S.
Pat. No. 5,476,364 granted to R. J. Kildea on Dec. 19, 1995
entitled "Tip Seal and Anti-Contamination for Turbine Blades"; and
U.S. Pat. No. 5,700,131 granted to Hall et al on Dec. 23, 1997
entitled "Cooled Turbine Blades for a Gas Turbine Engine".
[0007] Also well known by those skilled in this technology is that
the engine's efficiency increases as the pressure ratio of the
turbine, increases and the weight of the turbine decreases.
Needless to say these parameters have limitations. Increasing the
speed of the turbine also increases me airfoil loadings and, of
course, satisfactory operation of the turbine is to stay within
given airfoil loadings. The airfoil loadings are governed by cross
sectional area of the airfoil of the turbine multiplied by the
velocity of the tip of the turbine squared. Obviously, the
rotational speed of the turbine has a significant impact oh the
loadings.
[0008] The spar/shell construction contemplated by this invention
affords the turbine engine designer the option of reducing the
amount of cooling air that is required in any given engine design
and in addition, allowing the designer to fabricate the shell from
exotic high temperature materials that heretofore could not be
cast, or forged to define the surface profile of the airfoil
section. In other words, by virtue of this invention, the skin can
be made from Niobium or Molybdenum or (heir alloys, where the shape
is formed by a well known electric discharge process (EDM) or a
wire EDM process. In addition, because of the efficacious cooling
scheme of this invention, the shell portion could be made from
ceramics, or more conventional materials and still present an
advantage to me designer because a lesser amount of cooling air
would be required.
BRIEF SUMMARY OF THE INVENTION
[0009] An object of this invention is to provide a turbine rotor
for a gas turbine engine that is constructed, with in a spar/shell
configuration.
[0010] Another object of the present invention to provide for a
turbine blade that can make use of higher temperature materials
than are presently used in turbine blades.
[0011] Another object of the present invention is to form a shell
from, a high temperature resistant material that cannot be cast or
machined into a thin wall airfoil.
[0012] A feature of this invention is an inner spar that extends
from the root of the blade to the tip and is joined to me
attachment at the root by a pin or rod or the like.
[0013] Another feature of this invention is that the shell and/or
spar can be constructed from a high temperature material such as
ceramics, Molybdenum or Niobium (columbium) or a lesser temperature
resistive material such as Inco 718, Waspaloy or the well known
single crystal material currently being used in gas turbine
engines. For existing types of engine designs where it is desirable
of providing efficacious turbine blade cooling with the use of
compressor air at lower amounts and obtaining the same degree of
cooling. For advanced engine designs where it is desirable to
utilize more exotic materials such as Niobium or Molybdenum the
shell and spar can be made out of these materials or the spar can
be made from a lesser exotic material that is more readily cast or
forged.
[0014] The material of the shell may be taken from a group
consisting of stainless steel, molybdenum, niobium, ceramics,
molybdenum alloys, or niobium alloys. The material of the spar may
be taken from a group consisting of stainless steel, molybdenum,
niobium, ceramics, molybdenum alloys, or niobium alloys.
[0015] Another feature of this invention for engine designs that
require higher turbine rotational speeds, the spar can be made form
a dual spar system where the outer spar extends a shorted distance
radially relative to the inner spar and defines at the junction a
mid span shroud and the shell is formed in an upper section and a
lower section where each section is joined at the mid span shroud.
The pin in this arrangement couples the inner spar and outer spar
at the attachment formed at the root of the blade. This design can
utilized the same materials that are called out in the other
design.
[0016] A feature of this invention is an improved turbine blade
that is characterized, as being easy to fabricate, provide
efficacious cooling with lesser amounts of cooling air than
heretofore known designs, provides a shell or shells that can be
replaced and hence affords the user the option of repair or
replace. The materials selected can be conventional, or more
esoteric depending on the specification of the engine.
[0017] The foregoing and other features of the present invention
will become more apparent from me following description and
accompanying drawings.
[0018] BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS
[0019] FIG. 1 is an exploded view in perspective showing the
details of one embodiment of this invention.
[0020] FIG. 2 is a perspective-view illustrating the assembled
turbine blade of the embodiment depicted in FIG. 1 of this
invention.
[0021] FIG; 3 is a section taken from sectional lines 3-3 of FIG.
2.
[0022] FIG. 4 is a section taken along the sectional lines 4-4 of
FIG. 3 illustrating the attachment of the shell to the strut of
this invention.
[0023] FIG. 5 is a perspective view illustrating a second
embodiment of this invention.
[0024] FIG. 6 is a section view in elevation taken along the
sectional lines of 6-6 of FIG. 5.
DETAILED DESCRIPTION OF THE INVENTION
[0025] While this invention is described in its preferred
embodiment in two different, but similar configurations so as to
take advantage of engine's that are designed at higher speeds than
are heretofore encountered, this invention has the potential of
utilizing conventional materials and improving the turbine rotor by
enhancing its efficiency by providing the desired cooling with a
lesser amount of compressor air, and affords the designer to
utilize a more exotic material that has higher resistance
temperatures while also maintaining, the improved cooling aspects.
Hence, it will be understood to one skilled in this technology, the
material selected for the particular engine design is a option left
open to the designer while still employing the concepts of this
invention. For the sake of simplicity and convenience only a single
blade in each of the embodiments is described although one skilled
in this art that the turbine rotor consists of a plurality of
circumferentially spaced blades mounted in a rotor disk that makes
up the rotor assembly.
[0026] This disclosure is divided into two embodiments employing
the same concept of a spar and shell configuration of a turbine
blade, where one of the embodiments includes a single spar and the
other embodiment includes a double spar to accommodate higher
turbine rotational speeds. FIGS. 1 through 4 are directed to one of
the embodiments of a turbine blade generally illustrated as
reference numeral 10 as comprising a spar generally elliptical
shaped spar 12 extending longitudinally or in the radial direction
from the root portion 14 to the tip 16 with a downwardly extending
portion 18 that fairs into a rectangular shaped projection 26 that
is adapted to fit into the attachment 20. The spar 12 spans the
camber stations extending along the airfoil section defined by the
shell 28. The attachment 20 may include a fir tree attachment
portion 22 that fits into a complementary fir tree slot formed in
the turbine disk (not shown). The attachment 20 maybe formed with
the platform 24 or the platform may be formed separately and joined
thereto and projects in the circumferential direction to abut
against the platform in the adjacent blade in the turbine disk. A
seal, such as a feather seal (not shown), may be mounted between
platforms of adjacent blades to minimize or eliminate leakage
around the individual blades.
[0027] The spar may be formed as a single unit or may be made up in
complementary parts and as for example it may be formed in two
separate portions that are joined at the parting plane along the
leading edge facing portion 30 and trailing edge facing portion 32
and extending the longitudinal axis 31. Spar 12 is attached to the
attachment 20 by the pin 34 which fits through the hole 29 in the
attachment 20 and the aligned hole 31 formed, in the extending
portion 18. Pin 34 carries the head 36 that abuts against the face
38 of the attachment 20 and includes the flared put portion 40 at
the opposing end of head 36. This arrangement secures the spar 12
and assures that the load on the blade 10 is transmitted from the
airfoil section though the attachment 20 to the disk (not shown).
The tip of blade may be sealed by a cap 44 that may be formed
integrally with the spar 12 or may be a separate piece that is
suitably joined to the top end of the spar 12. It should be
appreciated that this design can accommodate a squealer cap, if
such is desired. The material of the spar will be predicated on the
usage of the blade and in a high temperature environment the
material can be a molybdenum or niobium and in a lesser temperature
environment the material can be a stainless steel like Inco 718 or
Waspaloy or the like.
[0028] Shell 48 extends over the surface of the spar 12 and is
hollow in the central portion 50 and spaced from the outer surface
of spar 12. The shell defines the pressure side 52, the suction
side 54, the leading edge 56 and the trailing edge 58. As mentioned
in the above paragraph the shell 48 may be made from different
materials depending on the specification of the gas turbine engine.
In the higher temperature requirements, the shell preferably will
be made from Molybdenum or Niobium and in a lesser temperature
environment the shell 48 maybe made from conventional materials. If
the material selected cannot be cast or forged then the shell will
be made, from a blank and the contour will be machined by a wire
EDM process. The shell can be made in a single unit or can be made
into two halves divided along the longitudinal axis, similar to the
spar 12. As best seen in FIG. 1, the attachment 20 is made to
include a stud portion 88 that complements the contoured surface of
spar 12 and the contoured surface of shell 48. Additionally the
shell 48 and spar 12 carry complementary male and female hooks 60
and 62. The top edge 80 of shell 48 is supported by the cap 44 and
fits into an annular groove 82 so that the upper edge 84 of shell
48 bears against the shoulder 86. The lowered edge 88 fits into an
annular complementary groove 90 formed on the upper edge of
platform 24 and bears against the opposing surfaces of the groove
90 and the outer surface of the attachment 20.
[0029] As mentioned in the above paragraphs, one of the important
features of this invention is that it affords efficacious cooling,
i.e. cooling that requires a lesser amount of air. This can be
readily seen by referring to FIG. 3. As shown the cooling air is
admitted through the inlet 66, the central, opening formed in the
spar 12 at the bottom face 68 of the attachment 20, and flows in a
straight passage or cavity 70 without having to flow through,
tortuous paths. The air that is admitted into cavity 70 flows out
of the feed holes 72 into the space or cavity 74 defined between
the spar 12 and the shell 48. Again, there are virtually no
tortuous passages that are typically found, in heretofore known
designs and hence the pressure drop is decreased requiring lesser
amount of air at a lower pressure, all of which enhances the
cooling efficiency of the blade. The air from the feed holes 72,
which may be formed integrally in the spar or drilled therein, can
serve to impinge on the inner wall of the shell 48 but primarily
feeds the space 74. It should be understood that this design can
include film cooling holes (as for example holes 71 and 73) formed
in the shell 48 on both the pressure surface 52 and the suction
surface 54 and may also include a shower head (depicted as holes
75) on the leading edge and codling holes (depicted as 77) on the
trailing edge 58. The design, and number of all of these cooling
holes i.e. shower head, film cooling, feed holes and the like are
predicated on the particular specification of the engine.
[0030] The other embodiment depicted in FIGS. 5 and 6 is similarly
constructed and is adapted to handle a higher rotational speed of
the turbine. In this embodiment, the shell 104 that is equivalent
to shell 48 depicted in FIGS. 14 is formed into two halves, the
upper halve 106 and the lower halve 108 and the attachment 110 that
is equivalent to the attachment 20 is extended in the longitudinal
and upwardly direction to extend almost midway along the airfoil
portion of the blade to form, another spar 112. This spar 112
surrounds the lower portion 114 of spar 12 (like numerals in all
the Figs. depict like or similar elements) and is contiguous
thereto along its inner surface. A ledge or platen 116 is formed
integrally therewith at the top end and extends in the spanwise
direction. Shell 106 and shell 108 are formed in an elliptical-like
shape to define the airfoil for defining the pressure surface 52,
suction surface 54, leading edge 56 and trailing edge 58. A groove
115 formed at the upper edge 117 of shell 106 bears against the
outer edge 118 of cap 120 which is the equivalent to cap 16 of
FIGS. 1 3 except it is a squealer cap. Obviously, when the blade is
rotating the shell 106 is loaded against the cap 120 and this force
is transmitted to the disk via the spar 12 and lower portion 114.
The lower edge 122 bears against the platen 116 and can be suitably
attached thereto by a suitable braze or weld. The lower shell 108
is similarly formed like shell 106 and defines the lower portion of
the airfoil. Lower shell 108 includes the groove 130 formed in the
increased diameter portion 132 of shell 108 and serves to receive
the outer edge 134 of platen 116. The lower edge 136 of shell 108
fits into an annular groove 138 formed in the platform 24. While
not shown in these Figs. the male and female hooks associated with
the spar and shell is also utilized in this embodiment and this
portion of the drawings are incorporated herein by reference. The
stud is like the embodiment depicted in FIGS. 1 3 is affixed to the
attachment via pin 34.
[0031] The cooling arrangement of the embodiment depicted in FIGS.
5 and 6 is almost identical to the cooling configuration of the
embodiment depicted in FIGS. 1 4. The only difference is that since
the platen 116 forms a barrier between the upper shell 106 and
lower shell 108, the cooling air to the lower portion of the
airfoil is directed from the inlet 66 and passage 70 via the radial
spaced holes 150 consisting of the aligned holes in the spars 12
and lower portion 114 that feeds space 156, and the holes 152
formed in the upper portion of the spar 12 that feed the space 158.
As is the case with the embodiment of FIGS. 1 4, the shell may
include a shower head at the leading edge, cooling passages at the
trailing edge, holes at the tip for cooling and discharging dirt
and foreign particles in the coolant and film cooling holes at the
surface of the pressure side and suction side.
[0032] Although this invention has been shown and described with
respect to detailed embodiments thereof, it will be appreciated and
understood by those skilled in the art that various changes in form
and detail thereof may be made without departing from the spirit
and scope of the claimed invention.
* * * * *