U.S. patent application number 09/681744 was filed with the patent office on 2002-12-05 for film cooled blade tip.
Invention is credited to Bunker, Ronald Scott.
Application Number | 20020182074 09/681744 |
Document ID | / |
Family ID | 24736597 |
Filed Date | 2002-12-05 |
United States Patent
Application |
20020182074 |
Kind Code |
A1 |
Bunker, Ronald Scott |
December 5, 2002 |
FILM COOLED BLADE TIP
Abstract
A turbine assembly having at least one rotor blade comprises an
airfoil having a pressure sidewall, a suction sidewall and a tip
portion having a tip cap. A tip is disposed on the tip cap. A
plurality of blade tip cooling holes are positioned within the
airfoil near the tip portion. Cooling grooves are disposed within
the airfoil to connect the blade tip cooling holes with the top
portion of the tip to transition cooling flow from the cooling
holes to the tip portion.
Inventors: |
Bunker, Ronald Scott;
(Niskayuna, NY) |
Correspondence
Address: |
GENERAL ELECTRIC COMPANY
GLOBAL RESEARCH CENTER
PATENT DOCKET RM. 4A59
PO BOX 8, BLDG. K-1 ROSS
NISKAYUNA
NY
12309
US
|
Family ID: |
24736597 |
Appl. No.: |
09/681744 |
Filed: |
May 31, 2001 |
Current U.S.
Class: |
416/97R |
Current CPC
Class: |
F01D 5/20 20130101; F01D
5/18 20130101 |
Class at
Publication: |
416/97.00R |
International
Class: |
F01D 005/18 |
Claims
1. A turbine assembly comprising: at least one rotor blade
comprising an airfoil having a pressure sidewall and a suction
sidewall defining an outer periphery and a tip portion having a tip
cap; a plurality of blade tip cooling holes disposed within said
airfoil adjacent to said tip portion; and at least one cooling
groove disposed within said airfoil connecting at least one of said
blade tip cooling holes with a top portion of said tip portion so
as to transition cooling flow from said cooling holes to said tip
portion.
2. A turbine assembly in accordance with claim 1, wherein said
blade tip film cooling holes are angled with respect to said
airfoil.
3. A turbine assembly in accordance with claim 1, wherein said
blade tip film cooling holes are angled in the range between about
2.degree. to about 7.degree. with respect to the surface of said
airfoil.
4. A turbine assembly in accordance with claim 1, wherein said
cooling grooves are disposed so as to have a substantially constant
width from said film cooling holes to said tip portion.
5. A turbine assembly in accordance with claim 1, wherein said
grooves are fan-type cooling grooves.
6. A turbine assembly in accordance with claim 1, wherein said
grooves are multiple-channel cooling grooves.
7. A turbine assembly in accordance with claim 1, further
comprising a pressure side winglet disposed upon an upper portion
of said airfoil, said winglet having a top portion contiguous with
said top portion of said tip and an angled body portion.
8. A turbine assembly in accordance with claim 7, wherein said
angled body portion is angled at substantially the same angle as
said film cooling holes.
9. A turbine assembly in accordance with claim 7, wherein said
angled body portion is positioned coextensively with a top portion
of a respective film cooling hole such that said top portion of
said film cooling hole and said angled body portion generally form
a straight line.
10. A turbine assembly in accordance with claim 7, wherein said
groove is disposed directly into a respective angled body portion
such that cooling flow issuing from a respective cooling hole flows
through said groove to a top portion of said winglet over said top
surface of said tip portion and on to said tip cap.
11. A turbine assembly in accordance with claim 7, wherein the
relative angle between said winglet and said film holes is between
about -15 to +15 degrees.
12. A turbine assembly in accordance with claim 1, wherein said
grooves are cast features of said blade tip.
13. A turbine assembly in accordance with claim 1, wherein said
grooves are machined into said blade tip after casting thereof.
14. A turbine assembly in accordance with claim 1, wherein said
grooves are formed by laser drilling said blade tip after casting
thereof.
15. A turbine assembly in accordance with claim 7, wherein said
winglet edge is rounded.
16. A turbine assembly in accordance with claim 7, wherein said
film cooling holes are contained within said winglet.
17. A turbine assembly in accordance with claim 1, wherein said tip
further includes a squealer tip.
18. A turbine assembly in accordance with claim 17, wherein said
squealer tip is a single-tooth squealer.
19. A turbine assembly in accordance with claim 18, wherein said
tip has a single-tooth squealer located approximately along a mean
chordline of said blade tip section.
20. A turbine blade comprising: an airfoil having a pressure
sidewall, a suction sidewall and a tip portion having a tip cap; a
plurality of blade tip cooling holes disposed within said airfoil
adjacent to said tip portion; and at least one cooling groove
disposed within said airfoil connecting at least one of said blade
tip cooling holes with a top portion of said tip so as to
transition cooling flow from said cooling holes to said tip
portion.
21. A turbine blade in accordance with claim 20, wherein said blade
tip film cooling holes are angled with respect to said airfoil.
22. A turbine blade in accordance with claim 20, wherein said blade
tip film cooling holes are angled in the range between about 20 to
about 70 with respect to the surface of said airfoil.
23. A turbine blade in accordance with claim 20, wherein said
cooling grooves are disposed so as to have a substantially constant
width from said film cooling holes to said tip portion.
24. A turbine blade in accordance with claim 20, wherein said
grooves are fan-type cooling grooves.
25. A turbine blade in accordance with claim 20, wherein said
grooves are multiple-channel cooling grooves.
26. A turbine blade in accordance with claim 20, further comprising
a pressure side winglet disposed upon an upper portion of said
airfoil, said winglet having a top portion contiguous with said top
portion of said tip and an angled body portion.
27. A turbine blade in accordance with claim 26, wherein said
angled body portion is angled at substantially the same angle as
said film cooling holes.
28. A turbine blade in accordance with claim 26, wherein said
angled body portion is positioned coextensively with a top portion
of a respective film cooling hole such that said top portion of
said film cooling hole and said angled body portion generally form
a straight line.
29. A turbine blade in accordance with claim 26, wherein said
groove is disposed directly into a respective angled body portion
such that cooling flow issuing from a respective cooling hole flows
through said groove to a top portion of said winglet over said top
surface of said tip portion and on to said tip cap.
30. A turbine blade in accordance with claim 26, wherein the
relative angle between said winglet and said film holes is between
about -15 to +15 degrees.
31. A turbine blade in accordance with claim 20, wherein said
grooves are cast features of said blade tip.
32. A turbine blade in accordance with claim 20, wherein said
grooves are machined into said blade tip after casting thereof.
33. A turbine blade in accordance with claim 20, wherein said
grooves are formed by laser drilling said blade tip after casting
thereof.
34. A turbine blade in accordance with claim 26, wherein said
winglet edge is rounded.
35. A turbine blade in accordance with claim 26, wherein said film
cooling holes are contained within said winglet.
36. A turbine blade in accordance with claim 20, wherein said tip
further includes a squealer tip.
37. A turbine blade in accordance with claim 36, wherein said
squealer tip is a single-tooth squealer.
38. A turbine blade in accordance with claim 37, wherein said tip
has a single-tooth squealer located approximately along a mean
chordline of said blade tip section.
Description
BACKGROUND OF INVENTION
[0001] The present invention relates generally to turbine engine
blades and, more particularly, to a turbine blade tip peripheral
end wall with a grooved cooling arrangement.
[0002] A reduction in turbine engine efficiency results from
leaking of hot expanding combustion gases in the turbine across a
gap between rotating turbine blades and stationary seals or shrouds
which surround the blades. The problem of sealing between such
relatively rotating members to avoid loss in efficiency is very
difficult in the turbine section of the engine because of high
temperatures and centrifugal loads.
[0003] One method of improving the sealing between a respective
turbine blade and shroud is the provision of squealer type tips on
turbine blades. A squealer tip includes a continuous peripheral end
wall of relatively small height typically surrounding and
projecting outwardly from an end cap on the outer end of a turbine
blade that encloses a cooling air plenum in the interior of the
blade.
[0004] During operation of the engine, temperature changes create
differential rates of thermal expansion and contraction on the
blade rotor and shroud that may result in rubbing between the blade
tips and shrouds. Centrifugal forces acting on the blades and
structural forces acting on the shrouds create distortions thereon
that may also result in rubbing interference.
[0005] Such rubbing interference between the rotating blade tips
and surrounding stationary shrouds causes heating of the blade tips
resulting in excessive wear or damage to the blade tips and
shrouds. Heating produced by the leakage flow of hot gases may
actually be augmented by the presence of a cavity defined by the
end cap and peripheral end wall of the squealer tip because of the
increased surface area of the peripheral end wall. The peripheral
end wall is especially difficult to cool, because the end wall
extends away from the internally cooled region of the blade.
Therefore, squealer type blade tips, though fostering improved
sealing, actually require additional cooling.
[0006] Because of the complexity and relative high cost of
replacing or repairing turbine blades, it is desirable to prolong
as much as possible the life of blade tips and respective blades.
Blade tip cooling is a conventional practice employed for achieving
that objective. The provision of holes for directing air flow to
cool blade tips is known in the prior art, for instance as
disclosed in U.S. Pat. No. 4,247,254 to Zelahy, and have been
applied to squealer type blade tips as disclosed in U.S. Pat. No.
4,540,339 to Horvath.
[0007] Turbine engine blade designers and engineers are constantly
striving to develop more efficient ways of cooling the tips of the
turbine blades to prolong turbine blade life and reduce engine
operating cost. Cooling air used to accomplish this is expensive in
terms of overall fuel consumption. Thus, more effective and
efficient use of available cooling air in carrying out cooling of
turbine blade tips is desirable not only to prolong turbine blade
life but also to improve the efficiency of the engine as well,
thereby again lowering engine operating cost. Consequently, there
is a continuing need for a cooling hole design that will make more
effective and efficient use of available cooling air.
SUMMARY OF INVENTION
[0008] A turbine assembly having at least one rotor blade comprises
an airfoil having a pressure sidewall, a suction sidewall, and a
tip portion having a tip cap. A squealer tip is disposed on the tip
cap. A plurality of blade tip cooling holes are positioned within
the airfoil near the tip portion. Cooling grooves are disposed
within the airfoil to connect the blade tip cooling holes with the
top portion of the squealer tip to transition cooling flow from the
cooling holes to the tip portion.
BRIEF DESCRIPTION OF DRAWINGS
[0009] FIG. 1 is a perspective view of a turbine blade having a
squealer tip with cooling holes through an end cap of the
blade;
[0010] FIG. 2 is a perspective view of a turbine blade having a
squealer tip and incorporating the cooling arrangement in
accordance with the present invention;
[0011] FIGS. 3-7 are fragmentary radial sectional views of the
turbine blade of FIG. 2 taken along line 3-3; and
[0012] FIGS. 8-10 are fragmentary longitudinal sectional views of
the turbine blade of FIG. 2 taken along line 4-4.
DETAILED DESCRIPTION
[0013] A turbine blade 10 includes an airfoil 12 having a pressure
side 14, a suction side 16, and a base 18 for mounting airfoil 12
to a rotor (not shown) of an engine (not shown) as shown in FIG. 1.
Base 18 has a platform 20 for rigidly mounting airfoil 12 and a
dovetail root 22 for attaching blade 10 to the rotor.
[0014] An outer end portion 24 of blade 10 has a tip 26. Tip 26
includes an end cap 28 which closes outer end portion 24 of blade
10, and an end wall 30 attached to, and extending along the
periphery 31 of, and projecting outwardly from, end cap 28 so as to
define a cavity 29 therewith. End cap 28 of tip 26 typically is
provided with an arrangement of tip cooling holes 32 formed
therethrough for permitting passage of cooling air flow from the
interior of blade 10 through end cap 28 to cavity 29 for purposes
of cooling blade tip 26.
[0015] The tip of a turbine blade is designed to serve many
purposes. One purpose is to maintain the blade integrity in the
event of rubbing between the blade tip and a stationary shroud (not
shown). A second purpose is to minimize the leakage flow across the
blade tip from the pressure side to the suction side and a third
purpose is to cool the blade tip within the material limit. Tip 26
provides the rubbing capability and also serves as a two-tooth seal
to discourage the leakage flow.
[0016] As shown in FIG. 1, at least one and typically a plurality
of blade tip film cooling holes 34 are disposed within outer end
portion 24 of airfoil 12. Typically, blade tip film cooling holes
34 provide external film cooling issued on the blade tip pressure
side 14 in the radial direction. Some designs use as many film
holes 34 as possible, in the limited space available, in an effort
to flood the pressures side tip region with coolant. It is desired
that this film cooling then carry over onto end wall 30 and into
cavity 29 to provide cooling there and also over the suction side
surfaces of tip 26. Film holes 34 are oriented in the radially
outward direction because the prevailing mainstream gas flows tend
to migrate in this manner in the tip region. In practice, it is
still very difficult and very inconsistent to cool the blade tip in
this manner due to the very complex nature of the cooling flow as
it mixes with very dynamic hot gases of the mainstream flow. Blade
tip film cooling holes 34 are typically angled with respect to the
surface of airfoil 12. In one embodiment, blade tip cooling holes
are angled in the range between about 20.degree. to about
70.degree. with respect to the surface of airfoil 12.
[0017] As shown in FIG. 1, hot air flows (generally illustrated as
arrows 36) over airfoil 12 and exerts motive forces upon the outer
surfaces of airfoil 12, in turn driving the turbine and generating
power. In some arrangements, cooling flow (generally illustrated by
arrows 38) exits film holes 34 and is swept by hot air flow 36
towards a trailing edge 40 of airfoil 12 and away from tip cap 28.
Typically, this results in a mixed effect, where some of the
cooling air is caught up and mixed with the hot gases and some goes
onto tip cap 28 and some goes axially along the airfoil to trailing
edge 40. This results in inadequate cooling of tip cap 28 and
endwall 30 and eventual temperature inflicted degradation of tip
cap 28 and endwall 30.
[0018] As shown in FIG. 2, hot air flow 36 passes over airfoil 12
and exerts motive forces upon the outer surfaces of airfoil 12,
driving the turbine and generating power. In accordance with one
embodiment of the instant invention, at least one and typically a
plurality of grooves 50 are disposed within outer portion of
airfoil 12 connecting at least one corresponding blade tip film
cooling hole 34 with top portion of the airfoil to transition
cooling flow 38 from blade tip film cooling holes 34 to tip cap 28
and to end wall 30.
[0019] As shown, in an exploded view of FIG. 2, cooling grooves 50
can be disposed so as to have a substantially constant width from
film cooling holes 34 to tip cap 28, as indicated by reference
numeral 80. Alternatively, a fan-type cooling groove 50 can be
utilized to spread the cooling air 30 as it exits film cooling
holes 34, as indicated by reference numeral 82. Also, a
multiple-channel cooling groove 50 can be utilized, as indicated by
reference numeral 84.
[0020] In one embodiment, airfoil 12 further comprises a pressure
side winglet 54 disposed upon an upper portion of airfoil 12, as
best shown in FIG. 3. Pressure side winglet 54 includes a top
portion 56 contiguous with top surface 52 of tip 26 and an angled
body portion 58.
[0021] Angled body portion 58 is typically angled at the same angle
as film cooling hole 34 in reference to the surface of airfoil 12.
In one embodiment, angled body portion 58 is positioned
coextensively with a top portion of a respective film cooling hole
34 such that the top portion of film cooling hole 34 and angled
body portion 58 generally form a straight line. In one embodiment,
groove 50 is disposed directly into a respective angled body
portion 58 such that cooling flow issuing from a respective cooling
hole 34 flows through groove 50 to top portion 56 of pressure side
winglet 54 over top surface 52 of tip 26 and on to tip cap 28.
[0022] As shown in FIG. 3, the addition of a pressure side tip
winglet 54, or angled projection of tip surface, performs the
function of adding resistance to the flow of gases into the gap
between the blade tip and the stationary shroud. Such a winglet 54
is known to reduce hot gas leakage flows into the blade tip gap.
With the added requirement of film cooling for the blade tip, these
two functions can be combined in novel ways to synergistically
improve performance and extend blade life. Blade tip film holes 34
are here provided with substantially the same angle as winglet 54.
Winglet 54 in this embodiment is a straight surface with a sharp
corner at the coincidence of surfaces 56 and 58. The film holes are
thus issued tangentially onto the surface with a O-degree relative
angle, which drastically limits the ability of the hot gases to get
under the film layer or film jets. It is a well established effect,
that tangential film cooling on a surface is more efficient than
film cooling issued at an angle. This increase in cooling
efficiency can be very large, as much as doubling or even tripling
the film cooling effectiveness locally. The relative angle between
winglet 54 and film holes 34 need not be exactly 0 degrees, but can
vary from -15 to +15 degrees, typically, and still achieve the
desired effect. Furthermore, in this embodiment, film holes 34 are
discharged into grooves 50 in winglet 54, which grooves 50 are at
the same angle as winglet 54. Grooves 50 may be of various depths
and shapes. Grooves 50 serve to contain the film cooling and
further protect it from mixing with the hot gases. Grooves 50, or
channels, also serve to increase the external surface area covered
by the film cooling. Grooves 50 may be cast features in the blade
tip, or machined after casting, or even simply formed by laser
drilling as part of the process of forming the film holes
themselves. Grooves 50 need not be of constant cross section, but
could also flare out in size with distance from the film hole,
which can provide added benefit in performance. The groove depth
into the surface can vary; this is not restricted by the dimension
of the film hole. Two or more grooves 50 may proceed from a single
film hole to help spread the cooling while also protecting the
coolant from mixing with hot gases.
[0023] As shown in FIG. 4, winglet 54 edge defined by the
coincidence of surfaces 56 and 58 need not be sharp, but can be
rounded. This in fact will allow the cooling air to negotiate the
turn onto the tip cap region better. This figure also shows an
embodiment which may be used in connection with the present
invention, namely that the squealer tip perimeter rim need not
extend completely around the pressure side and suction side of the
tip; ie. need not form a tip cavity. In this embodiment, the blade
tip 26 has a single-tooth squealer located only along the suction
side. The winglet 54 and novel tip film cooling may still be
employed on the pressure side.
[0024] As shown in FIG. 5, film cooling holes 34 can be entirely
contained within winglet 54, rather than being discharged near the
base of winglet 54. By routing the film holes within the winglet
54, these cooling holes cease to be film cooling holes, but instead
become internal cooling for the winglet 54. Given a suitably thin
amount of material between the cooling hole and the external
surface of the winglet 54, this can result in very efficient
cooling of the winglet 54. This embodiment in essence provides a
total shield to the film holes, preventing any mixing with the hot
gases on the pressure side of the blade tip.
[0025] As shown in FIG. 6, this embodiment is a combination of
FIGS. 4 and 5, in which the film cooling holes 34 are not entirely
contained within the winglet 54.
[0026] As shown in FIG. 7, this embodiment is the same as that of
FIG. 4, but with another single-tooth seal location. This figure
shows an embodiment which may be used in connection with the
present invention, namely that the tip perimeter rim need not
extend completely around the pressure side and suction side of the
tip; ie. need not form a tip cavity. In this embodiment, the blade
tip has a single-tooth squealer located along or approximately
along the mean chordline of the blade tip section. The winglet 54
and novel tip film cooling may still be employed on the pressure
side.
[0027] These figures depict examples of the shaping which the film
hole grooves 50 may assume. In FIG. 8, grooves 50 are made to be
cylindrical in shape, and can be either the same diameter as the
film hole or larger in diameter. A larger diameter will provide
additional coolant spreading and surface area for cooling. In FIG.
9, grooves 50 are flared or fan-shaped diffusers from the film hole
exit to the tip surface 52. The degree of flare may be altered
continuously or abruptly. In FIG. 10, grooves 50 are formed with
two branches both emanating from the film hole exit. The branches
may be cylindrical or flared, and may be from 0 to 45 degrees in
included angle.
[0028] As cooling air 38 exits blade tip film cooling holes 34,
cooling air 38 flows into groove 50 and travels to a top surface 52
of tip 26 and flows into tip cap 28 to provide cooling thereto as
best shown in FIGS. 3 and 4. Grooves 50 provide a safe passage for
cooling flow 38 issuing from film cooling holes 34 resulting in
appropriate cooling of the tip cap 28 region, lessening end cap
degradation.
[0029] While typical embodiments have been set forth for the
purpose of illustration, the foregoing description should not be
deemed to be a limitation on the scope of the invention.
Accordingly, various modifications, adaptations, and alternatives
may occur to one skilled in the art without departing from the
spirit and scope of the present invention.
* * * * *