U.S. patent application number 11/484143 was filed with the patent office on 2008-01-10 for integral main body-tip microcircuits for blades.
This patent application is currently assigned to United Technologies Corporation. Invention is credited to William Abdel-Messeh, Francisco J. Cunha.
Application Number | 20080008599 11/484143 |
Document ID | / |
Family ID | 38461959 |
Filed Date | 2008-01-10 |
United States Patent
Application |
20080008599 |
Kind Code |
A1 |
Cunha; Francisco J. ; et
al. |
January 10, 2008 |
Integral main body-tip microcircuits for blades
Abstract
A turbine engine component, such as a turbine blade, has an
airfoil portion having a tip and a root portion and a cooling
microcircuit arrangement within the airfoil portion. The cooling
microcircuit arrangement comprises a multi-leg main body portion
for allowing a flow of coolant to convectively cool the airfoil
portion and at least one integrally formed tip cooling microcircuit
for cooling the tip.
Inventors: |
Cunha; Francisco J.; (Avon,
CT) ; Abdel-Messeh; William; (Middletown,
CT) |
Correspondence
Address: |
BACHMAN & LAPOINTE, P.C. (P&W)
900 CHAPEL STREET, SUITE 1201
NEW HAVEN
CT
06510-2802
US
|
Assignee: |
United Technologies
Corporation
|
Family ID: |
38461959 |
Appl. No.: |
11/484143 |
Filed: |
July 10, 2006 |
Current U.S.
Class: |
416/97R |
Current CPC
Class: |
F01D 5/186 20130101;
Y02T 50/60 20130101; Y02T 50/676 20130101; F05D 2250/185 20130101;
F05D 2230/21 20130101 |
Class at
Publication: |
416/97.R |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Claims
1. A turbine engine component comprising: an airfoil portion having
a tip and a root portion; a cooling microcircuit arrangement within
said airfoil portion; and said cooling microcircuit arrangement
comprising a multi-leg main body portion for allowing a flow of
coolant to convectively cool said airfoil portion and at least one
integrally formed tip cooling microcircuit.
2. The turbine engine component according to claim 1, wherein said
multi-leg main body portion comprises a serpentine cooling
arrangement.
3. The turbine engine component according to claim 1, further
comprising a coolant inlet extending through said root portion and
communicating with said multi-leg main body portion.
4. The turbine engine component according to claim 3, wherein said
coolant inlet is located adjacent a leading edge of said root
portion.
5. The turbine engine component according to claim 3, wherein said
multi-leg main body portion has a first leg which allows coolant
from said coolant inlet to flow radially upwards towards said
tip.
6. The turbine engine component according to claim 5, further
comprising said at least one integrally formed cooling microcircuit
being in fluid communication with said first leg.
7. The turbine engine component according to claim 5, further
comprising a plurality of integrally formed cooling microcircuits
being in fluid communication with said first leg.
8. The turbine engine component according to claim 5, wherein said
multi-leg main body portion has a second leg which receives coolant
from said first leg and in which said coolant flows radially
downward.
9. The turbine engine component according to claim 8, further
comprising a plurality of film cooling slots located between said
first leg and said second leg for allowing coolant to flow over a
surface of said airfoil portion.
10. The turbine engine component according to claim 8, further
comprising said multi-leg main body portion having a third leg in
which coolant moves radially upward.
11. The turbine engine component according to claim 10, further
comprising a refresher inlet extending within said root portion and
supplying coolant to said third leg.
12. The turbine engine component according to claim 10, further
comprising said at least one tip cooling microcircuit communicating
with said third leg.
13. The turbine engine component according to claim 10, further
comprising a plurality of tip cooling microcircuits communicating
with said third leg.
14. The turbine engine component according to claim 1, wherein each
said tip cooling microcircuit has an exit for allowing cooling
fluid to flow over a pressure side surface of said airfoil
portion.
15. The turbine engine component according to claim 1, further
comprising a platform positioned intermediate said airfoil portion
and said root portion.
16. The turbine engine component according to claim 1, further
comprising said airfoil portion having a leading edge, a pressure
side wall, and a suction side wall, and said leading edge having a
plurality of fluid outlets for allowing said coolant to flow over
at least one of a leading edge portion of the pressure side wall
and a leading edge portion of the suction side wall.
Description
BACKGROUND OF THE INVENTION
[0001] (1) Field of the Invention
[0002] The present invention relates to a cooling microcircuit for
use in a turbine engine component such as a turbine blade.
[0003] (2) Prior Art
[0004] As it can be appreciated from FIGS. 1 and 2, prior art
turbine blades have a plurality of cavities with each blade
internal cavity feeding a cooling microcircuit located on a side of
the airfoil, either on a pressure side or on a suction side. For
blade cooling designs with double wall construction, or imbedded
microcircuits in the airfoils, such as those illustrated in FIGS. 1
and 2, the flow in the internal supply cavities assume relatively
low Mach number distribution when compared to more conventional
serpentine cooling designs. Due to these low Mach numbers, the
rotational forces, resulting from the angular speeds of the rotor,
become predominant. This, in turn, induces a series of cavity
vortices inside the supply cavities. Since the internal cavities
supply coolant air to the microcircuits embedded in the wall, the
location of these supply links, between cavity and microcircuit,
become extremely important.
[0005] There are two possible ways to solve this problem. First,
one can determine the characteristics of these in-plane secondary
vortices in a way so as to facilitate access of the cooling air
into the microcircuits. Second, one can de-sensitize the
microcircuit cooling to the supply links by having one supply
cavity per microcircuit. If neither of these options is followed, a
concern exists of having multiple feeds from one internal cavity to
more than one microcircuit, which would lead to a flow
imbalance.
[0006] In the FIGS. 1 and 2, there is shown an airfoil with
embedded wall circuits having cooling supply flows from a large
internal cavity. In the configuration shown, the embedded circuits
on the pressure side would experience a relative lower flow
compared to those on the suction side. As a result the cooling
effectiveness is much lower on the pressure side than on the
suction side.
[0007] The simpler option is the second option above where a blade
is designed to have dedicated and independent cooling supplies to
each microcircuit. As a result, the microcircuit flow cooling
characteristics are then de-sensitized from potential high
rotational effects and other interferences. As a consequence of
this cooling design philosophy, seven internal large cavities were
cored for the turbine blade of FIG. 1, leading to six internal ribs
within the airfoil portion. The existence of several cold ribs is
particularly significant for this application, as cold ribs provide
increased creep capability for the airfoil. However, as illustrated
in FIGS. 3-5, the direct relationship of one microcircuit per
supply cavity leads to potential assembly issues when microcircuit
cores are tied-in with the main-body cores of the supply cavities
during the pre-casting operation. In addition, the cooling
microcircuit is a separate and independent circuit.
[0008] Thus, there remains a problem of core assembly to be solved
when designing cooling microcircuits for turbine engine components
such as turbine blades.
SUMMARY OF THE INVENTION
[0009] In accordance with the present invention, there is provided
a solution to the core assembly problem which also takes advantage
of the isolation of each microcircuit from a total independent
supply.
[0010] In accordance with the present invention, there is provided
a turbine engine component which broadly comprises an airfoil
portion having a tip, a root portion, and a cooling microcircuit
arrangement within the airfoil portion. The cooling microcircuit
arrangement comprises a multi-leg main body portion for allowing a
flow of coolant to convectively cool the airfoil portion and at
least one integrally formed tip cooling microcircuit for cooling
the tip of the airfoil portion.
[0011] Other details of the integral main body-tip microcircuits
for blades of the present invention, as well as other objects and
advantages attendant thereto, are set forth in the following
detailed description and the accompanying drawings wherein like
reference numerals depict like elements.
BRIEF DESCRIPTION OF THE DRAWINGS
[0012] FIG. 1 is an illustration of a prior art turbine blade
having internal blade cavities for separate imbedded airfoil wall
microcircuits;
[0013] FIG. 2 is an illustration of a prior art turbine blade
having dedicated supply cavities per microcircuit to avoid cooling
effectiveness debits;
[0014] FIG. 3 is an illustration of a portion of a prior art
turbine blade having three pressure side microcircuits;
[0015] FIG. 4 is an illustration of a portion of a prior art
turbine blade having two suction side microcircuits;
[0016] FIG. 5 is an illustration of a prior art turbine blade
having a trailing edge microcircuit; and
[0017] FIG. 6 is an illustration of a turbine blade having a
serpentine airfoil microcircuit supplied from the blade root and
integrated main mid-body with tip microcircuit as one unit.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT(S)
[0018] To solve the problem of core assembly, while taking
advantage of the isolation of each microcircuit from a total
independent supply, the cooling scheme of FIG. 6 is presented. As
shown in this figure, there is provided a turbine engine component
10, such as a turbine blade, having an airfoil portion 12, a
platform 14, and a root portion 16. The airfoil portion 12 has a
tip 18. A cooling microcircuit 20 is imbedded within the airfoil
portion 12. The imbedded cooling microcircuit 20 receives a coolant
flow stream from an inlet 24 formed within the root portion 16. The
inlet 24 is preferably positioned adjacent a leading edge of the
root portion 16. The inlet 24 may communicate with any suitable
source of cooling fluid such as engine bleed air. The coolant flow
stream is allowed to flow radially upward (in a direction away from
the platform 14) through a first leg 26 of the cooling microcircuit
20 so as to take advantage of the natural pumping force. As can be
seen from FIG. 6, the cooling microcircuit 20 may have a serpentine
configuration. Thus, as the coolant flow stream reaches the
vicinity of the tip 18 of the airfoil portion 12, the coolant flow
bends and proceeds to a second leg 28. Within the second leg 28,
the coolant flows radially downward (in a direction toward the
platform 14). In this arrangement, some bypass coolant flow may be
used to cool the tip 18 via tip cooling circuits 30 and 32. As
shown in FIG. 6, the tip cooling circuit 30 comprises a plurality
of spaced apart flow passages 70. Each flow passage 70 has an inlet
which may communicate with and receive coolant from the first leg
26 as well as from a U-shaped flow turn portion 34 connecting the
legs 26 and 28.
[0019] The cooling microcircuit 30 may be provided with a third leg
36 in which the coolant flows radially upward. The tip circuit 32
also may comprise a plurality of spaced apart flow passages 72.
Each flow passage 72 may have an inlet which communicates with the
third leg 36 of the cooling microcircuit 20 so as to receive
coolant therefrom. Each cooling circuit passage 70 and 72 has a
fluid outlet or exit 33 which allows cooling fluid to flow over a
surface of the airfoil portion 12. Preferably, the exits 33 are
configured to allow the coolant to exit on the pressure side 35 of
the airfoil portion 12. The tip cooling exits 33 from the circuits
30 and 32 may extend from a point near the leading edge 44 to a
point near the trailing edge 50 of the airfoil portion 12. By
providing the cooling microcircuit arrangement described herein,
three separate circuits make up one unit and thus facilitate the
assembly process.
[0020] A root inlet refresher leg 38 may be fabricated within the
root portion 16. The root inlet refresher leg 38 is in fluid
communication with the third leg 36 and may be used to insure
adequate cooling flow in the third leg 36. The root inlet refresher
leg 38 may communicate with any suitable source (not shown) of
cooling fluid such as engine bleed air.
[0021] As can be seen from the foregoing description, an integral
main body and tip microcircuit arrangement 20 has been provided.
The turbine engine component 10 is cooled convectively in this
way.
[0022] If desired, exit tabs 40 forming film slots 42 may be
provided in the legs 26 and/or 28. The exit tabs 40 and film slots
42 allow coolant fluid to flow from the legs 26 and/or 28 onto a
surface of the airfoil portion. The surface may be the pressure
side surface 35 or the suction side surface 37. Fluid exiting the
slots 42 helps form a cooling film over one or more of the exterior
surfaces of the turbine engine component 10. Such film slots 42 may
be useful in an open-cooling system.
[0023] If desired, the leading edge 44 of the airfoil portion 12
may be provided with a plurality of fluid outlets or exits 46 which
allow a film of coolant to flow over the leading edge portions of
the pressure side 35 and the suction side 37 of the airfoil portion
12. The outlets or exits 46 may be supplied with coolant from a
supply cavity 48. The supply cavity 48 may communicate directly
with a source (not shown) of cooling fluid, such as engine bleed
air, or alternatively, the supply cavity 48 may be in fluid
communication with the first leg 26.
[0024] The cooling microcircuit of the present invention may also
be used in a closed loop system without film cooling for industrial
gas turbine applications where the external thermal load is not as
high as that for aircraft engine applications.
[0025] The cooling microcircuit arrangement of the present
invention may be formed using any suitable technique known in the
art. In a preferred method of forming the cooling microcircuit, one
or more sheets formed from a refractory metal material may be
configured in the shape of the cooling microcircuit arrangement 20
including the inlet 24 and the root inlet refresher leg 38, the
legs 26, 28, and 36, the tip cooling microcircuits 30 and 32, the
exits 33, the tabs 40, and the film slots 42. The refractory metal
material sheets may be placed or positioned within a mold cavity.
Thereafter, the turbine engine component 10 including the airfoil
portion 12, the platform 14, and the root portion 16 may be cast
from any suitable metal known in the art such as a nickel based
superalloy, a titanium based superalloy, or an iron based
superalloy. After the turbine engine component has been cast, the
refractory metal material sheets may be removed using any suitable
means known in the art, leaving the cooling microcircuit
arrangement 20 of the present invention.
[0026] It is apparent that there has been provided in accordance
with the present invention an integral main body-tip microcircuit
for blades which fully satisfies the objects, means, and advantages
set forth hereinbefore. While the present invention has been
described in the context of specific embodiments thereof, other
unforeseeable alternatives, modifications, and variations may
become apparent to those skilled in the art having read the
foregoing description. Accordingly, it is intended to embrace those
alternatives, modifications, and variations as fall within the
broad scope of the appended claims.
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