U.S. patent application number 11/494831 was filed with the patent office on 2009-08-20 for serpentine microcircuits for hot gas migration.
This patent application is currently assigned to United Technologies Corporation. Invention is credited to Francisco J. Cunha.
Application Number | 20090208343 11/494831 |
Document ID | / |
Family ID | 39121678 |
Filed Date | 2009-08-20 |
United States Patent
Application |
20090208343 |
Kind Code |
A1 |
Cunha; Francisco J. |
August 20, 2009 |
SERPENTINE MICROCIRCUITS FOR HOT GAS MIGRATION
Abstract
A turbine engine component, such as a turbine blade, has an
airfoil portion with a pressure side and a suction side. The
turbine engine component also has a first cooling circuit within
the pressure side for cooling the pressure side of the airfoil
portion and a second cooling circuit within the suction side for
cooling the suction side of the airfoil portion and for cooperating
with a wrap around leading edge cooling circuit for creating a
cooling film over the pressure side.
Inventors: |
Cunha; Francisco J.; (Avon,
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: |
39121678 |
Appl. No.: |
11/494831 |
Filed: |
July 28, 2006 |
Current U.S.
Class: |
416/97R |
Current CPC
Class: |
F01D 5/187 20130101;
F05D 2260/221 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 pressure side and a suction side; a first cooling circuit within
said pressure side for cooling said pressure side of said airfoil
portion; a second cooling circuit within said suction side for
cooling said suction side of said airfoil portion and for
cooperating with means for creating a cooling film over said
pressure side; a trailing edge internal circuit; and said first
cooling circuit having an exit which delivers cooling fluid to said
trailing edge internal circuit.
2. The turbine engine component according to claim 1, wherein said
means for creating a cooling film over said pressure side comprises
a cooling circuit wrapped around a leading edge of said airfoil
portion.
3. The turbine engine component according to claim 2, further
comprising said cooling circuit wrapped around said leading edge
having a first set of film holes for cooling said leading edge.
4. The turbine engine component according to claim 3 further
comprising said cooling circuit wrapped around said leading edge
having a second set of film holes for cooling said pressure side of
said airfoil portion.
5. The turbine engine component according to claim 1, further
comprising a leading edge internal circuit.
6. (canceled)
7. The turbine engine component according to claim 1, wherein said
first cooling circuit has a first leg, a second leg, and a bend
between said first leg and said second leg.
8. The turbine engine component according to claim 7, wherein said
second leg terminates in said exit.
9. The turbine engine component according to claim 7, further
comprising means for ensuring flow acceleration through the
bend.
10. The turbine engine component according to claim 9, wherein said
flow acceleration ensuring means comprises a plurality of
holes.
11. The turbine engine component according to claim 5, wherein each
of said internal circuits has a plurality of film holes for
creating a flow of cooling fluid over said pressure side and said
suction side.
12. The turbine engine component according to claim 5, wherein said
leading edge internal circuit has a plurality of cross-over holes
for supplying fluid to a leading edge cooling circuit.
13. The turbine engine component according to claim 5, wherein said
trailing edge internal circuit has a plurality of cross-over holes
for supplying fluid to a trailing edge cooling circuit.
14. A turbine engine component comprising: an airfoil portion
having a pressure side and a suction side; a first cooling circuit
within said pressure side for cooling said pressure side of said
airfoil portion; a second cooling circuit within said suction side
for cooling said suction side of said airfoil portion and for
cooperating with means for creating a cooling film over said
pressure side; a leading edge internal circuit and a trailing edge
internal circuit; and wherein each of said leading edge and said
trailing edge internal circuits has means for cooling a tip of said
airfoil portion.
15. A turbine engine component comprising: an airfoil portion
having a pressure side and a suction side; a first cooling circuit
within said pressure side for cooling said pressure side of said
airfoil portion; a second cooling circuit within said suction side
for cooling said suction side of said airfoil portion and for
cooperating with means for creating a cooling film over said
pressure side; and wherein said second cooling circuit has a first
leg, a second leg, and a bend between said first leg and said
second leg.
16. The turbine engine component according to claim 15, wherein
said second cooling circuit has a plurality of cross-over holes for
supplying cooling fluid to said means for creating a cooling film
over said pressure side.
17. The turbine engine component according to claim 15, wherein
said second leg communicates with said means for creating a cooling
film over said pressure side.
Description
BACKGROUND
[0001] (1) Field of the Invention
[0002] The present invention relates to a turbine engine component
having an improved scheme for cooling an airfoil portion.
[0003] (2) Prior Art
[0004] The overall cooling effectiveness is a measure used to
determine the cooling characteristics of a particular design. The
ideal non-achievable goal is unity, which implies that the metal
temperature is the same as the coolant temperature inside an
airfoil. The opposite can also occur when the cooling effectiveness
is zero implying that the metal temperature is the same as the gas
temperature. In that case, the blade material will certainly melt
and burn away. In general, existing cooling technology allows the
cooling effectiveness to be between 0.5 and 0.6. More advanced
technology such as supercooling should be between 0.6 and 0.7.
Microcircuit cooling as the most advanced cooling technology in
existence today can be made to produce cooling effectiveness higher
than 0.7.
[0005] FIG. 1 shows a durability map of cooling effectiveness
(x-axis) vs. the film effectiveness (y-axis) for different lines of
convective efficiency. Placed in the map is a point 10 related to a
new advanced serpentine microcircuit shown in FIGS. 2a-2c. This
serpentine microcircuit includes a pressure side serpentine circuit
20 and a suction side serpentine circuit 22 embedded in the airfoil
walls 24 and 26.
[0006] The Table I below provides the operational parameters used
to plot the design point in the durability map.
TABLE-US-00001 TABLE I Operational Parameters for serpentine
microcircuit beta 2.898 Tg 2581 [F.] Tc 1365 [F.] Tm 2050 [F.]
Tm_bulk 1709 [F.] Phi_loc 0.437 Phi_bulk 0.717 Tco 1640 [F.] Tci
1090 [F.] eta_c_loc 0.573 eta_f 0.296 Total Cooling Flow 3.503% WAE
10.8 Legend for Table I Beta = heat load Phi_loc = local cooling
effectiveness Phi_bulk = bulk cooling effectiveness Eta_c_loc =
local cooling efficiency Eta_f = film effectiveness Tg = gas
temperature Tc = coolant temperature Tm = metal temperature Tm_bulk
= bulk metal temperature Tco = exit coolant temperature Tci = inlet
coolant temperature WAE = compressor engine flow, pps
[0007] It should be noted that the overall cooling effectiveness
from the table is 0.717 for a film effectiveness of 0.296 and a
convective efficiency (or ability to pick-up heat) of 0.573. Also
note that the corresponding cooling flow for a turbine blade having
this cooling microcircuit is 3.5% engine flow. FIG. 3 illustrates
the cooling flow distribution for a turbine blade with the
serpentine microcircuits of FIGS. 2a-2c embedded in the airfoils
walls.
[0008] There are however field problems that can be addressed
efficiently with peripheral microcircuit designs. One such field
problem is illustrated in FIGS. 4A and 4B. In FIG. 4A, the
streamlines of the gas path close to the external surface of the
airfoil illustrate four different regions in which the gas flow
changes direction or migration: a tip region, two mid-section
regions, and a root region. In between the tip and the upper mid
region, the flow transitions through a pseudo stagnation point(s).
The momentum of the external gas seems to decelerate in such a way
as to impose a local thermal load to the part. This manifests
itself by regions where the propensity for erosion and oxidation
increase in the airfoil surface. The superposition of FIG. 4B
illustrates the local coincidence between the pseudo-stagnation
region and the blade distress in the part surface. In the mid
region, the upper and lower region also converge onto one another,
but even though the space between streamlines decreases, the flow
seems to accelerate and there is no pseudo-stagnation regions. A
mild manifestation of the same tip-to-mid phenomena seems to
initiate in the transition region between the mid-to-root regions.
It is therefore necessary to tailor the peripheral microcircuit in
such a manner as to address these local high thermal load
regions.
SUMMARY OF THE INVENTION
[0009] In accordance with the present invention, a turbine engine
component is provided with improved cooling. The turbine engine
component broadly comprises an airfoil portion having a pressure
side and a suction side. The turbine engine component further has a
first cooling circuit within the pressure side for cooling the
pressure side of the airfoil portion and a second cooling circuit
within the suction side for cooling the suction side of the airfoil
portion and for cooperating with means for creating a cooling film
over the pressure side.
[0010] Other details of the serpentine microcircuits for hot gas
migration 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
[0011] FIG. 1 is a graph showing cooling effectiveness versus film
effectiveness for a turbine engine component;
[0012] FIG. 2A shows an airfoil portion of a turbine engine
component having a pressure side cooling microcircuit embedded in
the pressure side wall and a suction side cooling microcircuit
embedded in the suction side wall;
[0013] FIG. 2B is a schematic representation of a pressure side
cooling microcircuit used in the airfoil portion of FIG. 2A;
[0014] FIG. 2C is a schematic representation of a suction side
cooling microcircuit used in the airfoil portion of FIG. 2A;
[0015] FIG. 3 illustrates the cooling flow distribution for a
turbine engine component with serpentine microcircuits embedded in
the airfoil walls;
[0016] FIG. 4A is a schematic representation illustrating the
pressure side distress on an airfoil surface;
[0017] FIG. 4B is a schematic representation of the local
coincidence between the pseudo-stagnation region and the blade
distress;
[0018] FIG. 5 is a schematic representation of a peripheral
pressure side cooling circuit;
[0019] FIG. 6 is a schematic representation of a peripheral suction
side cooling circuit; and
[0020] FIG. 7 is a schematic representation of main body internal
cooling circuits.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT(S)
[0021] Referring now to FIGS. 5 and 6, there are depicted two
peripheral cooling arrangements which may be used to address local
increases in the airfoil thermal load of a turbine engine component
90 such as a turbine blade. The two peripheral cooling arrangements
include a peripheral pressure side microcircuit 100 which may be
incorporated or embedded within the wall forming the pressure side
of an airfoil portion 104 and a suction side microcircuit 120 which
may be incorporated or embedded within the wall forming the suction
side of the airfoil portion 104.
[0022] In FIG. 5, the pressure side peripheral microcircuit 100 is
shown. In this circuit, the first leg 102 has an inlet 103 which
receives cooling fluid from a source (not shown). The leg 102
provides a flow of cooling fluid which quenches the hot spot in the
tip-to-mid region of the airfoil portion 104 shown in FIG. 4B. The
cooling fluid within the leg 102 proceeds around a 180 degree bend
106 which is supplemented with a plurality of film holes 108,
preferably three film holes. The film holes 108 ensure flow
acceleration through the bend 106 to a second downstream leg 110
which ends below the platform 112 of the turbine engine component
90 in an exit 164. Cooling fluid from the leg 110 is fed into an
internal trailing edge circuit 114 to be discussed hereinafter via
the exit 164 where it is used to further cool the airfoil portion
104.
[0023] Referring now to FIG. 6, there is shown a peripheral suction
side microcircuit 120. The circuit 120 has a first leg 122 which
communicates with a source (not shown) of cooling fluid. In the
first leg 122, the cooling flow convects heat away from the suction
side. Since the circuit 120 has no film holes, effective cooling
may not be done past the external gage point of the airfoil portion
104 where any film cooling would provide high aerodynamic penalties
due to mixing. (PLEASE CHECK THIS TO SEE IF IT MAKES SENSE) Thus,
the circuit 120 is used to feed cooling fluid to a leading edge
microcircuit 124 which wraps around the leading edge 126 of the
airfoil portion 104. The circuit 120 feeds or supplies cooling
fluid to the leading edge wrap around circuit 124 through a
plurality of wall cross over holes 128. As can be seen from FIG. 6,
the circuit 120 has a bend 130 and a second leg 132. The holes 128
are preferably located in the vicinity of the bend 130 and the
second leg 132. The second leg 132 may also communicate with the
wrap around circuit 124 via a passageway 134. As the microcircuit
124 wraps around the leading edge, several holes 136 are located in
the leading edge and are used to cool the leading edge of the
airfoil portion 104. Further, the microcircuit 124 is provided with
a plurality of film holes 138 for creating a film of cooling fluid
over the pressure side of the airfoil portion.
[0024] Referring now to FIG. 7, there is shown the main body
internal cooling circuits which include a leading edge internal
cooling circuit 150 and the trailing edge internal cooling circuit
114. The leading edge internal cooling circuit 150 communicates
with a source (not shown) of cooling fluid, such as engine bleed
air, via an inlet 151 and has one or more film cooling holes 152
adjacent the tip 154 of the airfoil portion 104 to provide tip
cooling. The circuit 150 also has a plurality of cross-over holes
156 for supplying cooling fluid to the leading edge microcircuit
124.
[0025] The trailing edge internal circuit 114 also communicates
with a source (not shown) of cooling fluid, such as engine bleed
air, via an inlet 157 and has one or more film cooling holes 158
adjacent the tip 154 to provide tip cooling. The circuit 114 also
has a plurality of cross-over holes 160 for communicating with a
trailing edge cooling circuit 162 for cooling the trailing edge of
the airfoil portion 104. As can be seen from FIG. 7, the tailing
edge internal circuit 114 also receives cooling fluid from the
peripheral pressure side microcircuit 100 via the exit 164.
[0026] Each of the leading edge internal circuit 150 and the
trailing edge internal circuit 114 may be provided with a plurality
of film cooling holes 170 and 172 respectively to form cooling
films over the pressure and suction sides of the airfoil portion
104.
[0027] Using the pressure and suction side cooling circuits of the
present invention, the airfoil portion of a turbine engine
component may be very effectively convectively cooled. Using the
pressure side circuit, the cooling flow is returned to the trailing
edge internal circuit for further cooling of the airfoil. Using the
suction side circuit, the leading edge of the airfoil is cooled
first before discharging in pressure side film. This effective use
of coolant allows for positive effects on cycle thermodynamic
efficiency, turbine efficiency, rotor inlet temperature impacts,
and specific fuel consumption.
[0028] It is apparent that there has been provided in accordance
with the present invention serpentine microcircuits for hot gas
migration which fully satisfy 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.
* * * * *