U.S. patent application number 13/544112 was filed with the patent office on 2013-08-15 for multi-lobed cooling hole array.
This patent application is currently assigned to UNITED TECHNOLOGIES CORPORATION. The applicant listed for this patent is JinQuan Xu. Invention is credited to JinQuan Xu.
Application Number | 20130205803 13/544112 |
Document ID | / |
Family ID | 48944506 |
Filed Date | 2013-08-15 |
United States Patent
Application |
20130205803 |
Kind Code |
A1 |
Xu; JinQuan |
August 15, 2013 |
MULTI-LOBED COOLING HOLE ARRAY
Abstract
A gas turbine engine component includes a wall having first and
second wall surfaces and first and second cooling holes extending
through the wall. The first and second cooling holes each include
an inlet located at the first wall surface, an outlet located at
the second wall surface, a metering section extending downstream
from the inlet and a diffusing section extending from the metering
section to the outlet. Each diffusing section includes first and
second lobes, each lobe diverging longitudinally and laterally from
the metering section. The outlets of each cooling hole include
first and second lateral ends and a trailing edge. One of the
lateral ends of the outlet of the first cooling hole and one of the
lateral ends of the outlet of the second cooling hole meet upstream
of the trailing edge of the first cooling hole and the trailing
edge of the second cooling hole.
Inventors: |
Xu; JinQuan; (Groton,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Xu; JinQuan |
Groton |
CT |
US |
|
|
Assignee: |
UNITED TECHNOLOGIES
CORPORATION
Hartford
CT
|
Family ID: |
48944506 |
Appl. No.: |
13/544112 |
Filed: |
July 9, 2012 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
61599385 |
Feb 15, 2012 |
|
|
|
Current U.S.
Class: |
60/806 ;
29/889.22 |
Current CPC
Class: |
Y02T 50/60 20130101;
F05D 2260/202 20130101; F01D 5/186 20130101; Y02T 50/676 20130101;
Y10T 29/49323 20150115; F05D 2250/324 20130101; F05D 2260/221
20130101; F05D 2250/52 20130101 |
Class at
Publication: |
60/806 ;
29/889.22 |
International
Class: |
F02C 7/18 20060101
F02C007/18; B23P 11/00 20060101 B23P011/00 |
Claims
1. A gas turbine engine component comprising: a wall having first
and second wall surfaces; a first cooling hole extending through
the wall and comprising: an inlet located at the first wall
surface; an outlet located at the second wall surface and
comprising: first and second lateral ends; and a trailing edge; a
metering section extending downstream from the inlet; and a
diffusing section extending from the metering section to the
outlet, the diffusing section comprising: first and second lobes,
each lobe diverging longitudinally and laterally from the metering
section; and a second cooling hole extending through the wall and
comprising: an inlet located at the first wall surface; an outlet
located at the second wall surface and comprising: first and second
lateral ends; and a trailing edge; a metering section extending
downstream from the inlet; and a diffusing section extending from
the metering section to the outlet, the diffusing section
comprising: first and second lobes, each lobe diverging
longitudinally and laterally from the metering section, wherein one
of the lateral ends of the outlet of the first cooling hole and one
of the lateral ends of the outlet of the second cooling hole meet
upstream of the trailing edge of the first cooling hole and the
trailing edge of the second cooling hole.
2. The component of claim 1, wherein the second lateral end of the
first outlet and the first lateral end of the second outlet form a
cusp.
3. The component of claim 2, wherein the cusp formed by the first
lateral end of the first outlet and the second lateral end of the
second outlet comprises an upstream end and a downstream end.
4. The component of claim 3, wherein the cusp is located upstream
of the first outlet and the second outlet.
5. The component of claim 1, wherein a region near where the
lateral end of the first cooling hole and the lateral end of the
second cooling hole meet is smoothed to eliminate sharp
corners.
6. The component of claim 1, wherein at least one of the diffusing
sections of the first and second cooling holes further comprises an
interlobe region having a portion that extends between the first
and second lobes of the at least one diffusing section, the
interlobe region having an end adjacent the second wall
surface.
7. The component of claim 1, wherein the trailing edge of the first
cooling hole and the trailing edge of the second cooling hole are
parallel and radially aligned.
8. The component of claim 1, wherein the diffusing section of the
first cooling hole further comprises a third lobe positioned
between the first and second lobes, the third lobe diverging
longitudinally from the metering section of the first cooling
hole.
9. The component of claim 8, wherein the diffusing section of the
second cooling hole further comprises a third lobe positioned
between the first and second lobes, the third lobe diverging
longitudinally from the metering section of the second cooling
hole.
10. A wall of a component of a gas turbine engine, the wall
comprising: first and second wall surfaces; a first inlet located
at the first wall surface; a first outlet located at the second
wall surface; a first metering section commencing at the first
inlet and extending downstream from the first inlet; a first
diffusing section extending from the first metering section and
terminating at the first outlet, the first diffusing section
comprising: first and second lobes, each lobe diverging
longitudinally and laterally from the first metering section,
wherein the second lobe comprises a first lateral end surface; and
a first trailing edge; a second inlet located at the first wall
surface; a second outlet located at the second wall surface; a
second metering section commencing at the second inlet and
extending downstream from the second inlet; a second diffusing
section extending from the second metering section and terminating
at the second outlet, the second diffusing section comprising:
third and fourth lobes, each lobe diverging longitudinally and
laterally from the second metering section, wherein the third lobe
comprises a second lateral end surface; and a second trailing edge;
wherein the first lateral end surface of the second lobe and the
second lateral end surface of the third lobe meet upstream of the
first and second trailing edges.
11. The wall of claim 10, wherein the first lateral end surface of
the second lobe and the second lateral end surface of the third
lobe form a cusp.
12. The wall of claim 11, wherein the cusp formed by the first
lateral end surface of the second lobe and the second lateral end
surface of the third lobe comprises an upstream end and a
downstream end.
13. The wall of claim 13, wherein the cusp is located upstream of
the first and second outlets.
14. The wall of claim 10, wherein a region near where the first
lateral end surface and the second lateral end surface meet is
smoothed to eliminate sharp corners.
15. The wall of claim 10, wherein the first diffusing section
further comprises a first interlobe region having a portion that
extends between the first and second lobes, the first interlobe
region having an end adjacent the first outlet, and wherein the
second diffusing section further comprises a second interlobe
region having a portion that extends between the third and fourth
lobes, the second interlobe region having an end adjacent the
second outlet.
16. The wall of claim 10, wherein the first diffusing section
further comprises a fifth lobe positioned between the first and
second lobes, the fifth lobe diverging longitudinally from the
first metering section.
17. The wall of claim 16, wherein the second diffusing section
further comprises a sixth lobe positioned between the third and
fourth lobes, the sixth lobe diverging longitudinally from the
second metering section.
18. A method for producing an array of multi-lobed cooling holes
between first and second wall surfaces, the method comprising:
forming a first cooling hole, wherein the first cooling hole
comprises: an inlet located at the first wall surface; an outlet
located at the second wall surface; a metering section commencing
at the inlet and extending downstream from the inlet; a diffusing
section extending from the metering section and terminating at the
outlet, the diffusing section comprising: first and second lobes,
each lobe diverging longitudinally and laterally from the metering
section; and a trailing edge; and forming a second cooling hole,
wherein the second cooling hole comprises: an inlet located at the
first wall surface; an outlet located at the second wall surface; a
metering section commencing at the inlet and extending downstream
from the inlet; a diffusing section extending from the metering
section and terminating at the outlet, the diffusing section
comprising: first and second lobes, each lobe diverging
longitudinally and laterally from the metering section; and a
trailing edge, wherein the second lobe of the first cooling hole
and the first lobe of the second cooling hole meet upstream of the
outlets of the first and second cooling holes.
19. The method of claim 18, wherein at least a portion of the
diffusing sections of the first and second cooling holes are formed
by casting.
20. The method of claim 18, wherein at least a portion of the
diffusing sections of the first and second cooling holes are formed
by masking.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application claims priority to U.S. Provisional
Application No. 61/599,385, filed on Feb. 15, 2012 and entitled
"MULTI-LOBED COOLING HOLE ARRAY", the disclosure of which is
incorporated by reference in its entirety.
BACKGROUND
[0002] This invention relates generally to turbomachinery, and
specifically to turbine flow path components for gas turbine
engines. In particular, the invention relates to cooling techniques
for airfoils and other gas turbine engine components exposed to hot
working fluid flow, including, but not limited to, rotor blades and
stator vane airfoils, endwall surfaces including platforms, shrouds
and compressor and turbine casings, combustor liners, turbine
exhaust assemblies, thrust augmentors and exhaust nozzles.
[0003] Gas turbine engines are rotary-type combustion turbine
engines built around a power core made up of a compressor,
combustor and turbine, arranged in flow series with an upstream
inlet and downstream exhaust. The compressor section compresses air
from the inlet, which is mixed with fuel in the combustor and
ignited to generate hot combustion gas. The turbine section
extracts energy from the expanding combustion gas, and drives the
compressor section via a common shaft. Expanded combustion products
are exhausted downstream, and energy is delivered in the form of
rotational energy in the shaft, reactive thrust from the exhaust,
or both.
[0004] Gas turbine engines provide efficient, reliable power for a
wide range of applications in aviation, transportation and
industrial power generation. Small-scale gas turbine engines
typically utilize a one-spool design, with co-rotating compressor
and turbine sections. Larger-scale combustion turbines including
jet engines and industrial gas turbines (IGTs) are generally
arranged into a number of coaxially nested spools. The spools
operate at different pressures, temperatures and spool speeds, and
may rotate in different directions.
[0005] Individual compressor and turbine sections in each spool may
also be subdivided into a number of stages, formed of alternating
rows of rotor blade and stator vane airfoils. The airfoils are
shaped to turn, accelerate and compress the working fluid flow, or
to generate lift for conversion to rotational energy in the
turbine.
[0006] Industrial gas turbines often utilize complex nested spool
configurations, and deliver power via an output shaft coupled to an
electrical generator or other load, typically using an external
gearbox. In combined cycle gas turbines (CCGTs), a steam turbine or
other secondary system is used to extract additional energy from
the exhaust, improving thermodynamic efficiency. Gas turbine
engines are also used in marine and land-based applications,
including naval vessels, trains and armored vehicles, and in
smaller-scale applications such as auxiliary power units.
[0007] Aviation applications include turbojet, turbofan, turboprop
and turboshaft engine designs. In turbojet engines, thrust is
generated primarily from the exhaust. Modern fixed-wing aircraft
generally employ turbofan and turboprop configurations, in which
the low pressure spool is coupled to a propulsion fan or propeller.
Turboshaft engines are employed on rotary-wing aircraft, including
helicopters, typically using a reduction gearbox to control blade
speed. Unducted (open rotor) turbofans and ducted propeller engines
also known, in a variety of single-rotor and contra-rotating
designs with both forward and aft mounting configurations.
[0008] Aviation turbines generally utilize two and three-spool
configurations, with a corresponding number of coaxially rotating
turbine and compressor sections. In two-spool designs, the high
pressure turbine drives a high pressure compressor, forming the
high pressure spool or high spool. The low-pressure turbine drives
the low spool and fan section, or a shaft for a rotor or propeller.
In three-spool engines, there is also an intermediate pressure
spool. Aviation turbines are also used to power auxiliary devices
including electrical generators, hydraulic pumps and elements of
the environmental control system, for example using bleed air from
the compressor or via an accessory gearbox.
[0009] Additional turbine engine applications and turbine engine
types include intercooled, regenerated or recuperated and variable
cycle gas turbine engines, and combinations thereof. In particular,
these applications include intercooled turbine engines, for example
with a relatively higher pressure ratio, regenerated or recuperated
gas turbine engines, for example with a relatively lower pressure
ratio or for smaller-scale applications, and variable cycle gas
turbine engines, for example for operation under a range of flight
conditions including subsonic, transonic and supersonic speeds.
Combined intercooled and regenerated/recuperated engines are also
known, in a variety of spool configurations with traditional and
variable cycle modes of operation.
[0010] Turbofan engines are commonly divided into high and low
bypass configurations. High bypass turbofans generate thrust
primarily from the fan, which accelerates airflow through a bypass
duct oriented around the engine core. This design is common on
commercial aircraft and transports, where noise and fuel efficiency
are primary concerns. The fan rotor may also operate as a first
stage compressor, or as a pre-compressor stage for the low-pressure
compressor or booster module. Variable-area nozzle surfaces can
also be deployed to regulate the bypass pressure and improve fan
performance, for example during takeoff and landing. Advanced
turbofan engines may also utilize a geared fan drive mechanism to
provide greater speed control, reducing noise and increasing engine
efficiency, or to increase or decrease specific thrust.
[0011] Low bypass turbofans produce proportionally more thrust from
the exhaust flow, generating greater specific thrust for use in
high-performance applications including supersonic jet aircraft.
Low bypass turbofan engines may also include variable-area exhaust
nozzles and afterburner or augmentor assemblies for flow regulation
and short-term thrust enhancement. Specialized high-speed
applications include continuously afterburning engines and hybrid
turbojet/ramjet configurations.
[0012] Across these applications, turbine performance depends on
the balance between higher pressure ratios and core gas path
temperatures, which tend to increase efficiency, and the related
effects on service life and reliability due to increased stress and
wear. This balance is particularly relevant to gas turbine engine
components in the hot sections of the compressor, combustor,
turbine and exhaust sections, where active cooling is required to
prevent damage due to high gas path temperatures and pressures.
SUMMARY
[0013] A gas turbine engine component includes a wall having first
and second wall surfaces and first and second cooling holes
extending through the wall. The first and second cooling holes each
include an inlet located at the first wall surface, an outlet
located at the second wall surface, a metering section extending
downstream from the inlet and a diffusing section extending from
the metering section to the outlet. Each diffusing section includes
first and second lobes, each lobe diverging longitudinally and
laterally from the metering section. The outlets of each cooling
hole include first and second lateral ends and a trailing edge. One
of the lateral ends of the outlet of the first cooling hole and one
of the lateral ends of the outlet of the second cooling hole meet
upstream of the trailing edge of the first cooling hole and the
trailing edge of the second cooling hole.
[0014] A gas path wall of a component of a gas turbine engine
includes first and second wall surfaces, first and second inlets
located at the first wall surface, and first and second outlets
located at the second wall surface. A first metering section
commences at the first inlet and extends downstream from the first
inlet. A first diffusing section extends from the first metering
section and terminates at the first outlet. The first diffusing
section includes a first trailing edge and first and second lobes,
each lobe diverging longitudinally and laterally from the first
metering section. The second lobe includes a first lateral end
surface. A second metering section commences at the second inlet
and extends downstream from the second inlet. A second diffusing
section extends from the second metering section and terminates at
the second outlet. The second diffusing section includes a second
trailing edge and third and fourth lobes, each lobe diverging
longitudinally and laterally from the second metering section. The
third lobe includes a second lateral end surface. The first lateral
end surface of the second lobe and the second lateral end surface
of the third lobe meet upstream of the first and second trailing
edges.
[0015] A method for producing an array of multi-lobed cooling holes
between first and second wall surfaces includes forming a first
cooling hole and a second cooling hole. The first cooling hole
includes an inlet located at the first wall surface, an outlet
located at the second wall surface, a metering section commencing
at the inlet and extending downstream from the inlet and a
diffusing section extending from the metering section and
terminating at the outlet. The diffusing section includes first and
second lobes, each lobe diverging longitudinally and laterally from
the metering section, and a trailing edge. The second cooling hole
includes an inlet located at the first wall surface, an outlet
located at the second wall surface, a metering section commencing
at the inlet and extending downstream from the inlet and a
diffusing section extending from the metering section and
terminating at the outlet. The diffusing section includes first and
second lobes, each lobe diverging longitudinally and laterally from
the metering section, and a trailing edge. The second lobe of the
first cooling hole and the first lobe of the second cooling hole
meet upstream of the outlets of the first and second cooling
holes.
BRIEF DESCRIPTION OF THE DRAWINGS
[0016] FIG. 1 is a cross-sectional view of a gas turbine
engine.
[0017] FIG. 2A is a perspective view of an airfoil for the gas
turbine engine, in a rotor blade configuration.
[0018] FIG. 2B is a perspective view of an airfoil for the gas
turbine engine, in a stator vane configuration.
[0019] FIG. 3 is a view of a wall having an array of multi-lobed
cooling holes.
[0020] FIG. 4 is a sectional view of one of the multi-lobed cooling
holes of FIG. 3 taken along the line 4-4.
[0021] FIG. 5 is a view of the multi-lobed cooling hole of FIG. 4
taken along the line 5-5.
[0022] FIG. 6 is a view of two adjacent multi-lobed cooling holes
of FIG. 3.
[0023] FIG. 7 is a sectional view of the multi-lobed cooling holes
of FIG. 6 taken along the line 7-7.
DETAILED DESCRIPTION
[0024] FIG. 1 is a cross-sectional view of gas turbine engine 10.
Gas turbine engine (or turbine engine) 10 includes a power core
with compressor section 12, combustor 14 and turbine section 16
arranged in flow series between upstream inlet 18 and downstream
exhaust 20. Compressor section 12 and turbine section 16 are
arranged into a number of alternating stages of rotor airfoils (or
blades) 22 and stator airfoils (or vanes) 24.
[0025] In the turbofan configuration of FIG. 1, propulsion fan 26
is positioned in bypass duct 28, which is coaxially oriented about
the engine core along centerline (or turbine axis) C.sub.L. An
open-rotor propulsion stage 26 may also provided, with turbine
engine 10 operating as a turboprop or unducted turbofan engine.
Alternatively, fan rotor 26 and bypass duct 28 may be absent, with
turbine engine 10 configured as a turbojet or turboshaft engine, or
an industrial gas turbine.
[0026] For improved service life and reliability, components of gas
turbine engine 10 are provided with an improved cooling
configuration, as described below. Suitable components for the
cooling configuration include rotor airfoils 22, stator airfoils 24
and other gas turbine engine components exposed to hot gas flow,
including, but not limited to, platforms, shrouds, casings and
other endwall surfaces in hot sections of compressor 12 and turbine
16, and liners, nozzles, afterburners, augmentors and other gas
wall components in combustor 14 and exhaust section 20.
[0027] In the two-spool, high bypass configuration of FIG. 1,
compressor section 12 includes low pressure compressor (LPC) 30 and
high pressure compressor (HPC) 32, and turbine section 16 includes
high pressure turbine (HPT) 34 and low pressure turbine (LPT) 36.
Low pressure compressor 30 is rotationally coupled to low pressure
turbine 36 via low pressure (LP) shaft 38, forming the LP spool or
low spool. High pressure compressor 32 is rotationally coupled to
high pressure turbine 34 via high pressure (HP) shaft 40, forming
the HP spool or high spool.
[0028] Flow F at inlet 18 divides into primary (core) flow F.sub.P
and secondary (bypass) flow F.sub.S downstream of fan rotor 26. Fan
rotor 26 accelerates secondary flow F.sub.S through bypass duct 28,
with fan exit guide vanes (FEGVs) 42 to reduce swirl and improve
thrust performance. In some designs, structural guide vanes (SGVs)
42 are used, providing combined flow turning and load bearing
capabilities.
[0029] Primary flow F.sub.P is compressed in low pressure
compressor 30 and high pressure compressor 32, then mixed with fuel
in combustor 14 and ignited to generate hot combustion gas. The
combustion gas expands to provide rotational energy in high
pressure turbine 34 and low pressure turbine 36, driving high
pressure compressor 32 and low pressure compressor 30,
respectively. Expanded combustion gases exit through exhaust
section (or exhaust nozzle) 20, which can be shaped or actuated to
regulate the exhaust flow and improve thrust performance.
[0030] Low pressure shaft 38 and high pressure shaft 40 are mounted
coaxially about centerline C.sub.L, and rotate at different speeds.
Fan rotor (or other propulsion stage) 26 is rotationally coupled to
low pressure shaft 38. In advanced designs, fan drive gear system
44 is provided for additional fan speed control, improving thrust
performance and efficiency with reduced noise output.
[0031] Fan rotor 26 may also function as a first-stage compressor
for gas turbine engine 10, and LPC 30 may be configured as an
intermediate compressor or booster. Alternatively, propulsion stage
26 has an open rotor design, or is absent, as described above. Gas
turbine engine 10 thus encompasses a wide range of different shaft,
spool and turbine engine configurations, including one, two and
three-spool turboprop and (high or low bypass) turbofan engines,
turboshaft engines, turbojet engines, and multi-spool industrial
gas turbines.
[0032] In each of these applications, turbine efficiency and
performance depend on the overall pressure ratio, defined by the
total pressure at inlet 18 as compared to the exit pressure of
compressor section 12, for example at the outlet of high pressure
compressor 32, entering combustor 14. Higher pressure ratios,
however, also result in greater gas path temperatures, increasing
the cooling loads on rotor airfoils 22, stator airfoils 24 and
other components of gas turbine engine 10. To reduce operating
temperatures, increase service life and maintain engine efficiency,
these components are provided with improved cooling configurations,
as described below. Suitable components include, but are not
limited to, cooled gas turbine engine components in compressor
sections 30 and 32, combustor 14, turbine sections 34 and 36, and
exhaust section 20 of gas turbine engine 10.
[0033] FIG. 2A is a perspective view of rotor airfoil (or blade) 22
for gas turbine engine 10, as shown in FIG. 1, or for another
turbomachine. Rotor airfoil 22 extends axially from leading edge 51
to trailing edge 52, defining pressure surface 53 (front) and
suction surface 54 (back) therebetween.
[0034] Pressure and suction surfaces 53 and 54 form the major
opposing surfaces or walls of airfoil 22, extending axially between
leading edge 51 and trailing edge 52, and radially from root
section 55, adjacent inner diameter (ID) platform 56, to tip
section 57, opposite ID platform 56. In some designs, tip section
57 is shrouded.
[0035] Cooling holes or outlets 60 are provided on one or more
surfaces of airfoil 22, for example along leading edge 51, trailing
edge 52, pressure (or concave) surface 53, or suction (or convex)
surface 54, or a combination thereof. Cooling holes or passages 60
may also be provided on the endwall surfaces of airfoil 22, for
example along ID platform 56, or on a shroud or engine casing
adjacent tip section 57.
[0036] FIG. 2B is a perspective view of stator airfoil (or vane) 24
for gas turbine engine 10, as shown in FIG. 1, or for another
turbomachine. Stator airfoil 24 extends axially from leading edge
61 to trailing edge 62, defining pressure surface 63 (front) and
suction surface 64 (back) therebetween. Pressure and suction
surfaces 63 and 64 extend from inner (or root) section 65, adjacent
ID platform 66, to outer (or tip) section 67, adjacent outer
diameter (OD) platform 68.
[0037] Cooling holes or outlets 60 are provided along one or more
surfaces of airfoil 24, for example leading or trailing edge 61 or
62, pressure (concave) or suction (convex) surface 63 or 64, or a
combination thereof. Cooling holes or passages 60 may also be
provided on the endwall surfaces of airfoil 24, for example along
ID platform 66 and OD platform 68.
[0038] Rotor airfoils 22 (FIG. 2A) and stator airfoils 24 (FIG. 2B)
are formed of high strength, heat resistant materials such as high
temperature alloys and superalloys, and are provided with thermal
and erosion-resistant coatings. Airfoils 22 and 24 are also
provided with internal cooling passages and cooling holes 60 to
reduce thermal fatigue and wear, and to prevent melting when
exposed to hot gas flow in the higher temperature regions of a gas
turbine engine or other turbomachine. Cooling holes 60 deliver
cooling fluid (e.g., steam or air from a compressor) through the
outer walls and platform structures of airfoils 22 and 24, creating
a thin layer (or film) of cooling fluid to protect the outer (gas
path) surfaces from high temperature flow.
[0039] While surface cooling extends service life and increases
reliability, injecting cooling fluid into the gas path also reduces
engine efficiency, and the cost in efficiency increases with the
required cooling flow. Cooling holes 60 are thus provided with
improved metering and inlet geometry to reduce jets and blow off,
and improved diffusion and exit geometry to reduce flow separation
and corner effects. Cooling holes 60 reduce flow requirements and
improve the spread of cooling fluid across the hot outer surfaces
of airfoils 22 and 24, and other gas turbine engine components, so
that less flow is needed for cooling and efficiency is maintained
or increased.
[0040] The array of multi-lobed cooling holes described herein
provide a cooling solution that offers improved lateral film
cooling coverage and eliminates or reduces the problems associated
with conventional diffusion film cooling holes, such as flow
separation, blow off and low resistance to thermo-mechanical
fatigue. Multi-lobed cooling holes provide improved film
effectiveness and reduce the likelihood of film separation so that
they work as intended at high blowing ratios and reduce the
detrimental effects of kidney vortices. The array of cooling holes
described herein also provide increased resistance to
thermo-mechanical fatigue by eliminating "sharp" corners within the
cooling holes.
[0041] FIG. 3 illustrates a view of a wall of a gas turbine engine
component having an array of multi-lobed film cooling holes. Wall
100 includes first wall surface 102 and second wall surface 104. As
described in greater detail below, wall 100 is primarily metallic
and second wall surface 104 can include a thermal barrier coating.
Multi-lobed film cooling holes 106 are oriented so that their
inlets are positioned on the first wall surface 102 and their
outlets are positioned on second wall surface 104. During gas
turbine engine operation, second wall surface 104 is in proximity
to high temperature gases (e.g., combustion gases, hot air).
Cooling air is delivered inside wall 100 where it exits the
interior of the component through cooling holes 106 and forms a
cooling film on second wall surface 104. As shown in FIG. 3,
cooling holes 106 have two lobes in the diffusing section of the
cooling hole outlet on second wall surface 104. Multiple
multi-lobed film cooling holes 106 are positioned side-by-side in a
row to form cooling hole array 107.
[0042] As described below in greater detail, cooling air flows out
of cooling holes 106, with cooling air flowing through each of the
lobes in the diffusing section. Cooling holes 106 of cooling hole
array 107 can be arranged in a row on wall 100 as shown in FIG. 3
and positioned axially so that the cooling air flows in
substantially the same direction longitudinally as the high
temperature gases flowing past wall 100. In this embodiment,
cooling air passing through cooling holes 106 exits cooling holes
traveling in substantially the same direction as the high
temperature gases flowing along second wall surface 104
(represented by arrow H). Here, the linear row of cooling holes 106
is substantially perpendicular to the direction of flow H to create
array 107. In alternate embodiments, the orientation of cooling
holes 106 can be arranged on second wall surface 104 so that the
flow of cooling air is substantially perpendicular to the hot air
flow (i.e. cooling air exits cooling holes 106 radially) or at an
angle between parallel and perpendicular. Array 107 can also
include staggered cooling holes 106 on wall 100. Cooling holes 106
can be located on a variety of components that require cooling.
Suitable components include, but are not limited to, turbine vanes
and blades, combustors, blade outer air seals, and augmentors, etc.
Cooling holes 106 can be located on the pressure side or suction
side of vanes and blades. Cooling holes 106 can also be located on
the blade tip or blade or vane platforms.
[0043] FIGS. 4 and 5 illustrate one embodiment of a single cooling
hole 106 in greater detail. A single cooling hole is described
below to illustrate the features present within cooling holes 106
of array 107. FIG. 4 illustrates a sectional view of multi-lobed
film cooling hole 106 of FIG. 3 taken along the line 4-4. FIG. 5
illustrates a view of cooling hole 106 of FIG. 4 taken along the
line 5-5. For the purposes of illustration, wall 100 has been
removed from FIG. 5 to better show cooling hole 106. Cooling hole
106 includes inlet 110, metering section 112, diffusing section 114
and outlet 116. Inlet 110 is an opening located on first wall
surface 102. Cooling air C enters cooling hole 106 through inlet
110 and passes through metering section 112 and diffusing section
114 before exiting cooling hole 106 at outlet 116 along second wall
surface 104.
[0044] Metering section 112 is adjacent to and downstream from
inlet 110 and controls (meters) the flow of cooling air through
cooling hole 106. In exemplary embodiments, metering section 112
has a substantially constant flow area from inlet 110 to diffusing
section 114. Metering section 112 can have circular, oblong (oval
or elliptical), racetrack (oval with two parallel sides having
straight portions) crescent, cusp or dual-cusp shaped axial cross
sections. In FIGS. 4 and 5, metering section 112 has a circular
cross section. Circular metering sections 112 have a length l and
diameter d. In exemplary embodiments, inlet 110 and metering
section 112 have the same diameter d. In some embodiments, circular
metering section 112 has a length l according to the relationship:
d.ltoreq.l.ltoreq.3d. That is, the length of metering section 112
is between one and three times its diameter. The length of metering
section 112 can exceed 3d, reaching upwards of 30d. In alternate
embodiments, metering section 112 has an oblong or racetrack-shaped
or other shaped cross section. As oblong and racetrack
configurations are not circular, their metering sections 112 have a
length l and hydraulic diameter d.sub.h. In some embodiments,
metering section 112 has a length l according to the relationship:
d.sub.h.ltoreq.l.ltoreq.3d.sub.h. That is, the length of metering
section 112 is between one and three times its hydraulic diameter.
Again, the length of metering section 112 can exceed 3d.sub.h,
reaching upwards of 30d.sub.h. In exemplary embodiments, metering
section 112 is inclined with respect to wall 100 as illustrated in
FIG. 4 (i.e. metering section 112 is not perpendicular to wall
100). Metering section 112 has a longitudinal axis represented by
numeral 118.
[0045] Diffusing section 114 is adjacent to and downstream from
metering section 112. Cooling air C diffuses within diffusing
section 114 before exiting cooling hole 106 along second wall
surface 104. Second wall surface 104 includes upstream end 120
(upstream of cooling hole 106) and downstream end 122 (downstream
from cooling hole 106). Diffusing section 114 opens along second
wall surface 104 between upstream end 120 and downstream end 122.
As shown in FIG. 4, cooling air C diffuses away from longitudinal
axis 118 in diffusing section 114 as it flows towards outlet
116.
[0046] As shown best in FIG. 5, diffusing section 114 includes two
channel-like lobes 124 and 126 as described in the U.S. Provisional
Application No. 61/599,372, filed on Feb. 15, 2012 and entitled
"MULTI-LOBED COOLING HOLE AND METHOD OF MANUFACTURE", which is
incorporated by reference. Lobes 124 and 126 are surfaces of wall
100 which define the void of cooling hole 106 at diffusing section
114. Each lobe 124, 126 diverges longitudinally and laterally from
metering section 112 and has a bottom surface (bottom surfaces 128
and 130, respectively), a side wall along the outer edge of
diffusing section 114 (the side walls are represented by lines 132
and 134, respectively) and a trailing edge (trailing edges 136 and
138, respectively). FIG. 4 best illustrates the longitudinal
divergence (from longitudinal axis 118), while FIG. 5 best
illustrates the lateral divergence (from centerline axis 140). As
shown in FIG. 5, first lobe 124 laterally diverges upwards from
centerline axis 140 and second lobe 126 laterally diverges
downwards from centerline axis 140. Cooling air C leaving metering
section 112 and entering diffusing section 114 diffuses into lobes
124 and 126, causing the cooling air to spread laterally within
diffusing section 114. Side wall 132 and bottom surface 128 direct
cooling air C through first lobe 124, and side wall 134 and bottom
surface 130 direct cooling air C through second lobe 126.
[0047] Diffusing section 114 also includes interlobe region 142.
Portion 144 of interlobe region 142 is located between first lobe
124 and second lobe 126. End 146 of interlobe region 142 is
adjacent outlet 116 where the outlet meets second wall surface 104.
Portion 144, located between first lobe 124 and second lobe 126,
can extend towards metering section 112 to varying degrees. The
location of end 146 of interlobe region 142 relative to trailing
edges 136 and 138 can also vary. In the embodiment shown in FIG. 5,
end 146 meets trailing edges 136 and 138 of lobes 124 and 126,
respectively at outlet 116. In this embodiment, trailing edges 136
and 138 and hence, first lobe 124 and second lobe 126, extend to
outlet 116 at second wall surface 104. In other embodiments, end
146 of interlobe region 142 is spaced from trailing edges 136 and
138. In these embodiments, trailing edges 136 and 138 and hence,
first lobe 124 and second lobe 126, do not extend to outlet 116 at
second wall surface 104.
[0048] In the embodiment illustrated in FIG. 5, diffusing section
114 also includes first inclined portion 148 and second inclined
portion 150. First inclined portion 148 is located adjacent to and
extends from bottom surface 128 of first lobe 124. First inclined
portion 148 extends from first lobe 124 towards centerline axis 140
and second lobe 126. Second inclined portion 150 is located
adjacent to and extends from bottom surface 130 of second lobe 126.
Second inclined portion 150 extends from second lobe 126 towards
centerline axis 140 and first lobe 124. Depending on the location
of cooling hole 106, first inclined portion 148 and second inclined
portion 150 can have varying lateral and longitudinal lengths and
extend from lobes 124 and 126 at various angles (inclinations).
Like the side walls and bottom surfaces, first and second inclined
portions 148 and 150 direct cooling air C through lobes 124 and 126
of diffusing section 114.
[0049] In some embodiments, first inclined portion 148 and second
inclined portion 150 meet together to form a ridge as shown in FIG.
5. Ridge 152 is located between first lobe 124 and second lobe 126
at the intersection of first inclined portion 148 and second
inclined portion 150. Ridge 152 aids in separating and directing
the flow of cooling air C into first lobe 124 and second lobe 126.
The location and angle of ridge 152 within diffusing section 114
can vary to direct cooling air C within diffusing section 114 to
suit the location and desired flow profile of cooling hole 106.
[0050] Ridge 152 can extend longitudinally to varying degrees
between metering section 112 and interlobe region 142. Ridge 152
can extend upstream all the way to metering section 112, beginning
where metering section 112 and diffusing section 114 meet as shown
in FIG. 4. Alternatively, ridge 152 can begin farther downstream
(closer to outlet 116). Ridge 152 can extend downstream to
interlobe region 142 as shown in FIG. 4. Alternatively, ridge 152
can converge with bottom surfaces 128 and 130 upstream of interlobe
region 142. Corresponding changes to the longitudinal lengths of
first inclined portion 148 and second inclined portion 150 must
accompany any change in the longitudinal extension of ridge 152. As
shown in FIG. 4, ridge 152 does not extend to outlet 116.
[0051] Interlobe region 142 (and portions 144 and 145) can take
various shapes and have different configurations depending on the
location and desired flow profile of cooling hole 106. The bottom
surface of interlobe region 142 can be flat or curved. A curved
(longitudinally convex) bottom surface of interlobe region 142 can
facilitate improved flow attachment due to the Coanda effect.
Interlobe region 142 can have a compound trapezoidal shape as shown
in FIG. 5. In some embodiments, ridge 154 separates interlobe
region 142 into two sides having surfaces in two different planes.
Ridge 154 converges with bottom surface 130 of second lobe 126 at
outlet 116 at second wall surface 104 as shown in FIG. 4. The
intersection of ridges 152 and 154 at the point where interlobe
region 142 meets first inclined portion 148 and second inclined
portion 150 forms apex 156. By forming apex 156 upstream of outlet
116, diffusing section 114 facilitates improved flow
attachment.
[0052] In other embodiments, cooling hole 106 has diffusing section
114 with three channel-like lobes 124, 126 and 128 as described in
the U.S. Provisional Application No. 61/599,381, filed on Feb. 15,
2012 and entitled "TRI-LOBED COOLING HOLE AND METHOD OF
MANUFACTURE", which is incorporated by reference.
[0053] FIGS. 6 and 7 illustrate two adjacent multi-lobed cooling
holes 106. FIG. 6 shows a plan view of two multi-lobed cooling
holes 106A and 106B. FIG. 7 is a sectional view of multi-lobed
cooling holes 106A and 106B of FIG. 6 taken along the line 7-7. For
the purposes of illustration, wall 100 has been removed from FIGS.
6 and 7 to better show cooling holes 106A and 106B. In some
embodiments, cooling holes 106A and 106B have trailing edges that
are substantially straight, or parallel or aligned with each other
as shown in FIGS. 6 and 7. End 146A of cooling hole 106A and end
146B of cooling hole 106B are each located at second wall surface
104 at the same axial position (i.e. axially aligned). Ends 146A
and 146B are also parallel. In alternative embodiments, ends 146A
and 146B can be offset (i.e. staggered) or non-parallel (i.e.
cooling air C from cooling holes 106A and 106B can laterally
converge or diverge). In other embodiments, ends 146A and 146B can
be curved instead of straight. Staggered and non-parallel
arrangements of adjacent cooling holes 106 can allow individual
cooling holes 106 to be oriented and aligned with high temperature
gases passing over the cooling hole.
[0054] Cooling hole 106A abuts cooling hole 106B at outlets 116A
and 116B. Cooling hole 106A and cooling hole 106B meet along ridge
157 upstream of outlets 116A and 116B (i.e. ridge 157 is located
within the diffusing sections and not at the outlets). In some
embodiments, ridge 157 is rounded or smoothed so that it does not
create a feature having a sharp (acute) angle. Side wall 134A
(first lateral end surface) of second lobe 126A of cooling hole
106A meets with side wall 132B (second lateral end surface) of
first lobe 124B (third lobe) of cooling hole 106B at outlets 116A
and 116B upstream of the trailing edge (ends 146A and 146B) as
shown in FIG. 6. By locating adjacent cooling holes 106A and 106B
so that their lateral end surfaces meet at outlets 116A and 116B
upstream of the trailing edge, a continuous and uninterrupted film
of cooling air C is formed along second wall surface 104 spanning
both cooling hole 106A and cooling hole 106B.
[0055] In some embodiments, cooling hole 106A and cooling hole 106B
meet to form cusp 158 (shown best in FIG. 7). Lobe 126A forms one
side of cusp 158, while lobe 124B forms the other side of cusp 158.
Cusp 158 includes upstream end 160 and downstream end 162. In some
embodiments, cusp 158 does not extend all the way downstream to
outlet 116A or outlet 116B as shown in FIG. 6 by the dashed line.
Locating the ends of cusp 158 and/or common wall 157 upstream of
outlets 116A and 116B creates a continuous trailing edge (146A and
146B) along both cooling holes as shown in FIG. 6. Creating a
continuous trailing edge reduces the potential for
thermo-mechanical fatigue effects on cooling holes 106A and 106B.
In some embodiments, cusp 158 is rounded or smoothed so that it
does not create a feature having a sharp angle. A continuous
trailing edge can also be easier to manufacture by casting and/or
masking methods. In some embodiments, at least a portion of
diffusing sections 114A and 114B are formed by casting and/or
masking.
[0056] To further reduce the likelihood of thermo-mechanical
fatigue, regions near upstream end 160 of cusp 158 can be smoothed
to eliminate sharp corners and edges at outlets 116A and 116B. As
shown in FIG. 7, region 164 near upstream end 160 is rounded to
prevent sharp corners or points along cusp 158 and the areas
between outlets 116A and 116B.
[0057] FIGS. 6 and 7 illustrate two cooling holes (106A and 106B)
with each cooling hole having two lobes 128 and 130. In other
embodiments, adjacent cooling holes 106 can both have three lobes.
In still other embodiments, adjacent cooling holes 106 can include
one cooling hole having two lobes and the other cooling hole having
three lobes. Cooling holes having multiple diffusing sections can
also be placed in array 107. Cooling holes with multiple diffusing
sections are described in U.S. Provisional Application No.
61/599,384, filed on Feb. 15, 2012 and entitled "MULTIPLE DIFFUSING
COOLING HOLE", which is incorporated by reference.
[0058] The gas turbine engine components, gas path walls and
cooling passages described herein can thus be manufactured using
one or more of a variety of different processes. These techniques
provide each cooling hole and cooling passage with its own
particular configuration and features, including, but not limited
to, inlet, metering, transition, diffusion, outlet, upstream wall,
downstream wall, lateral wall, longitudinal, lobe and downstream
edge features, as described above. In some cases, multiple
techniques can be combined to improve overall cooling performance
or reproducibility, or to reduce manufacturing costs.
[0059] Suitable manufacturing techniques for forming the cooling
configurations described here include, but are not limited to,
electrical discharge machining (EDM), laser drilling, laser
machining, electrical chemical machining (ECM), water jet
machining, casting, conventional machining and combinations
thereof. Electrical discharge machining includes both machining
using a shaped electrode as well as multiple pass methods using a
hollow spindle or similar electrode component. Laser machining
methods include, but are not limited to, material removal by
ablation, trepanning and percussion laser machining. Conventional
machining methods include, but are not limited to, milling,
drilling and grinding.
[0060] The gas flow path walls and outer surfaces of some gas
turbine engine components include one or more coatings, such as
bond coats, thermal barrier coatings, abrasive coatings, abradable
coatings and erosion or erosion-resistant coatings. For components
having a coating, the inlet, metering portion, transition,
diffusion portion and outlet cooling features may be formed prior
to coating application, after a first coating (e.g., a bond coat)
is applied, or after a second or third (e.g., interlayer) coating
process, or a final coating (e.g., environmental or thermal
barrier) coating process. Depending on component type, cooling hole
or passage location, repair requirements and other considerations,
the diffusion portion and outlet features may be located within a
wall or substrate, within a thermal barrier coating or other
coating layer applied to a wall or substrate, or based on
combinations thereof. The cooling geometry and other features may
remain as described above, regardless of position relative to the
wall and coating materials or airfoil materials.
[0061] In addition, the order in which cooling features are formed
and coatings are applied may affect selection of manufacturing
techniques, including techniques used in forming the inlet,
metering portion, transition, outlet, diffusion portion and other
cooling features. For example, when a thermal barrier coat or other
coating is applied to the outer surface of a gas path wall before
the cooling hole or passage is produced, laser ablation or laser
drilling may be used. Alternatively, either laser drilling or water
jet machining may be used on a surface without a thermal barrier
coat. Additionally, different machining methods may be more or less
suitable for forming different features of the cooling hole or
cooling passage, for example, different EDM, laser machining and
other machining techniques may be used for forming the outlet and
diffusion features, and for forming the transition, metering and
inlet features.
[0062] While the invention has been described with reference to
exemplary embodiments, it will be understood by those skilled in
the art that various changes may be made and equivalents may be
substituted for elements thereof without departing from the scope
of the invention. In addition, many modifications may be made to
adapt a particular situation or material to the teachings of the
invention without departing from the essential scope thereof.
Therefore, it is intended that the invention not be limited to the
particular embodiments disclosed, but that the invention will
include all embodiments falling within the scope of the appended
claims.
Discussion of Possible Embodiments
[0063] The following are non-exclusive descriptions of possible
embodiments of the present invention.
[0064] A gas turbine engine component can include a wall having
first and second wall surfaces and first and second cooling holes
extending through the wall. The first and second cooling holes can
each include an inlet located at the first wall surface, an outlet
located at the second wall surface, a metering section extending
downstream from the inlet and a diffusing section extending from
the metering section to the outlet. Each diffusing section can
include first and second lobes, each lobe diverging longitudinally
and laterally from the metering section. The outlets of each
cooling hole can include first and second lateral ends and a
trailing edge. One of the lateral ends of the outlet of the first
cooling hole and one of the lateral ends of the outlet of the
second cooling hole can meet upstream of the trailing edge of the
first cooling hole and the trailing edge of the second cooling
hole.
[0065] The system of the preceding paragraph can optionally
include, additionally and/or alternatively any, one or more of the
following features, configurations and/or additional
components:
[0066] the second lateral end of the first outlet and the first
lateral end of the second outlet can form a cusp;
[0067] the cusp formed by the first lateral end of the first outlet
and the second lateral end of the second outlet can include an
upstream end and a downstream end;
[0068] the cusp can be located upstream of the first outlet and the
second outlet;
[0069] a region near where the lateral end of the first cooling
hole and the lateral end of the second cooling hole meet can be
smoothed to eliminate sharp corners;
[0070] at least one of the diffusing sections of the first and
second cooling holes can further include a first interlobe region
having a portion that extends between the first and second lobes of
the at least one diffusing section, the first interlobe region
having an end adjacent the second wall surface;
[0071] the diffusing section of the first cooling hole can further
include a third lobe positioned between the first and second lobes,
the third lobe diverging longitudinally from the metering section
of the first cooling hole; and/or
[0072] the diffusing section of the second cooling hole can further
include a third lobe positioned between the first and second lobes,
the third lobe diverging longitudinally from the metering section
of the second cooling hole.
[0073] A wall of a component of a gas turbine engine can include
first and second wall surfaces, first and second inlets located at
the first wall surface, and first and second outlets located at the
second wall surface. A first metering section can commence at the
first inlet and extend downstream from the first inlet. A first
diffusing section can extend from the first metering section and
terminate at the first outlet. The first diffusing section can
include a first trailing edge and first and second lobes, each lobe
diverging longitudinally and laterally from the first metering
section. The second lobe can include a first lateral end surface. A
second metering section can commence at the second inlet and extend
downstream from the second inlet. A second diffusing section can
extend from the second metering section and terminate at the second
outlet. The second diffusing section can include a second trailing
edge and third and fourth lobes, each lobe diverging longitudinally
and laterally from the second metering section. The third lobe can
include a second lateral end surface. The first lateral end surface
of the second lobe and the second lateral end surface of the third
lobe can meet upstream of the first and second trailing edges.
[0074] The system of the preceding paragraph can optionally
include, additionally and/or alternatively any, one or more of the
following features, configurations and/or additional
components:
[0075] the first lateral end surface of the second lobe and the
second lateral end surface of the third lobe can form a cusp;
[0076] the cusp formed by the first lateral end surface of the
second lobe and the second lateral end surface of the third lobe
can include an upstream end and a downstream end;
[0077] the cusp can be located upstream of the first and second
outlets;
[0078] a region near where the first lateral end surface and the
second lateral end surface meet can be smoothed to eliminate sharp
corners;
[0079] the first diffusing section can further include a first
interlobe region having a portion that extends between the first
and second lobes, the first interlobe region having an end adjacent
the first outlet; and the second diffusing section can further
include a second interlobe region having a portion that extends
between the third and fourth lobes, the second interlobe region
having an end adjacent the second outlet;
[0080] the first trailing edge and the second trailing edge can be
parallel and radially aligned;
[0081] the first diffusing section can further include a fifth lobe
positioned between the first and second lobes, the fifth lobe
diverging longitudinally from the first metering section;
and/or
[0082] the second diffusing section can further include a sixth
lobe positioned between the third and fourth lobes, the sixth lobe
diverging longitudinally from the second metering section.
[0083] A method for producing an array of multi-lobed cooling holes
between first and second wall surfaces can include forming a first
cooling hole and a second cooling hole. The first cooling hole can
include an inlet located at the first wall surface, an outlet
located at the second wall surface, a metering section commencing
at the inlet and extending downstream from the inlet and a
diffusing section extending from the metering section and
terminating at the outlet. The diffusing section can include first
and second lobes, each lobe diverging longitudinally and laterally
from the metering section, and a trailing edge. The second cooling
hole can include an inlet located at the first wall surface, an
outlet located at the second wall surface, a metering section
commencing at the inlet and extending downstream from the inlet and
a diffusing section extending from the metering section and
terminating at the outlet. The diffusing section can include first
and second lobes, each lobe diverging longitudinally and laterally
from the metering section, and a trailing edge. The second lobe of
the first cooling hole and the first lobe of the second cooling
hole can meet upstream of the outlets of the first and second
cooling holes.
[0084] The system of the preceding paragraph can optionally
include, additionally and/or alternatively any, one or more of the
following features, configurations and/or additional
components:
[0085] at least a portion of the diffusing sections of the first
and second cooling holes can be formed by casting; and/or
[0086] at least a portion of the diffusing sections of the first
and second cooling holes can be formed by masking.
* * * * *