U.S. patent application number 15/138624 was filed with the patent office on 2017-10-26 for airfoil for a turbine engine.
The applicant listed for this patent is General Electric Company. Invention is credited to Robert David Briggs, Scott Ronald Bunker, Douglas Gerard Konitzer.
Application Number | 20170306764 15/138624 |
Document ID | / |
Family ID | 58632269 |
Filed Date | 2017-10-26 |
United States Patent
Application |
20170306764 |
Kind Code |
A1 |
Konitzer; Douglas Gerard ;
et al. |
October 26, 2017 |
AIRFOIL FOR A TURBINE ENGINE
Abstract
A method and apparatus for minimizing engine weight for a
turbine engine by utilizing one or more discrete protuberances
disposed on an engine component wall. The wall can have a nominal
thickness to decrease engine weight while the protuberances can
provide increased discrete thicknesses for providing one or more
cooling holes. The increased thickness at the protuberances
provides for an increased thickness to provide sufficient length to
increase cooling hole effectiveness.
Inventors: |
Konitzer; Douglas Gerard;
(West Chester, OH) ; Bunker; Scott Ronald; (West
Chester, OH) ; Briggs; Robert David; (West Chester,
OH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Family ID: |
58632269 |
Appl. No.: |
15/138624 |
Filed: |
April 26, 2016 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F01D 5/147 20130101;
F05D 2260/22141 20130101; F01D 5/187 20130101; F01D 5/186 20130101;
F01D 25/002 20130101; Y02T 50/60 20130101; F05D 2230/20 20130101;
F05D 2240/30 20130101; F05D 2260/202 20130101; F05D 2250/185
20130101; F05D 2220/323 20130101; F05D 2250/232 20130101 |
International
Class: |
F01D 5/18 20060101
F01D005/18; F01D 25/00 20060101 F01D025/00 |
Claims
1. A component for a turbine engine, with the turbine engine
generating a hot combustion gas flow and providing a cooling fluid
flow, the component comprising: a wall having a nominal thickness
separating the hot combustion gas flow from the cooling fluid flow
having a hot surface facing the hot combustion gas flow and a cool
surface facing the cooling fluid flow; at least one localized,
protuberance extending from the cool surface; and a film hole
extending through the protuberance and the wall, having a length
greater than the nominal thickness of the wall.
2. The component of claim 1 wherein the film hole is angled
relative to a local normal between the hot surface and the cool
surface.
3. The component of claim 1 wherein the film hole is
non-linear.
4. The component of claim 1 wherein the film hole includes an inlet
and an outlet, having a passage connecting the inlet to the
outlet.
5. The component of claim 4 wherein at least one of the inlet,
outlet, or passage is shaped.
6. The component of claim 4 wherein the protuberance includes an
upstream side and a downstream side having the inlet disposed on
the upstream side.
7. The component of claim 1 wherein the protuberance includes a
height and the height is at least 50% of the nominal thickness.
8. The component of claim 7 wherein the height is at least 100% of
the nominal thickness.
9. The component of claim 1 wherein the protuberance is symmetrical
about an axis parallel to a direction of the cooling fluid flow
within the component.
10. The component of claim 1 wherein the protuberance includes a
height and the height is a function of one of a length, diameter,
or length-to-diameter ratio of the film hole.
11. The component of claim 1 wherein the protuberance includes an
upstream side and a downstream side having the film hole disposed
on the upstream side.
12. The component of claim 1 wherein the at least one protuberance
includes a recess.
13. The component of claim 12 wherein the film hole is disposed in
the recess.
14. The component of claim 1 wherein the nominal thickness is a
function of at least one of a thermal load, vibratory force,
pressure differential between the fluid flows, or manufacturer
required load for the component.
15. The component of claim 1 wherein the protuberances are formed
by additive manufacturing.
16. The component of claim 1 wherein the protuberance is one of
rounded, conical, frustoconical, or non-rectilinear.
17. The component of claim 1 wherein there is one film hole
disposed in the at least one protuberance.
18. An airfoil for a turbine engine comprising: a wall having a
first side adjacent to a first fluid flow and a second side
adjacent to a second fluid flow and having a nominal thickness; at
least one localized, protuberance extending from the wall; and a
cooling hole extending through the protuberance and the wall,
having a length greater than the nominal thickness of the wall.
19. The airfoil of claim 18 wherein the first side is an interior
surface of the airfoil and the protuberance is disposed on the
first side.
20. The airfoil of claim 18 wherein the wall is a rib of the
airfoil.
21. The airfoil of claim 18 wherein the cooling hole includes an
inlet and an outlet, having a passage connecting the inlet to the
outlet.
22. The airfoil of claim 21 wherein at least one of the inlet,
outlet, or passage is shaped.
23. The airfoil of claim 18 wherein the protuberance includes an
upstream side and a downstream side having the inlet disposed on
the upstream side.
24. The airfoil of claim 18 wherein the protuberance includes a
height and the height is at least 50% of the nominal thickness.
25. The airfoil of claim 24 wherein the height is at least 100% of
the nominal thickness.
26. The airfoil of claim 18 wherein the protuberance includes a
height and the height is a function of one of a length, diameter,
or length-to-diameter ratio of the cooling hole.
27. The airfoil of claim 18 wherein the at least one protuberance
includes a recess.
28. The airfoil of claim 27 wherein the cooling hole is disposed in
the recess.
29. The airfoil of claim 18 wherein the nominal thickness is a
function of at least one of a thermal load, vibratory force,
pressure differential between the first and second fluids, or
manufacturer required load for the airfoil.
30. The airfoil of claim 18 wherein the protuberances are formed by
additive manufacturing.
31. The airfoil of claim 18 wherein the protuberance is one of
rounded, conical, frustoconical, or non-rectilinear.
32. The airfoil of claim 18 wherein there is one film hole disposed
in the at least one protuberance.
33. A method of cooling an engine component having a hot surface
and a cool surface comprising: passing a cooling fluid flow along
the cool surface; and providing at least a portion of the cooling
fluid flow through a film hole in a localized protuberance
extending from the cool surface.
34. The method of claim 33 wherein providing at least a portion of
the cooling fluid flow includes providing the portion of the
cooling fluid flow through a recess in the protuberance prior to
providing the cooling fluid flow to the film hole.
35. The method of claim 34 wherein providing the portion of the
cooling fluid flow through the recess minimizes dust accumulation.
Description
BACKGROUND OF THE INVENTION
[0001] Turbine engines, and particularly gas or combustion turbine
engines, are rotary engines that extract energy from a flow of
combusted gases passing through the engine onto a multitude of
rotating turbine blades.
[0002] Turbine engines for aircraft are often designed to operate
at high temperatures to maximize engine efficiency, so cooling of
certain engine components can be beneficial or necessary.
Typically, cooling is accomplished by ducting cooler air from the
high and/or low pressure compressors to the engine components that
require cooling. Temperatures in the high pressure turbine can be
1000.degree. C. to 2000.degree. C. and the cooling air from the
compressor can be around 500.degree. C. to 700.degree. C. While the
compressor air is a high temperature, it is cooler relative to the
turbine air, and can be used to cool the turbine.
[0003] Contemporary engine components, such as the rotating blades,
necessarily account for a portion of the overall engine weight.
Decreasing the weight of these engine components is desirable to
increase engine efficiency. Decreasing weight of the engine
components can be accomplished by utilizing thinner walls for the
components, for example. However, thinner walls include a decreased
volume through which film holes can extend, which can decrease the
effectiveness of the film holes. Thus, it is desirable to utilize
thinner walls for the engine components to decrease system weight
while providing sufficient length for the film holes to maintain
cooling effectiveness.
BRIEF DESCRIPTION OF THE INVENTION
[0004] In one aspect, embodiments of the invention relate to a
component for a turbine engine, which generates a hot combustion
gas flow, and provides a cooling fluid flow defining a cooling
fluid flow, includes a wall separating the hot combustion gas flow
from the cooling fluid flow. The wall includes a hot surface facing
the hot combustion gas flow and a cool surface facing the cooling
fluid flow, having a nominal thickness. The component further
includes at least one localized, radiused protuberance extending
from the cool surface and a film hole extending through the
protuberance and the wall. The film hole has a greater length than
the nominal thickness of the wall.
[0005] In another aspect, embodiments of the invention relate to an
airfoil for a turbine engine including a wall having a first side
adjacent to a first fluid flow and a second side adjacent to a
second fluid flow, having a nominal thickness. The component
further includes at least one localized, radiused protuberance
extending from the wall and a film hole extending through the
protuberance and the wall, having a length greater than the nominal
thickness of the wall.
[0006] In yet another aspect, embodiments of the invention relate
to a method of cooling an engine component having a cool surface.
The method includes passing a cooling fluid flow along the cool
surface and providing at least a portion of the cooling fluid flow
through a film hole in a protuberance extending from the cool
surface.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] In the drawings:
[0008] FIG. 1 is a schematic cross-sectional diagram of a gas
turbine engine for an aircraft.
[0009] FIG. 2 is an isometric view of an airfoil of the gas turbine
engine of FIG. 1.
[0010] FIG. 3 is a cross-sectional view of the airfoil of FIG. 2
having walls including protuberances with film holes.
[0011] FIG. 4 is a perspective view of the wall having the
protuberance with the film hole extending through the protuberance
and the wall.
[0012] FIG. 5 is a cross-sectional view of the wall of FIG. 4
illustrating a profile of the protuberance and film hole across a
direction of a cooling fluid flow.
[0013] FIG. 6 is a cross-sectional view of the wall of FIG. 4
illustrating a profile of the protuberance and film hole in the
direction of the cooling fluid flow.
[0014] FIG. 7 is a cross-sectional view of another embodiment of
the film hole extending in a direction opposite of the cooling
fluid flow.
[0015] FIG. 8 is a cross-sectional view of yet another embodiment
the film hole having an inlet on a forward face of the
protuberance.
[0016] FIG. 9 is a cross-sectional view of the protuberance having
a recess with the film hole inlet disposed in the recess.
[0017] FIG. 10 is a perspective view of an elongated protuberance
having a film hole offset form the direction of the cooling fluid
flow.
DESCRIPTION OF EMBODIMENTS OF THE INVENTION
[0018] The described embodiments of the present invention are
directed to an engine component for a turbine engine having at
least one protuberance disposed on a wall of the engine component
with a film hole extending through the protuberance and the wall.
For purposes of illustration, the present invention will be
described with respect to the turbine for an aircraft gas turbine
engine. It will be understood, however, that the invention is not
so limited and may have general applicability within an engine, as
well as in non-aircraft applications, such as other mobile
applications and non-mobile industrial, commercial, and residential
applications.
[0019] As used herein, the term "forward" or "upstream" refers to
moving in a direction toward the engine inlet, or a component being
relatively closer to the engine inlet as compared to another
component. The term "aft" or "downstream" used in conjunction with
"forward" or "upstream" refers to a direction toward the rear or
outlet of the engine or being relatively closer to the engine
outlet as compared to another component.
[0020] Additionally, as used herein, the terms "radial" or
"radially" refer to a dimension extending between a center
longitudinal axis of the engine and an outer engine
circumference.
[0021] All directional references (e.g., radial, axial, proximal,
distal, upper, lower, upward, downward, left, right, lateral,
front, back, top, bottom, above, below, vertical, horizontal,
clockwise, counterclockwise, upstream, downstream, forward, aft,
etc.) are only used for identification purposes to aid the reader's
understanding of the present invention, and do not create
limitations, particularly as to the position, orientation, or use
of the invention. Connection references (e.g., attached, coupled,
connected, and joined) are to be construed broadly and can include
intermediate members between a collection of elements and relative
movement between elements unless otherwise indicated. As such,
connection references do not necessarily infer that two elements
are directly connected and in fixed relation to one another. The
exemplary drawings are for purposes of illustration only and the
dimensions, positions, order and relative sizes reflected in the
drawings attached hereto can vary.
[0022] FIG. 1 is a schematic cross-sectional diagram of a gas
turbine engine 10 for an aircraft. The engine 10 has a generally
longitudinally extending axis or centerline 12 extending forward 14
to aft 16. The engine 10 includes, in downstream serial flow
relationship, a fan section 18 including a fan 20, a compressor
section 22 including a booster or low pressure (LP) compressor 24
and a high pressure (HP) compressor 26, a combustion section 28
including a combustor 30, a turbine section 32 including a HP
turbine 34, and a LP turbine 36, and an exhaust section 38.
[0023] The fan section 18 includes a fan casing 40 surrounding the
fan 20. The fan 20 includes a plurality of fan blades 42 disposed
radially about the centerline 12. The HP compressor 26, the
combustor 30, and the HP turbine 34 form a core 44 of the engine
10, which generates combustion gases. The core 44 is surrounded by
core casing 46, which can be coupled with the fan casing 40.
[0024] A HP shaft or spool 48 disposed coaxially about the
centerline 12 of the engine 10 drivingly connects the HP turbine 34
to the HP compressor 26. A LP shaft or spool 50, which is disposed
coaxially about the centerline 12 of the engine 10 within the
larger diameter annular HP spool 48, drivingly connects the LP
turbine 36 to the LP compressor 24 and fan 20. The spools 48, 50
are rotatable about the engine centerline and couple to a plurality
of rotatable elements, which can collectively define a rotor
51.
[0025] The LP compressor 24 and the HP compressor 26 respectively
include a plurality of compressor stages 52, 54, in which a set of
compressor blades 56, 58 rotate relative to a corresponding set of
static compressor vanes 60, 62 (also called a nozzle) to compress
or pressurize the stream of fluid passing through the stage. In a
single compressor stage 52, 54, multiple compressor blades 56, 58
can be provided in a ring and can extend radially outwardly
relative to the centerline 12, from a blade platform to a blade
tip, while the corresponding static compressor vanes 60, 62 are
positioned upstream of and adjacent to the rotating blades 56, 58.
It is noted that the number of blades, vanes, and compressor stages
shown in FIG. 1 were selected for illustrative purposes only, and
that other numbers are possible.
[0026] The blades 56, 58 for a stage of the compressor can be
mounted to a disk 61, which is mounted to the corresponding one of
the HP and LP spools 48, 50, with each stage having its own disk
61. The vanes 60, 62 for a stage of the compressor can be mounted
to the core casing 46 in a circumferential arrangement.
[0027] The HP turbine 34 and the LP turbine 36 respectively include
a plurality of turbine stages 64, 66, in which a set of turbine
blades 68, 70 are rotated relative to a corresponding set of static
turbine vanes 72, 74 (also called a nozzle) to extract energy from
the stream of fluid passing through the stage. In a single turbine
stage 64, 66, multiple turbine blades 68, 70 can be provided in a
ring and can extend radially outwardly relative to the centerline
12, from a blade platform to a blade tip, while the corresponding
static turbine vanes 72, 74 are positioned upstream of and adjacent
to the rotating blades 68, 70. It is noted that the number of
blades, vanes, and turbine stages shown in FIG. 1 were selected for
illustrative purposes only, and that other numbers are
possible.
[0028] The blades 68, 70 for a stage of the turbine can be mounted
to a disk 71, which is mounted to the corresponding one of the HP
and LP spools 48, 50, with each stage having a dedicated disk 71.
The vanes 72, 74 for a stage of the compressor can be mounted to
the core casing 46 in a circumferential arrangement.
[0029] Complementary to the rotor portion, the stationary portions
of the engine 10, such as the static vanes 60, 62, 72, 74 among the
compressor and turbine section 22, 32 are also referred to
individually or collectively as a stator 63. As such, the stator 63
can refer to the combination of non-rotating elements throughout
the engine 10.
[0030] In operation, the airflow exiting the fan section 18 is
split such that a portion of the airflow is channeled into the LP
compressor 24, which then supplies pressurized airflow 76 to the HP
compressor 26, which further pressurizes the air. The pressurized
airflow 76 from the HP compressor 26 is mixed with fuel in the
combustor 30 and ignited, thereby generating combustion gases. Some
work is extracted from these gases by the HP turbine 34, which
drives the HP compressor 26. The combustion gases are discharged
into the LP turbine 36, which extracts additional work to drive the
LP compressor 24, and the exhaust gas is ultimately discharged from
the engine 10 via the exhaust section 38. The driving of the LP
turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP
compressor 24.
[0031] A portion of the pressurized airflow 76 can be drawn from
the compressor section 22 as bleed air 77. The bleed air 77 can be
draw from the pressurized airflow 76 and provided to engine
components requiring cooling. The temperature of pressurized
airflow 76 entering the combustor 30 is significantly increased. As
such, cooling provided by the bleed air 77 is necessary for
operating of such engine components in the heightened temperature
environments.
[0032] A remaining portion of the airflow 78 bypasses the LP
compressor 24 and engine core 44 and exits the engine assembly 10
through a stationary vane row, and more particularly an outlet
guide vane assembly 80, comprising a plurality of airfoil guide
vanes 82, at the fan exhaust side 84. More specifically, a
circumferential row of radially extending airfoil guide vanes 82
are utilized adjacent the fan section 18 to exert some directional
control of the airflow 78.
[0033] Some of the air supplied by the fan 20 can bypass the engine
core 44 and be used for cooling of portions, especially hot
portions, of the engine 10, and/or used to cool or power other
aspects of the aircraft. In the context of a turbine engine, the
hot portions of the engine are normally downstream of the combustor
30, especially the turbine section 32, with the HP turbine 34 being
the hottest portion as it is directly downstream of the combustion
section 28. Other sources of cooling fluid can be, but are not
limited to, fluid discharged from the LP compressor 24 or the HP
compressor 26.
[0034] FIG. 2 is a perspective view of an engine component in the
form the turbine blades 68 of the engine 10 of FIG. 1. It should be
understood that the turbine blade 68 is exemplary and that the
engine component can include other components requiring cooling.
The turbine blade 68 includes a dovetail 90 and an airfoil 92. The
airfoil 92 includes a tip 94 to a root 96. A span-wise direction
can be defined between the tip 94 and the root 96. The dovetail 90
includes a platform 98 and one or more inlet passages 100 having an
outlet 102. The dovetail 90 and platform 98 can be integral with
the airfoil 92 adjoining at the root 96. The platform 98 helps to
radially contain the turbine airflow driven by the airfoil 92. The
dovetail 90 can be configured to mount to a turbine rotor disk 71
(FIG. 1) to rotate the airfoil 92 about the engine centerline 10.
The inlet passages 100 can be fed with a flow of air, such as
bypass air 104. The bypass air 104 is provided to the airfoil 92 at
the root 96 exhausting through the outlets 102. It should be
appreciated that the dovetail 90 is shown in cross-section, such
that the inlet passages 100 are housed within the body of the
dovetail 90.
[0035] Referring to FIG. 3, the airfoil 92, shown in cross-section,
has an interior 110 bounded by an outer wall 112. The outer wall
112 includes a concave-shaped pressure sidewall 114 and a
convex-shaped suction sidewall 116. A leading edge 118 and a
trailing edge 120 are defined at the junction between the pressure
and suction sidewalls 114, 116, defining a chord-wise distance
between the leading and trailing edges 118, 120. The airfoil 92,
when implemented as a rotating blade as compared to a stationary
vane, rotates in a direction such that the pressure sidewall 114
follows the suction sidewall 116. Thus, as shown in FIG. 3, the
airfoil 92 would rotate upward toward the top of the page.
[0036] One or more ribs 130 are included in the interior 110. The
ribs 130 can extend between the pressure and suction sidewalls 114,
116 to define internal chambers 132. The chambers 132 can be
discrete compartments defined within the airfoil 92. Alternatively,
the chambers 132 can be in fluid communication with one another,
such as defining a serpentine flow path snaking through the airfoil
92 in the span-wise direction. It should be understood that the
ribs 130 and chambers 132 defined by the ribs 130 are exemplary and
should not be construed as limiting. It is contemplated that the
interior 110 or chambers 132 defined therein can also include a
plurality of plenums, circuits, micro-circuits, near wall cooling
systems, pin banks, or similar structures in non-limiting
examples.
[0037] A protuberance 134 can be disposed on the outer wall 112.
The protuberances 134 are discrete members, defining an increased
thickness for the wall 112. In non-limiting examples, the
protuberance can be radiused, rounded, conical, frustoconical,
bell-shaped, or non-linear. Additional examples of protuberances
can include, but are not limited to, radiused, circular, oval,
elliptical, spherical, ellipsoidal, or curvilinear. The
protuberances 134 can be integral to the outer wall 112, or can be
mounted thereto. In one non-limiting example, the protuberances 134
can be formed on the airfoil 92 using additive manufacturing. Any
number of protuberances 134 can be included on the outer wall 112
or the ribs 130 and can be organized in any manner, such as a
linear grouping in the span-wise or chord-wise direction, a
pattern, or random placement.
[0038] Additionally, it is contemplated that the protuberances 134
can be formed on an interior wall of the airfoil 92 or an engine
component. In one example, the protuberances 134 can be formed on
the rib 130. In other non-limiting examples, the protuberance 134
can be disposed on walls of cooling structures, such as
micro-circuits, cooling mesh, plenums, pin banks, or other
component structures requiring cooling.
[0039] A film hole 136 can extend through the protuberances 134.
The film holes 136 can extend through the protuberances 134 on the
outer wall 112 to provide a cooling film along the external surface
of the outer wall 112 for cooling the airfoil 92. Additionally, in
the case where the protuberance 134 is disposed on an interior wall
or structure, such as the rib 130 in one non-limiting example, the
film hole 136 can be a cooling hole such as a cross-over hole. In
additional examples, such a cooling hole can provide a flow of
cooling fluid among internal cavities or chambers of the engine
component, such as adjacent areas channels or a micro circuit.
[0040] The film hole 136 can be shaped to direct a flow of fluid
entering the film hole 132, passing through the film hole 132, or
exhausting from the film hole 132. Such shaping, for example, can
include a converging, diverging, or metering section to direct the
flow of fluid, in non-limiting examples. The converging section can
increase the flow velocity of the flow of fluid, the diverging
section can decrease the flow velocity of the flow of fluid, and
the metering section can meter the flow of fluid passing through
it. Additional shaping can include an expansion section or a
reduction section. The expansion section can include an increasing
cross-sectional area to form a diffusion section and the reduction
section can include a decreasing cross-sectional area.
Additionally, the shaping of the film hole 136 can include a
non-linear film hole 136. Such a film hole 136 could include curved
passages or follow the curvature of the protuberance 134.
[0041] It should be understood that the protuberance 134 could be
placed on any wall having opposing flows on opposing sides of the
wall 118. Additionally, the film hole 136 can pass through the
protuberance to provide a flow between the opposing sides of the
wall 118.
[0042] In an example where the engine component is not a blade,
such as a vane, combustion liner, shroud, or other component
requiring cooling, the protuberances can be disposed an any wall,
such as an internal or external wall, and can include a film hole
to provide a flow of fluid through such a wall for providing a
cooling film. Thus, it should be understood that the airfoil 92 as
illustrated is exemplary and non-limiting, and the protuberances
134 can have equal applicability in any other engine component
utilizing film holes.
[0043] FIG. 4 illustrates one protuberance 134 having a film hole
136 extending through the protuberance 134. The protuberance can be
non-rectilinear, including a non-linear surface extending from the
first side 144 to the inlet 150, with the inlet 150 being rounded,
transitioning into the film hole 136. In the top view shown, the
protuberance is circular. In other non-limiting examples, the
protuberance can be oval, elliptical, spherical, ellipsoidal, or
curvilinear. The protuberance 134 can be symmetrical, being even
about an axis parallel to the direction of a cooling fluid flow C.
The protuberance 134 is disposed on a wall 140. The wall 140 can be
the outer wall 112, for example, or any other component wall having
a film hole 136. The wall 140 has a nominal thickness 142, having a
first side 144 and a second side 146 defining a consistent nominal
thickness 142 between the sides 144, 146. The protuberance 134 is a
circular extension, extending from the first side 144 into a first
fluid flow 148. The film hole 136 is disposed in the center of the
circular protuberance 134, having an inlet 150 and an outlet 152. A
passage 154 couples the inlet 150 to the outlet 152.
[0044] The nominal thickness 142 can be the thickness for the wall
140 defined as the distance between the first and second side 144,
146. Such a nominal thickness 142 can be determined in many
different ways. For example, the nominal thickness 142 for the wall
can be a function of the thermal load on the wall 140, the airfoil
92, or the engine component. In other examples, the nominal
thickness 142 can be a function of a vibratory force acting on the
wall 140, a pressure differential between on opposing sides of the
wall 140, or the manufacturer required load for the wall 140 during
operation. It should be appreciated that the nominal thickness 142
can be determined by multiple methods, such that a minimum
operational thickness for the particular wall 140 is determined.
Additionally, the nominal thickness 142 is a minimal operational
thickness of the wall 140, being a function of the thermal load,
vibratory force, pressure differentials, load requirements, or
other similar method can be respective of minimal operation
requirements to maintain safe operation of the engine and
individual engine components. The nominal thickness 142 for the
wall 140 can reduce engine weight, increasing engine efficiency or
performance.
[0045] Referring to FIG. 5, the first side 144 can be a cool side
or cool surface, adjacent to a first fluid flow 148, such as a
cooling fluid flow C. The second side 146 can be a hot side, or a
hot surface, adjacent to a second fluid flow 170, such as a hot gas
flow H. The profile view of the protuberance 134 in FIG. 5
illustrates a height 160 of the protuberance 134. The height 160 is
the maximum distance the protuberance 134 extends from the wall
140. The height 160 can be determined in multiple different ways.
For example, the height 160 can be a function of the nominal
thickness 142. In one example, the height 160 can be at least 50%
of the nominal thickness 142. In another example, the height 160
can be equal to the nominal thickness 142, or greater. In yet
another example, the height can be equal to at least 100% of the
nominal thickness 142. It should be appreciated that the height 160
can be anywhere from 5% of the nominal thickness 142 to 200% of the
nominal thickness 142 or greater.
[0046] Alternatively, the height 160 can be a function of the film
hole 136. A length 162 of the film hole 136 can be defined as the
distance between the inlet 150 and the outlet 152. The height 160
can be a function of the length 162, where a particular film hole
136 can require a particular length 162 to provide an effective
flow of fluid. For example, the height 160 can be at least 50% of
the length 162. While FIG. 5 illustrates a linear film hole 136, it
should be understood that the film hole 136 need not be linear, and
with such a film hole, the length 162 can be measured as the
streamline distance between the inlet 150 and the outlet. It should
be appreciated that the film hole 136 as illustrated in FIG. 5 is a
perpendicular film hole 136. In the case of the non-linear film
hole, the length 162 will increase. As such, it should be
appreciated that in such a case, the height 160 can be at least 30%
of the length 162.
[0047] In yet another example, the height 160 can be determined as
a function of a diameter 164 of the film hole 136. A particular
diameter 164 for a film hole may be required by an engine
component, in order to keep structural integrity of the engine
component. The diameter 164 can require a particular length 162 for
the film hole 136 to maintain an effectiveness, defining a required
length-to-diameter ratio (L/D) for the film hole 136. As such, the
diameter 164, or the L/D ratio can be used to define the height 160
in order to provide sufficient film hole effectiveness.
[0048] Referring now to FIG. 6, a side profile view of the
protuberance 134 illustrates one orientation of the film hole 136.
The protuberance 134 can be conical, having a portion removed at
the inlet 150 of the film hole 136. Alternatively, it is
contemplated that the protuberance 135 can have a conic profile,
having the inlet 150 disposed on one of the sides of the
protuberance. The first cooling flow 148 can be a cooling fluid
flow C. A second fluid flow 170, adjacent to the second side 146 of
the wall 140, can be a hot gas flow H. The film hole 136 can be
angled in the direction of one of the first and second fluid flows
148, 170, or both. As such, the inlet 150 can be positioned
upstream of the outlet 152 relative to the cooling fluid flow C. In
an example where the film hole 136 is a film hole, the cooling
fluid flow C can be provided through the film hole 136 to the
second side 146 as a cooling film 172 to cool the engine
component.
[0049] FIG. 7 illustrates another embodiment of the protuberance
134 having a rounded dimension with a frustoconical shape at the
inlet 150 of the film hole 136. The film hole 136 includes the
inlet 150 disposed downstream of the outlet 152 relative to the
cooling fluid flow C. Such an orientation can be advantageous for
providing an effective film hole length as well as providing
multiple directional capabilities for exhausting a fluid from the
film hole 136. FIG. 8 illustrates another embodiment, having a
rounded protuberance having the film hole 136 with an inlet 150
offset from the center of the protuberance 134. The protuberance
134 can be divided into an upstream side 174 and a downstream side
176. The inlet 150 can be disposed on the upstream side 174. Such
an orientation can be advantageous for determining flow rate
entering the film hole 136. Alternatively, it is contemplated that
the inlet 150 can be disposed on the downstream side 176, or any
other position on the protuberance 134.
[0050] It should be appreciated that the position and orientation
of the film hole 136 of FIGS. 5-8 is exemplary. The position of the
inlet 150, outlet 152, and dimension of the passage 154 disposed
therebetween can be adapted to control flow rates through the film
hole 136, or adapt the length 162, diameter 164, or
length-to-diameter ratio for the film hole 136 to provide effective
cooling through the film hole 136. Furthermore, it is contemplated
that the film hole 136 can be provided with inlet shaping or outlet
shaping, to provide a more deterministic flow for a cooling fluid
passing through the cooling flow. Such an example would be a
diverging outlet which can provide a cooling fluid over a greater
cross-sectional area of the engine component.
[0051] Additionally, it should be appreciated that the protuberance
134 can have a height 160 dependent on portions of the engine
component, such as the nominal thickness 142, the length 162,
diameter 164, or L/D ratio of the film hole 136.
[0052] Referring now to FIG. 9, a recess 180 can be formed in the
protuberance 134. The recess 180 can be machined as part of the
protuberance 134, such as during additive manufacturing, or can be
removed from the protuberance 134 to form the recess 180. The
recess 180 can be symmetrical, such as a hemispherical shape, while
any shape is contemplated. In other non-limiting examples, the
recess 180 can be a rectilinear shape, or an arcuate or radiused
shape, or any combination thereof. The film hole 136 can be
disposed in the recess 180, having the inlet 150 at least partially
formed within the recess 180. The size or shape of the recess 180
can be used to control the flow rate of a flow of fluid provided to
the film hole 136, or to further reduce component weight in
combination with the nominal thickness 142 for the wall 140.
[0053] Referring now to FIG. 10, the protuberance 134 can be
asymmetrical, having an elongated or offset shape. Such as shape
may be desirable to optimize fluid flows within the engine
component or for directing a flow toward one or more film holes
136. Additionally the film hole 136 can be offset from the
direction of the first fluid flow 148 along the engine component.
The discrete direction of the first fluid flow 148 at the
protuberance 134 can be transposed on the protuberance 134 as a
transposed axis 190. The passage 154, shown in a linear example
while non-linear film holes are contemplated, can define a passage
axis 192. A film hole angle 194 can be defined between the
transposed axis 190 and the passage axis 192 to define the offset
relationship of the film hole 136 to the first fluid flow 148.
Additionally, the film hole 136 can be angled relative to a local
normal between the first and second sides 144, 146.
[0054] It should be understood that the offset orientation of the
film hole 136 or the protuberance 134 can be discrete, relative to
an adjacent flow of fluid which can change direction or magnitude
at different portions of the engine component. As such, a plurality
of protuberances 134 along the engine component can be aligned or
patterned, while some of the protuberances 134 or film holes 136
are offset from the direction of the first fluid flow 148 at a
portion of the engine component.
[0055] A method of cooling an engine component, such as the airfoil
92, can include a cool surface, such as the first side 144. The
method can include passing a cooling fluid flow C along the cool
surface 144 and providing at least a portion of the cooling fluid
flow C through a film hole, such as the film hole 136, in a
protuberance 134 extending from the cool surface 144. Providing at
least a portion of the cooling fluid flow C can include providing a
portion of the cooling fluid flow through a recess 180 in the
protuberance 134 prior to providing the cool fluid flow C to the
film hole 136. Additionally, providing a portion of the cooling
fluid flow C through the recess 180 can minimize dust accumulation
at the film hole 136 or along the cool surface 144 of the engine
component.
[0056] It should be understood that the airfoil 92 or other engine
component requiring cooling can utilizing the film hole 136 such as
a film hole disposed within the protuberance 134. The protuberance
134 provides for an increased thickness permitting an increased
film hole length 162 to provide effective cooling through the film
hole 136. At the same time, the use of a protuberance 134 permits
the remaining portions of the engine component to have a nominal
thickness 142, which reduces component weight, reducing overall
engine weight. A reduced weight provides for better engine
efficiency.
[0057] It should also be understood that the protuberances 134 are
discrete, having no greater an area than necessary to provide for
the casting, drilling, or otherwise forming the film holes 136
through the protuberances 134 in order to have an increased length,
diameter, or L/D ratio for the film hole 136 which would otherwise
be unachievable within the nominal thickness 142 of the engine
component, due to the nominal thickness 142 to manufacturing
capabilities of the engine component at the nominal thickness
142.
[0058] Furthermore, the protuberances 134 are radiused, reducing
drag or resistance caused by the extension of the protuberance 134
into the flow of fluid, such as the cooling fluid flow C, adjacent
the protuberance 134. Further still, the radiused protuberances 134
or recesses 180 therein can provide for reduced dust accumulation,
increasing component lifetime or reducing required maintenance.
[0059] It should be appreciated that the airfoil, engine
components, protuberances, or film holes described herein can be
formed by additive manufacturing. Such manufacturing can be used to
develop the intricate details of the aforementioned, such as
specific film hole shaping without the poor yields of such
manufacturing as casting, or the imperfections associated with
other manufacturing methods such as film hole drilling.
[0060] It should be appreciated that application of the disclosed
design is not limited to turbine engines with fan and booster
sections, but is applicable to turbojets and turbo engines as
well.
[0061] This written description uses examples to disclose the
invention, including the best mode, and also to enable any person
skilled in the art to practice the invention, including making and
using any devices or systems and performing any incorporated
methods. The patentable scope of the invention is defined by the
claims, and may include other examples that occur to those skilled
in the art. Such other examples are intended to be within the scope
of the claims if they have structural elements that do not differ
from the literal language of the claims, or if they include
equivalent structural elements with insubstantial differences from
the literal languages of the claims.
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