BACKGROUND
[0001] This disclosure relates to a gas turbine engine, and more particularly to a gas turbine
engine component having an internal cooling circuit. The internal cooling circuit
may include a cooling cavity and a rib that separates the cooling cavity into separate
portions.
[0002] Gas turbine engines typically include a compressor section, a combustor section and
a turbine section. During operation, air is pressurized in the compressor section
and is mixed with fuel and burned in the combustor section to generate hot combustion
gases. The hot combustion gases are communicated through the turbine section, which
extracts energy from the hot combustion gases to power the compressor section and
other gas turbine engine loads.
[0003] Because they are commonly exposed to hot combustion gases, many gas turbine engine
components employ internal cooling circuits that channel a dedicated cooling fluid
for cooling regions of the component. Thermal energy is transferred from the component
to the cooling fluid to cool the component.
SUMMARY
[0004] A component for a gas turbine engine according to an example of the present disclosure
includes a wall that extends about a cooling cavity. The cooling cavity is a dual-fed
cavity that is fed from at least two different locations. A rib separates the cooling
cavity into a first portion and a second portion that is fluidly isolated from the
first portion. The component is an airfoil. The first portion is fed with a first
cooling fluid from a first coolant source, and the second portion is fed with a second,
different cooling fluid from a second coolant source. The first and second coolant
sources are separate and distinct from the component.
[0005] In a further non-limiting embodiment of any of the foregoing embodiments, the first
and second coolant sources are stages of a compressor section or a turbine section.
[0006] In a further non-limiting embodiment of any of the foregoing embodiments, the wall
circumscribes the cooling cavity.
[0007] In a further non-limiting embodiment of any of the foregoing embodiments, the rib
is offset from a midspan of the airfoil.
[0008] In a further non-limiting embodiment of any of the foregoing embodiments, the rib
is defined at a location between 10% and 90% span or between 30% and 70% span of the
airfoil.
[0009] A further non-limiting embodiment of any further embodiments includes a plurality
of cooling features defined along the wall.
[0010] In a further non-limiting embodiment of any of the foregoing embodiments, the plurality
of cooling features include pedestals that extend between opposed surfaces of the
cooling cavity.
[0011] In a further non-limiting embodiment of any of the foregoing embodiments, the plurality
of cooling features include trip strips that protrude from surfaces of the cooling
cavity.
[0012] In a further non-limiting embodiment of any of the foregoing embodiments, the rib
is skewed in a radial direction towards one of an inner diameter and an outer diameter
of the airfoil.
[0013] In a further non-limiting embodiment of any of the foregoing embodiments, the first
portion is an outer diameter portion and the second portion is an inner diameter portion
of the cooling cavity.
[0014] In a further non-limiting embodiment of any of the foregoing embodiments, the outer
diameter portion defines a first serpentine passage and the inner diameter portion
defines a second serpentine passage. The first and second serpentine passages are
bounded by the rib.
[0015] In a further non-limiting embodiment of any of the foregoing embodiments, the rib
is spaced apart from leading and trailing edges of the airfoil.
[0016] In a further non-limiting embodiment of any of the foregoing embodiments, the rib
extends between leading and trailing edges of the airfoil.
[0017] In a further non-limiting embodiment of any of the foregoing embodiments, the first
portion is circumferentially offset from the second portion.
[0018] In a further non-limiting embodiment of any of the foregoing embodiments, the rib
connects between opposing sides of the wall.
[0019] In a further non-limiting embodiment of any of the foregoing embodiments, the rib
extends in an axial direction inside of the cooling cavity.
[0020] A further non-limiting embodiment of any further embodiments includes a plurality
of openings through portions of the wall associated with both the first portion and
the second portion, wherein the plurality of openings are film cooling holes.
[0021] A gas turbine engine according to an example of the present disclosure includes a
component that defines a cooling circuit that cools the component with a cooling fluid.
The cooling circuit is disposed inside an airfoil of the component. The cooling circuit
has a cooling cavity disposed inside of the component, and an axial rib that divides
the cooling cavity into a first portion and a second portion that is separate from
the first portion. The first portion is fed with a first cooling fluid from a first
coolant source, and the second portion is fed with a second, different cooling fluid
from a second coolant source. The first and second coolant sources are separate and
distinct from the component.
[0022] In a further non-limiting embodiment of any of the foregoing embodiments, the rib
fluidly isolates the first portion from the second portion.
[0023] In further non-limiting embodiment of any further embodiments, a mid-turbine frame
includes the airfoil.
[0024] A component according to an exemplary aspect of the present disclosure includes,
among other things, a wall that extends about a cooling cavity. The cooling cavity
is a dual-fed cavity that is fed from at least two different locations. A rib separates
the cooling cavity into a first portion and a second portion that is fluidly isolated
from the first portion.
[0025] In a further non-limiting embodiment of the foregoing component, the component is
one of a vane, a blade, a blade outer air seal (BOAS), and a liner.
[0026] In a further non-limiting embodiment of either of the foregoing components, the wall
circumscribes the cooling cavity.
[0027] In a further non-limiting embodiment of any of the foregoing components, the first
portion of the cooling cavity is fed with a first cooling fluid and the second portion
of the cooling cavity is fed with a second, different cooling fluid.
[0028] In a further non-limiting embodiment of any of the foregoing components, the first
portion is an outer diameter portion and the second portion is an inner diameter portion
of the cooling cavity.
[0029] In a further non-limiting embodiment of any of the foregoing components, the first
portion is circumferentially offset from the second portion.
[0030] In a further non-limiting embodiment of any of the foregoing components, the rib
connects between opposing sides of the wall.
[0031] In a further non-limiting embodiment of any of the foregoing components, the rib
extends in an axial direction inside of the cooling cavity.
[0032] In a further non-limiting embodiment of any of the foregoing components, the components
include a plurality of openings through portions of the wall associated with both
the first portion and the second portion.
[0033] In a further non-limiting embodiment of any of the foregoing components, the plurality
of openings are film cooling holes.
[0034] A gas turbine engine according to another exemplary aspect of the present disclosure
includes, among other things, a component that defines a cooling circuit configured
to cool the component with a cooling fluid. The cooling circuit includes a cooling
cavity disposed inside of the component and an axial rib that divides the cooling
cavity into a first portion and a second portion that is separate from the first portion.
[0035] In a further non-limiting embodiment of the foregoing gas turbine engine, the rib
fluidly isolates the first portion from the second portion.
[0036] In a further non-limiting embodiment of either of the foregoing gas turbine engines,
the cooling circuit is disposed inside an airfoil of the component.
[0037] In a further non-limiting embodiment of any of the foregoing gas turbine engines,
the cooling circuit is disposed inside of a body of the component.
[0038] In a further non-limiting embodiment of any of the foregoing gas turbine engines,
the cooling circuit includes a plurality of openings that expel the cooling fluid
from both the first portion and the second portion.
[0039] A method of cooling a gas turbine engine component according to another exemplary
aspect of the present disclosure includes, among other things, dividing a cooling
cavity disposed inside the gas turbine engine component into a first portion and a
second portion with a rib, communicating a first cooling fluid from a first location
into the first portion, and communicating a second cooling fluid from a second location
into the second portion.
[0040] In a further non-limiting embodiment of the foregoing method, the method includes
expelling the first and second cooling fluids from the cooling cavity through a plurality
of openings.
[0041] In a further non-limiting embodiment of either of the foregoing methods, the second
cooling fluid is separate from the first cooling fluid.
[0042] In a further non-limiting embodiment of any of the foregoing methods, the rib fluidly
isolates the first portion from the second portion.
[0043] In a further non-limiting embodiment of any of the foregoing methods, one of the
first portion and the second portion of the cooling cavity is positioned in a high
pressure area and the other of the first portion and the second portions is positioned
in a low pressure area.
[0044] The embodiments, examples and alternatives of the preceding paragraphs, the claims,
or the following description and drawings, including any of their various aspects
or respective individual features, may be taken independently or in any combination.
Features described in connection with one embodiment are applicable to all embodiments,
unless such features are incompatible.
[0045] The various features and advantages of this disclosure will become apparent to those
skilled in the art from the following detailed description. The drawings that accompany
the detailed description can be briefly described as follows.
BRIEF DESCRIPTION OF THE DRAWINGS
[0046]
Figure 1 illustrates a schematic, cross-sectional view of a gas turbine engine.
Figure 2 illustrates a gas turbine engine component according to a first embodiment
of this disclosure.
Figure 3 illustrates a cross-sectional view through Section A-A of Figure 2.
Figure 4 illustrates a flow separating rib of a gas turbine engine component cooling
circuit.
Figure 5A illustrates cooling features of the component of Figure 2.
Figure 5B illustrates a cross-sectional view through Section B-B of Figure 5A.
Figure 6 illustrates a gas turbine engine component according to a second embodiment
of this disclosure.
Figure 7 illustrates another flow separating rib.
Figure 8 illustrates a gas turbine engine component according to an embodiment of
this disclosure.
Figure 9 illustrates a gas turbine engine component according to another embodiment
of this disclosure.
DETAILED DESCRIPTION
[0047] This disclosure relates to a gas turbine engine component that includes an internal
cooling circuit. The cooling circuit employs one or more cooling cavities disposed
inside of the component. A flow separating rib is positioned to divide the cooling
cavity into at least two portions. The cooling cavity may be fed with separate cooling
fluids at opposite sides of the cavity. These opposite fluid flows are fluidly isolated
between the first portion and the second portion by the rib in order to maintain a
constant fluid flow within each portion even where pressure differentials may exist
between the opposite sides. A more evenly cooled part is achieved by maintaining constant
fluid flows within each portion of the cooling cavity. These and other features are
discussed in greater detail herein.
[0048] Figure 1 schematically illustrates a gas turbine engine 20. The gas turbine engine
20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section
22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative
engines might include an augmenter section (not shown) among other systems or features.
The fan section 22 drives air along a bypass flow path B in a bypass duct defined
within a nacelle 15, while the compressor section 24 drives air along a core flow
path C for compression and communication into the combustor section 26 then expansion
through the turbine section 28. Although depicted as a two-spool turbofan gas turbine
engine in the disclosed non-limiting embodiment, it should be understood that the
concepts described herein are not limited to use with two-spool turbofans as the teachings
may be applied to other types of turbine engines including three-spool architectures.
[0049] The exemplary engine 20 generally includes a low speed spool 30 and a high speed
spool 32 mounted for rotation about an engine central longitudinal axis A relative
to an engine static structure 36 via several bearing systems 38. It should be understood
that various bearing systems 38 at various locations may alternatively or additionally
be provided, and the location of the bearing systems 38 may be varied as appropriate
to the application.
[0050] The low speed spool 30 generally includes an inner shaft 40 that interconnects a
fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine
46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism,
which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48
to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool
32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor
52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary
gas turbine 20 between the high pressure compressor 52 and the high pressure turbine
54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally
between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine
frame 57 further supports bearing systems 38 in the turbine section 28. The inner
shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems
38 about the engine central longitudinal axis A which is collinear with their longitudinal
axes.
[0051] The core airflow is compressed by the low pressure compressor 44 then the high pressure
compressor 52, mixed and burned with fuel in the combustor 56, then expanded over
the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57
includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally
drive the respective low speed spool 30 and high speed spool 32 in response to the
expansion. It will be appreciated that each of the positions of the fan section 22,
compressor section 24, combustor section 26, turbine section 28, and fan drive gear
system 48 may be varied. For example, gear system 48 may be located aft of combustor
section 26 or even aft of turbine section 28, and fan section 22 may be positioned
forward or aft of the location of gear system 48.
[0052] The engine 20 in one example is a high-bypass geared aircraft engine. In a further
example, the engine 20 bypass ratio is greater than about six (6), with an example
embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic
gear train, such as a planetary gear system or other gear system, with a gear reduction
ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio
that is greater than about five. In one disclosed embodiment, the engine 20 bypass
ratio is greater than about ten (10:1), the fan diameter is significantly larger than
that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure
ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio
is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure
at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The gear
system 48 may be an epicycle gear train, such as a planetary gear system or other
gear system, with a gear reduction ratio of greater than about 2.3:1. It should be
understood, however, that the above parameters are only exemplary of one embodiment
of a geared architecture engine and that the present invention is applicable to other
gas turbine engines including direct drive turbofans and turboshafts.
[0053] A significant amount of thrust is provided by the bypass flow B due to the high bypass
ratio. The fan section 22 of the engine 20 is designed for a particular flight condition
-- typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight
condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption
- also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the
industry standard parameter of lbm of fuel being burned divided by lbf of thrust the
engine produces at that minimum point. "Low fan pressure ratio" is the pressure ratio
across the fan blade alone, without a Fan Exit Guide Vane ("FEGV") system. The low
fan pressure ratio as disclosed herein according to one non-limiting embodiment is
less than about 1.45. "Low corrected fan tip speed" is the actual fan tip speed in
ft/sec divided by an industry standard temperature correction of [(Tram °R) / (518.7
°R)]
0.5. The "Low corrected fan tip speed" as disclosed herein according to one non-limiting
embodiment is less than about 1,150 ft/second (350.5 meters/second).
[0054] Each of the compressor section 24 and the turbine section 28 may include alternating
rows of rotor assemblies and vane assemblies (shown schematically) that define a plurality
of stages 31 of the compressor section 24 and a plurality of stages 33 of the turbine
section 28. For example, the rotor assemblies can carry a plurality of rotating blades
25, while each vane assembly can carry a plurality of vanes 27 that extend into the
core flow path C. The blades 25 may either create or extract energy in the form of
pressure from the core airflow as it is communicated along the core flow path C. The
vanes 27 direct the core airflow to the blades 25 to either add or extract energy.
[0055] Figure 2 illustrates a component 60 that can be incorporated into a gas turbine engine,
such as one or more airfoils including blades 25, vanes 27 or airfoils 59 of the gas
turbine engine 20 of Figure 1. In this non-limiting embodiment, the component 60 is
represented as a turbine vane, which can be utilized in a stage 33 of the high pressure
turbine 54 or low pressure turbine 46, for example. However, the teachings of this
disclosure are not limited to turbine vanes and could extend to other components of
a gas turbine engine, including but not limited to, other vanes, blades, blade outer
air seals (BOAS) (see, for example, the BOAS illustrated in Figure 6), or other components
such as a blade or vane of the compressor section 24.
[0056] In one embodiment, the component 60 includes an outer platform 62, an inner platform
64, and an airfoil 66 that extends in a chordwise direction X between leading and
trailing edges L/E, T/E, in a radial direction R between the outer platform 62 and
the inner platform 64, and in a thickness direction T between pressure and suction
sides P, S. The thickness direction T is generally perpendicular to the chordwise
and radial directions X, R. The outer platform 62 connects the component 60 to an
engine casing (not shown) and the inner platform 64 affixes a radially inboard portion
of the component 60 to securely position the component 60 within the core flow path
C.
[0057] The component 60 can include one or more internal cooling cavities 72 that are disposed
inside of the component 60. In one embodiment, the cooling cavities 72 extend inside
of the airfoil 66 of the component 60. In another embodiment, one or more cooling
cavities may extend inside a body or platform portion of the component, such as in
components that do not have an airfoil (e.g., a BOAS, liner, panel, etc.).
[0058] The internal cooling cavities 72 define a cooling circuit 74 for cooling the component
60. The illustrated cooling circuit 74 represents but one non-limiting example of
many potential cooling circuits. In other words, the component 60 could be manufactured
to include various alternatively shaped and sized cooling passages as part of an internal
circuitry within the scope of this disclosure.
[0059] Figure 3 illustrates a portion of the cooling circuit 74 described in Figure 2. In
this embodiment, a cross-section through a cooling cavity 72 is depicted. The cooling
cavity 72 is generally surrounded or circumscribed by a wall 80. Opposing sides 90,
92 of the wall 80 define flow boundaries of the cooling cavity 72. The wall 80 may
embody any of a variety of sizes and shapes within the scope of this disclosure.
[0060] A rib 82 may axially extend inside of the cooling cavity 72 to separate the cooling
cavity 72 into a first portion 84 and a second portion 86. In one embodiment, the
first portion 84 is an outer diameter portion of the cooling cavity 72 and the second
portion 86 is an inner diameter portion. However, other configurations are also contemplated
as being within the scope of this disclosure, including but not limited to circumferentially
spaced portions such that the first portion 84 is circumferentially offset from the
second portion 86 (see, e.g., Figure 6) and axially spaced portions (see, e.g., Figure
7).
[0061] Rib 82 can be defined at various span positions of the airfoil 66 relative to the
radial direction R. Span position may be relative to the inner platform 64, such as
0% span at the inner platform 64 and 100% span at outer platform 62 (or a tip of blade
25). In an embodiment, rib 82 is defined at about 50% span (i.e., mid-span position).
In other embodiments, rib 82 is defined at a location between 10-90% span, or more
narrowly between 30-70% span such that rib 82 is radially offset from the mid-span
position. The rib 82 can be situated relatively closer to one of the platforms 62,
64 to vary the relative sizes of the first and second portions 84, 86 of the cooling
cavity 72.
[0062] In one embodiment, the rib 82 extends between the opposing sides 90, 92 of the wall
80 to completely seal and separate the first portion 84 of the cooling cavity 72 from
the second portion 86 of the cooling cavity 72. Said another way, the rib 82 is a
solid flow separator that fluidly isolates the first portion 84 from the second portion
86 of the cooling cavity 72.
[0063] The rib 82 may be positioned at a mid-span M of the cooling cavity 72. The actual
location of the rib 82 could vary part-by-part and may depend on pressure differentials
that exist between the first portion 84 and the second portion 86 of the cooling cavity
72, among other factors.
[0064] In one embodiment, the cooling cavity 72 is a dual-fed cavity that is fed with a
cooling fluid at both of its opposite sides (i.e., fed from two distinct locations).
In embodiments, the first and second portions 84, 86 of the cooling cavity 72 are
fed from distinct first and second coolant sources 35, 37 that communicate coolant
from different locations. The first and second coolant sources 35, 37 are separate
and distinct from the component 60. For example, the first portion 84 of the cooling
cavity 72 may be fed with a first cooling fluid F1 from the first coolant source 35,
such as a first bleed airflow from a first stage 31A of the compressor section 24
(Figure 1), and the second portion 86 of the cooling cavity 72 may be fed with a second
cooling fluid F2 from the second coolant source 37, such as a second bleed airflow
from a second, different stage 31B of the compressor section 24 (Figure 1). In other
words, the cooling fluids F1 and F2 may be separate from one another.
[0065] The first and second coolant sources 35, 37 can be defined by various locations or
components of the engine 20. In embodiments, one of the first and second bleed airflows
is supplied by a first stage 31A of the compressor section 24, and another one of
the first and second bleed airflows is supplied by a second stage 31B of the compressor
section 24. In some embodiments, the first stage 31A is an upstream stage of high
pressure compressor 52 that supplies the cooling fluid Flat a relatively low pressure,
including a forwardmost or intermediate stage, and the second stage 31B is a downstream
stage of the high pressure compressor 52 at a relatively higher pressure than the
first stage 31A, such as an intermediate or aftmost stage. In another embodiment,
the first coolant source 35 supplies the cooling fluid F1 from the bypass flow path
B at a relatively lower pressure and temperature than the second coolant source 37.
In embodiments, the coolant source 35/37 is a stage 31 of the low pressure compressor
44 or another portion of the engine 20. In yet another embodiment, the coolant source
35/37 is a stage 33 of the turbine section 28, such as a first stage 33A and a second
stage 33B of the turbine section 28. One of first and second stages 33A, 33B can be
an upstream stage of the turbine section 28, such as a stage of the high pressure
turbine 54, and another one of the first and second stages 33A, 33B can be a downstream
stage of the turbine section 28, such as a stage of the low pressure turbine 46.
[0066] In the embodiment illustrated in Figure 3, an inlet 85 of the first portion 84 of
the cooling cavity 72 is positioned in a relatively high pressure area and an inlet
87 of the second portion 86 of the cooling cavity 72 is positioned at a relatively
low pressure area. Of course, an opposite configuration is also possible in which
the inlet 85 of the first portion 84 is located at a relatively low pressure area
and the inlet 87 of the second portion 86 is within a relatively high pressure area
(see Figure 4). Despite the pressure differentials that may exist at the inlets 85,
87, flow of the first and second cooling fluids F1, F2 remains constant within both
the first portion 84 and the second portion 86 because these portions are sealed from
one another by the rib 82. Maintaining consistent flow in this manner results in relatively
consistent Mach numbers, pressure losses, heat transfer and metal temperatures throughout
the cooling cavity 72. In other words, the component 60 is more evenly cooled by virtue
of the flow separating rib 82.
[0067] A plurality of openings 96 may extend through portions of the wall 80 associated
with both the first portion 84 and the second portion 86 of the cooling cavity 72.
The cooling fluids F1, F2 that are circulated in the first and second portions 84,
86, respectively, may be expelled through the openings 96. In one embodiment, the
openings 96 are film cooling holes. In another embodiment, the openings 96 are slots.
Any type of opening may extend through the wall 80 for expelling the cooling fluids
F1, F2 from the cooling cavity 72.
[0068] Referring to Figures 5A and 5B, one or more cooling features 93 are defined along
the wall 80. The cooling features 93 can extend from surfaces of the wall 80, such
as one of sides 90, 92, and into the first and/or second portions 84, 86 of the cooling
cavity 72. The cooling features 93 can be situated to provide additional surface area
for convective cooling and/or direct or meter fluid flow within or through localized
regions of the first and/or second portions 84, 86. Example cooling features 93 can
include pedestals 93A extending between opposed surfaces of the cooling cavity 72,
for example. Other cooling features 93 can include features having a curved or complex
geometry such as teardrop shaped features 93B to direct flow through the cooling cavity
72, and sinusoidal shaped features 93C and trip strips 93D protruding from surfaces
of the cooling cavity 72 to cause turbulence in the flow of cooling fluid F1/F2. Other
example cooling features 93 can include recesses such as dimples 93E extending inwardly
from surfaces of the wall 80.
[0069] Figure 6 illustrates another component 160 that can be incorporated into a gas turbine
engine. In this disclosure, like reference numbers designate like elements where appropriate
and reference numerals with the addition of 100 or multiples thereof designate modified
elements that are understood to incorporate the same features and benefits of the
corresponding original elements.
[0070] In this embodiment, the component 160 is represented as a BOAS. The BOAS can be situated
adjacent to a tip of one the blades 25 and can be utilized to seal or otherwise bound
the core flow path C (Figure 1), for example. The component 160 includes a body 161
having a radially inner face 163 and a radially outer face 165. The radially inner
face 163 and the radially outer face 165 extend circumferentially between a first
mate face 167 and a second mate face 169 and extend axially between a leading edge
171 and a trailing edge 173.
[0071] A cooling cavity 172 may be disposed inside the body 161. The cooling cavity 172
of this embodiment circumferentially extends between the first mate face 167 and the
second mate face 169. A wall 180 may extend about the cooling cavity 172. The cooling
cavity 172 is divided into a first portion 184 and a second portion 186 by a rib 182.
The rib 182 fluidly isolates the first portion 184 from the second portion 186.
[0072] The first portion 184 of the cooling cavity 172 may be fed with a first cooling fluid
F1 at a location adjacent to the first mate face 167 and the second portion 186 may
be fed with a second cooling fluid F2 at a location adjacent to the second mate face
169. The rib 182 is adapted to maintain these split flows at relatively constant flow
levels despite potential pressure differentials that may exist between the first mate
face 167 and the second mate face 169.
[0073] Figure 8 illustrates another component 260 that can be incorporated into a gas turbine
engine. In the illustrated embodiment, the component includes an airfoil 266 that
extends in a chordwise direction X between leading and trailing edges L/E, T/E. Rib
282 extends in the chordwise direction X between the leading and trailing edges L/E,
T/E to fluidly isolate first and second portions 284, 286 such that the cooling fluids
F1, F2 do not intermix in cooling cavity 272.
[0074] One or more radially extending ribs 295 extend from walls 280 to establish serpentine
passages in first and second portions 284, 286. One or more of the ribs 295 can extend
or be spaced apart from rib 282 to define sections of the respective serpentine passages,
with rib 282 bounding each of the serpentine passages. In the illustrated embodiment
of Figure 8, rib 282 has a major component that extends in the chordwise direction
X, and ribs 295 each have a major component that extends in a radial direction R.
[0075] An axis E1 of rib 82 can be oriented relative to an axis E2 that extends in the radially
direction R through the airfoil 266 to establish a radial angle α. In embodiments,
the axis E1 is substantially perpendicular to the axis E2. In other embodiments, the
angle α is non-perpendicular such that the axis E1 has a component that extends in
the radial direction R and the axis E1 is skewed toward an inner diameter 266A or
an outer diameter 266B of the airfoil 266. The axis E1 can be skewed to adjust a pressure
of the cooling fluid F1, F2 that is discharged by opening(s) 296. In embodiments,
the radial angle α is between 10-30° or between 70-90°.
[0076] Figure 9 illustrates yet another component 360 that can be incorporated into a gas
turbine engine. In the illustrated embodiment, the component includes an airfoil 366
including a rib 382 that fluidly isolates first and second portions 384, 386 of cooling
cavity 372. Rib 382 spans between ribs 395 and is spaced apart from leading and trailing
edges L/E, T/E. Flow of cooling fluids F1, F2 can be directed from inlets 385, 387
and towards the rib 382. Wall 380 can include one or more openings 396 to expel the
cooling fluids F1, F2 from the first and second portions 384, 386 of the internal
cavity 372 to provide film cooling, for example.
[0077] Although the different non-limiting embodiments are illustrated as having specific
components, the embodiments of this disclosure are not limited to those particular
combinations. It is possible to use some of the components or features from any of
the non-limiting embodiments in combination with features or components from any of
the other non-limiting embodiments.
[0078] It should be understood that like reference numerals identify corresponding or similar
elements throughout the several drawings. It should also be understood that although
a particular component arrangement is disclosed and illustrated in these exemplary
embodiments, other arrangements could also benefit from the teachings of this disclosure.
[0079] The foregoing description shall be interpreted as illustrative and not in any limiting
sense. A worker of ordinary skill in the art would understand that certain modifications
could come within the scope of this disclosure. For these reasons, the following claims
should be studied to determine the true scope and content of this disclosure.
1. A component for a gas turbine engine, comprising:
a wall that extends about a cooling cavity, wherein said cooling cavity is a dual-fed
cavity that is fed from at least two different locations; and
a rib that separates said cooling cavity into a first portion and a second portion
that is fluidly isolated from said first portion;
wherein the component is an airfoil; and
wherein said first portion is fed with a first cooling fluid from a first coolant
source, and said second portion is fed with a second, different cooling fluid from
a second coolant source, the first and second coolant sources being separate and distinct
from the component.
2. The component as recited in claim 1, wherein said first and second coolant sources
are stages of a compressor section or a turbine section.
3. The component as recited in claim 1 or 2, wherein said wall circumscribes said cooling
cavity.
4. The component as recited in claim 1, 2, or 3, wherein said rib is offset from a midspan
of said airfoil,
wherein, optionally, said rib is defined at a location between 10% and 90% span or
between 30% and 70% span of said airfoil.
5. The component as recited in any preceding claim, further comprising a plurality of
cooling features defined along said wall,
wherein, optionally, said plurality of cooling features include pedestals that extend
between opposed surfaces of said cooling cavity, or
wherein said plurality of cooling features include trip strips that protrude from
surfaces of said cooling cavity.
6. The component as recited in any preceding claim, wherein said rib is skewed in a radial
direction towards one of an inner diameter and an outer diameter of said airfoil.
7. The component as recited in any preceding claim, wherein said first portion is an
outer diameter portion and said second portion is an inner diameter portion of said
cooling cavity,
wherein, optionally, said outer diameter portion defines a first serpentine passage
and said inner diameter portion defines a second serpentine passage, the first and
second serpentine passages bounded by said rib.
8. The component as recited in claim 7, wherein said rib is spaced apart from leading
and trailing edges of said airfoil,
or wherein said rib extends between leading and trailing edges of said airfoil.
9. The component as recited in any preceding claim, wherein said first portion is circumferentially
offset from said second portion.
10. The component as recited in any preceding claim, wherein said rib connects between
opposing sides of said wall.
11. The component as recited in any preceding claim, wherein said rib extends in an axial
direction inside of said cooling cavity.
12. The component as recited in any preceding claim, comprising a plurality of openings
through portions of said wall associated with both said first portion and said second
portion, wherein said plurality of openings are film cooling holes.
13. A gas turbine engine, comprising the component as recited in any preceding claim.
14. The gas turbine engine as recited in claim 13, wherein said rib fluidly isolates said
first portion from said second portion.
15. The gas turbine engine as recited in claim 13 or 14, further comprising a mid-turbine
frame that includes said airfoil.