Note: Descriptions are shown in the official language in which they were submitted.
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METHOD FOR PREVENTING BACKFLOW AND FORMING A
COOLING LAYER IN AN AIRFOIL
FIELD OF THE INVENTION
The invention relates to a method for preventing backflow and forming a
cooling layer in an airfoil. More specifically, the invention relates to
method for
preventing backflow and forming a cooling layer where separation regions are
formed
at a cooling slot inlet.
BACKGROUND OF THE INVENTION
Gas turbine engines extract energy from a stream of hot combustion gases
that flow through a flow path defined by the turbine. A typical turbine engine
includes at least one stage of turbine blades and one stage of vanes spaced
from the
turbine blades. Each turbine stage comprises a plurality of turbine blades or
airfoils
spaced circumferentially around, and extending radially outward from, a
rotatable hub
or disk so that a portion of each turbine blade extends into the flow path and
comes in
contact with the flow of the combustion gases through the flow path. In
practice,
turbine engines comprise multiple stages of vanes and blades.
During engine operation it is necessary to cool turbine blades and vanes to
improve their ability to endure extended exposure to the hot combustion gases.
Frequently, blade cooling is achieved by creating a cooling film along the
blade. In
order to develop the desired cooling film, the turbine blades include one or
more
rows of spanwisely distributed cooling air supply holes, referred to as film
holes and
these holes are located along the surface of the blade. The film holes
penetrate the
walls of the airfoil to establish fluid flow communication between cooling
fluid
passing through the interior of the blade and the externally located hot
combustion
gases. Additionally, the blade includes a plurality of cooling slots spaced
along the
trailing edge of the blade. The slots are located within the blade and have
outlet
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openings spaced along the trailing blade edge. During engine operation,
cooling fluid
or air is typically supplied to the blade by a compressor upstream of the
airfoil
compressor. The cooling air passes through the interior of the blade,
including the
slots, and exits the blade through the film holes and outlet openings. The
cooling air
flows from the holes and the cooling slots as a series of discrete jets. The
air
discharged from the slots and holes is intended to form the cooling film along
the
blade surface.
A conventional airfoil in Figure 2 provides an example of a turbine blade
70 of the prior art. As shown in Figure 2, the blade 70 includes leading edge
71,
trailing edge 72 and a plurality of parallel cooling slots 75 at the blade
trailing edge.
In prior art blade 70, each of the cooling slots has an associated axially
extending slot
reference line 80. Each slot has an inlet 62 and an outlet 63. The outlet is
located at
the trailing edge of the blade. The inlet and outlet are located at
substantially the same
radial position along the radially extending blade length. For simplicity, in
Figure 2
reference lines 80 are provided for fewer than all of the slots however, the
reference
lines apply to all of the cooling slots 75. Each of the cooling slots is
parallel to its
respective reference line 80.
Film cooling provides an effective means for controlling the temperature
of airfoil surfaces, however in practice, cooling films are difficult to
effectively
produce. One shortcoming associated with the conventional parallel cooling
slot
orientation is that the blade is susceptible to the backflow of combustion
gases
through the cooling slots. Backflow occurs when the static pressure of the
cooling air
does not exceed the static pressure of the combustion gases flowing through
the flow
path. When backflow occurs, the combustion gases flow through the cooling
holes
and into the cooling slots
In order to overcome the susceptibility to backflow in conventional blades,
the high cooling air is discharged from the slots and holes at a high pressure
to
prevent backflow. The relatively high pressure cooling air can cause the
cooling air to
be discharged from the cooling slots with a velocity that prevents the cooling
air from
effectively adhering to the surface and edges of the airfoil. As a result, the
desired
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cooling film does not form on the blade. Instead the cooling air is directly
flowed into
and entrained with the combustion gases. As a result, a portion of the blade
airfoil
surface immediately downstream of each cooling hole or cooling slot is exposed
to
the combustion gases and is not protected by a cooling film. Additionally,
each of
the cooling air jets may locally intersect and bifurcate the stream of
combustion gases
into a pair of minute, oppositely swirling vortices. The combustion gases
enter the
exposed portion of the airfoil and can cause irreparable damage to the
airfoil. The
intense heat of backflow gases can quickly and irreparably damage an airfoil.
What is therefore needed is an airfoil with cooling slots arranged in a
manner that promotes effective formation of a cooling film along the airfoil
surface.
BRIEF DESCRIPTION OF THE INVENTION
A method for preventing backflow and forming a cooling layer in an
airfoil, said airfoil comprising a leading edge, a trailing edge, a blade tip
at a first
blade end and a blade root at a second blade end, the tip and root being
separated by a
radial distance, a cooling passage extending between the leading and trailing
edges,
and at least one cooling slot having an inlet end in fluid receiving
communication
with the cooling passage and an outlet end proximate the trailing edge, and
wherein
for the at least one slot the inlet and outlet are located at different radial
locations
within the airfoil, the method comprising flowing a cooling fluid in a first
direction
through said cooling passage toward a cooling slot, flowing the cooling medium
in a
second direction through said cooling passage toward the cooling slot, forming
a
separation region proximate the cooling slot inlet and flowing the cooling
medium
through said cooling slot and out the slot to form a layer at the trailing
edge of the
airfoil.
Thus, by the described invention improved the cooling of an airfoil is
achieved. This improvement is accomplished by metering airflow through a
plurality
of angled cooling slots. Also, instead of drilling cooling slots into an
airfoil, one may
cast cooling slots into an airfoil and thus decrease manufacturing costs and
increase
the beneficial variability of cooling slots at their creation.
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BRIEF DESCRIPTION OF THE DRAWINGS
While the specification concludes with claims particularly pointing out and
distinctly claiming the invention, it is believed that the embodiments set
forth herein
will be better understood from the following description in conjunction with
the
accompanying figures, in which like reference numerals identify like elements
and in
which:
Figure 1 shows a schematic representation of a gas turbine;
Figure 2 is a sectional view of a prior art turbine blade comprising a
conventional cooling slot configuration;
Figure 3 is a sectional view of a turbine blade comprising a cooling slot
arrangement according to an embodiment of the present invention;
Figure 4 is a sectional view of a turbine blade comprising an alternate
embodiment of the invention; and
Figure 5 is an enlarged detailed view of the portion of Figure 4 within the
circle identified as 5.
DETAILED DESCRIPTION OF THE INVENTION
Figure 1 is a schematic representation of an exemplary gas turbine engine
10. Engine 10 includes a fan assembly 12, a core engine 13, a high-pressure
compressor 14, and a combustor 16. Engine 10 also includes a high-pressure
turbine
18, a low-pressure turbine 20, and a booster 22. Fan assembly 12 includes an
array of
fan blades 24 extending radially outward from a rotor disc 26. Engine 10 has
an
intake side 27 through which air flows into and an exhaust side 29 through
which air
flows out of the engine. In one embodiment, the gas turbine engine is a GE90-
115B
that is available from General Electric Company, Cincinnati, Ohio. Fan
assembly 12
and turbine 20 are coupled by shaft 31. Compressor 14 and turbine 18 are
coupled by
shaft 33.
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During operation, air flows axially through fan assembly 12 in a direction
that is substantially parallel to central axis 34 extending through engine 10.
Compressed air is supplied primarily to combustor 16 by high-pressure
compressor
14. Most of the highly compressed air is delivered to combustor 16. Airflow
(not
shown in Figure 1) from combustor 16 drives turbines 18 and 20, and turbine 20
drives fan assembly 12 through shaft 31. High-pressure turbine 18 includes an
array
of blades 60.
Blade or airfoil 60 is shown in greater detail in Figure 3. Additionally the
airfoil may be a vane. Airfoil 60 comprises leading edge 74, and a trailing
edge 76
opposite the leading edge. The blade also comprises radially opposed blade tip
81
and root 79. The tip and root are separated by a radially extending distance.
The
blade is coupled with the rotor (not shown) at the root. Air flowing through
the gas
turbine engine along the flow path flows across the blade 60 in an axial
direction from
the leading edge 74 to the trailing edge 76. Compressed cooling air flows into
the
blade through openings at the leading edge 74 of the airfoil and also through
inlet
passages 77. The cooling air that flows through passages 77 flows radially
outward
toward blade tip 81. As the inlet passages extend toward tip 81, they combine
into a
single cooling passage 91. The cooling passage extends in a serpentine manner
through the interior of the blade. As shown in Figure 3, blade 60 includes two
inlets
but it should be understood that blade 60 may include any suitable number of
inlet
passages 77. Arrows in Figure 3 generally represent the flow direction of
cooling air
within blade 60.
A plurality of spaced apart vanes 92 are located in cooling passage 91
between inlet passages 77 and tip 81. The vanes are oriented in a parallel
array, with
each vane being substantially parallel to the other vanes in the array. Each
vane has a
first end 94 and a second end 95. For each vane the first end 94 of each
discrete vane
is located closer to root 79 than second end 95 of the same vane. For each
discrete
vane each second vane end 95 is located closer to tip 81 than first vane end
94 for the
same vane. The vanes are fixed to the wall that defines the portion of cooling
passage
91 at the trailing blade edge. The vanes are oriented at an angle relative to
generally
axially extending axis 99. Each vane is oriented relative to axis 99 at an
angle that is
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less than ninety degrees. By orienting the vanes in this manner, with the
first and
second ends for each vane at different radial locations, cooling air is more
effectively
directed into cooling slots 45.
As shown in Figure 3, blade 60 includes a plurality of cooling slots 45.
The cooling slots are oriented in a generally parallel array. For purposes of
disclosing
a preferred embodiment of the invention blade 60 comprises seven slots however
it
should be understood that any suitable number of slots 45 may be provided in
the
blade. Each slot has an inlet 96 and an outlet 97. The outlets 97 are located
at the
trailing edge 76 of blade 60. The slots are formed in the blade proximate the
trailing
edge. The inlet is in flow communication with the cooling slot 91 and cooling
air in
the cooling passage 91 enters the cooling slot through inlet 96. The slots 45
of blade
60 are of substantially constant radial dimension and the radial dimension may
be a
diameter for example. For each cooling slot, the outlet 97 is located closer
to the root
79 than the slot inlet 96 for the same cooling slot. For each discrete slot,
the slot inlet
96 is located nearer the blade tip 81 than the slot outlet 97 for the same
cooling slot.
As a result of positioning the inlet and outlet for each cooling slot at a
different radial
locations along the blade, the airfoil of the present invention more
effectively
produces a cooling film along the blade. More specifically, airfoil 60 more
effectively
forms a cooling film along the trailing edge 76 of the blade.
Figure 4 discloses an alternate embodiment blade 61 that comprises slots
48, similar to slots 45.. Slots 48 include inlet 106 and outlet 107. Like
slots 45, the
inlet and outlet for each slot is located at a different radial location along
the blade
with each inlet 106 located closer to tip 81 than outlet 107. The outlet 107
is located
closer to root 79 than inlet 106. The radial dimensions for inlets 106 and 107
are not
the same. As shown in Figure 4, the inlet has a smaller radial dimension than
the
outlet. The radial dimension may be a diameter for example with the diameter
of inlet
106 being smaller than the diameter of outlet 107. Blade 61 includes passages
77, 91
leading edge 74, trailing edge 76, tip 81, root 79 and vanes as described in
blade 60.
Note that unless specifically indicated to the contrary, as the description
proceeds the description relating to slot 45 shall also apply to slot 48. For
simplicity,
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the description shall refer to slot 45. As is shown in Figures 3 and 4,
substantially all
of the cooling slots 45, 48 may be oriented in a parallel array, at
substantially the
same angle Alpha (a) as shown in detail in Figure 5. The angle alpha,
identified at
110 is measured between reference line 35 and the central axis of slot 45. The
central
axis is identified as 120. The reference line 35 is substantially horizontal.
In an
alternate embodiment, fewer than substantially all of the slots may be
arranged in
parallel. For example, fifty percent of the slots may be arranged in parallel
at the same
angle 1 l0. Angle 110 of cooling slot 45 is shown in which the angle is less
than 90
and greater than 0 .
In practice, the flow of air through the cooling slot 45 of the present
embodiment invention is distinguishable from the flow of air through
conventional
slots where the slot inlet and outlet are located at the same radial positions
along the
length of the blade. Cooling slots 45 minimize the mass flow of air through
the slots
45 thus providing a controlled flow through the blade that is discharged from
the slot
outlet 97 at a velocity that is greatly reduced relative to prior art cooling
slots. Such
metered or controlled airflow creates a partial restriction of cooling air
passing
through the cooling slots 45. It should be understood that such restriction
does not
diminish the quality of the cooling layer formed on blade 60. Rather, the
controlled,
metered flow serves to enhance the formation of cooling film layer 30 and also
to
prevent both the escape of cooling air into the flow path of combustion gases
and the
formation of a backflow condition. By decreasing the cooling air mass flow
through
cooling slot 45 the velocity of the cooling air exiting the slots is reduced,
thereby
providing a cooler, slower moving boundary layer. As a result, upon exiting
the slot
the cooling air remains close to the surface and edges of turbine blade 60,
ensuring
that a suitable cooling layer is formed.
Figure 5 provides a more detailed view of cooling airflow entering,
traveling through and exiting cooling slot 45. Although the flow of cooling
air
entering, flowing through and exiting is only shown relative to one slot 45,
the flow
represents the flow for all slots 45 and 48. Cooling air flows to slot 45
through
passage 91, from a first flow position 126 toward cooling slot inlet 96.
Oppositely,
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cooling air flows through passage 91, in from second flow position 127 toward
cooling slot inlet 96. First flow position cooling air enters the blade
through openings
at the blade leading edge 74 and passes through upstream portion of passage 91
toward the slots. As cooling air flows to the cooling slot from flow position
126 it
may substantially move unobstructed into cooling slot 45. As cooling air
enters from
second flow position 127 the flow may be obstructed by one or more separation
regions 136 created at or proximate cooling slot inlet 96. A separation region
136
occurs in a region adjacent cooling slot inlet 96. When cooling air from the
flow
position 127 approaches slot 45, cooling air from flow position 127 abruptly
meets the
flow 126, and thus creates one or more areas in which the air swirls or
separates from
its original flow stream, producing separation region 136.
In addition to the angled orientation of cooling slot 45, the separation
region 136 can aid in metering the flow of cooling air through cooling slot 45
since it
can at least partially block the flow of air from flow position 127 from
moving into
cooling slot 45. This prevents the formation of backflow as well as
controlling the
flow of cooling air into the slot. Cooling film layer 130 is formed by the
cooling air
exiting from cooling slot outlet 45. Cooling film layer 130 is formed on the
leading
edge 76 of blade 60 and serves to help cool the surface of turbine blade 60
and protect
the blade against the harmful effects associated with hot combustion gases.
Cooling slot 45 is oriented at an angle 110 that may range from about I
degree (1 ) to about 88 degrees (88 ). In another embodiment the angle 110 may
range from about 10 degrees (10 ) to about 75 degrees (75 ). In still another
embodiment the angle may range from about 20 degrees (20 ) to about 60 degrees
(60 )(30 ) to about 50 degrees (50 ).
The pressure ratio for each turbine blade 60 at the inlet 96 of each cooling
slot 45 ranges from a pressure ratio of about 1.05 to about 2Ø The term
"pressure
ratio" means the ratio of the internal blade pressure to the external flow
path pressure.
It is desired to produce a pressure ratio greater than 1.0 since a pressure
ratio lower
than that would produce a backflow condition. Also, the movement of air within
the
airfoil through the cooling passage, slots and vanes is desired to have a Mach
number
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ranging from about 0.03 Mach number to about 1.0 Mach number. The Mach number
is defined as a ratio of the speed of an object or flow relative to the speed
of sound in
the medium through which it is traveling. In the present invention the Mach
number
falls into the desired range.
Additional benefits associated with the blade of the present invention
include the fact that more cooling slots 45 can be used in engines having
smaller
turbine blades. By the term "smaller turbine blades" it is meant herein a
turbine blade
in an aircraft engine application in which the engine core flow rate is less
than 13.61
kg/s at take-off power level. An exemplary engine having smaller turbine
blades of
the type discussed is a CT7 or T700 available from General Electric Company,
Cincinnati, Ohio.
The blade of the present invention allows cooling slots 45 to be cast rather
than drilled. The use of cast slots instead of drilled holes presents a
significant cost
savings in manufacturing, use of resources and material usage. In one
embodiment, at
least a portion of cooling slots 45 may be cast along trailing edge 76 of
turbine blade
60.
Cooling slots 45 of the invention also allow for beneficial variability. The
term "beneficial variability" means that one or more cooling slots 45 may have
a
varying diameter along its length and/or because of casting may have much
larger
diameters in comparison to drilled cooling slots 75. One example of beneficial
variability is the use of larger holes, i.e., the exits of the cooling slots
along the
trailing edge of the turbine blades 70 (see Figure 4). By having larger exit
holes than
those provided by drilling, e.g., laser drilling, greater cooling film
coverage is
achieved about the surface of turbine blade 60. Also, since outlets 107 can be
made to
be larger, than current slot technology, fewer cooling slots 45 may be used
than in
blades where constant radial dimension/diameter slots are used.
This written description uses examples to disclose the invention, including
the best mode, and also to enable any person skilled in the art to make and
use the
invention. The patentable scope of the invention is defined by the claims, and
may
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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 language
of the
claims.
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