Note: Descriptions are shown in the official language in which they were submitted.
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COOL TIP COMBUSTOR
~ACKGROVND OF THE INVENTION
1. Field of the In~ention
This invention relates to means for directing
cooling air to critical parts of hot section turbine
blades in gas turbine engines.
2. Description of the Prior Art
In the course of gas turbine engine development,
tremendous effort has been directed atraising ~he
internal operating temperatures of such engines to improve
thermodynamic efficiency. As turbine inlet temperatures
have been increased in pursuit of this -goal, it has
become necessary to provide cooling air to hot section
turbine blades and vanes in order to limi-t temperatures
of those components to levels that can be accommodated
by the blade and vane materials. The air that is used
for this cooling function is usually compressed to
pressures that meet or exceed the gas pressures inside
the turbine section. Because the air has undergone
the work necessary for compression, this cooling air must
be used as efficiently as possible to limit the power
required by the engine's compressor section in order to
compress that air. To limit the amount of cooling air
used, intricate cooling air flowpaths and passages
are utilized that are intended to use the cooling air
in a highly efficient manner.
In smaller aixflow size engines, blade cooling
configurations are generally restricted to fairly
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simple designs because of small dimensions and limitations
of current manufacturing technologies. The implication
is a typical smaller engine turbine blade or vane cannot
be provided with the highl~ complex, internal air
cooling passage configuration typically used today in
larger gas turbine engines.
One particular problem with smaller engines is that
tip sections of turbine blades are extremely difficult
to cool efficiently. The cooling air used to internally
cool turhine blade tips has increased its tempexature
by thermal pickup in ~he lowex portion o~ the blade
rendering it less effective for cooling purposes.
In downstream sections of the turbine blade tips some
of the cooling air has been bled out of the trailing
edge cooling holes before it reaches the blade tip
region, thereby reducing cooling air velocity and,
consequently, its cooling effectiveness. Adding to these
difficulties of cooling small turbine blades, the down-
stream trail;~ng edge of the blade tip region is usually
very thin for aerodynamic performance reasons, which
limits the ability to duct cooling air into this region.
As a result of these inherent limitations, design
cycle temperatures of these small engines are restricted
and engine performance is thereby limited. Further,
~5 the turbine blade tips often become a life-limiting
engine component problem area. As the turbine tips
deteriorate, due to Gxidation and corrosion accumulating
during engine use, the engine performance drops below
minimum acceptable levels. The engine must then be
removed from the aircraft and the turbine section
refurbished. Maintenance and overhaul of the kurbine
section to correct deteriorated blade tips is both
expensive and time consuming.
It is, therefore, an object of the present invention
to provide a means for cooling tips of turbine blades
in turbine sections of gas turbine engines with a system
:,', ~ ' ' .
~4L8~
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that can be utilized in relatively small engine con-
figurations.
Another object of the present invention is to
provide a source of cooling air that can be directed
specifically to turbine blade tips in a turbine section
of a small gas turbine engine.
It is another object of the present invention to
provide a film of cooling air along a radially outer
most wall of a turbine section of a small gas turbine
engine for the purpose of cooling turbine blade tips
with a limited amount of cooling air.
These and other objects will become more readily
apparent upon reference to the following description
in conjunct:ion with the appended drawings.
SUMMARY OF THE INVENTION
-
Briefly, in accordance with one embodiment of the
present invention, means are provided for introducing
cooling air into a turbine section of a gas turbine
engine in the region of tip sections of turbine blades.
The source of this cooling air is compressor discharge
air that has bypassed the combustor. This compressor
discharge air is introduced at an aft section of the
combustor through inlet air holes just upstream of the
turbine section. The air is introduced along a radially
outer section of the combustor only. The cooling air
flows initially into annulus regions within the
combustor that are protected from hot combustion gases.
From these annulus regions, the air flows downstream
in the combustor and forms a thick film that blankets
the combustor wall. Because it is introduced at the
downstream section of the combustor, there is no
combustion of this cooling air, and i-t enters the
turbine section at close to the same temperature as
it enters the turbine section at close to the same
temperature as when the cooling air entered the
combustor section. This temperature is must lower than
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the hot gases which have just undergone combustion.
This thick,low-temperature cooling air film flows into
the turbine section along a radially outer wall of the
turbine flow path. The cooling air film provides a
relatively cooler gas flow along the tips of the turbine
blades rotating in the turbine section. It ls primarily
the tips of the turbine blades only that are thus cooled,
and this limits the amount of cooling air employed.
A BRIEF DESCRIPTION OF THE D~AWIN~S
Figure 1 is a schematic cross-sectional illustration
of a central section o~ a gas turbine engine.
Figure 2 is a schematic cross-sectional illustration
of a combustor and high pressure turbine section of a gas
turbine engine with the present invention embodied therein.
Figure 3 is a cross-sectional illustration of a
downstream portion of a combustor wall with one embodiment
of part of the present invention incorporated therein.
Figure 4 i5 a graphical representation of test
results of turbine blade temperatures.
DESCRIPTION OF THE PREFERRED EMBODIMENT
Referring now to Figure 1, a central section of a
typical gas turbine engine 10 is shown -that involves
substantial turhomachinery that rota~es about an engine
centerline 12. Components of this turbomachinery include,
in serial flow relationship, a compressor 14, a combustor
16, a high-pressure turbine section 18, and a low-
pressure turbine section 20. In conventional operation,
inlet air is directed into and pressurized by the
compressor 14 from which the air is discharged through
a diffuser 22. A major portion of this compressor
discharge air is then passed into the combustor 16
where it is mixed with fuel and vaporized to form high-
pressure, high-temperature combustion gases which flow
downstream into the high-pressure turbine 18. The high-
pressure gases cause turbine blades 24 in the high-
pressure turbine 18 -to rotate at high velocities thereby
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providing mechanical power. These high-temperature, high-
pressure gases then continue to flow downsteam into the low-
pressure tu.rbine 20 where they cause low-pressure turbine
blades 26 to rotate thereby providing additional mechanical
power. From the low-pressure turbine 20, the gases are
discharged downstream so as to pass out of the engine 10.
A portion of the air discharge from the compressor 14
that passes through the diffuser 22 is circulated to cool
a vari.ety of hot parts of the engine 10. Some of that air
used for cooling flows to the region of the combustor 16 and
surrounds the combustor walls. In some engines, small
cooling holes are provided in combustor walls so that cooling
air can enter the combustor to cool interior combustor
surfaces. Other portions of the cooling air are directed
internally to hot temperature parts inside the high-pressure
turbine 18. A part of this air used to cool the high-pressure
turbine is directed into the interior of a high-pressure
turbine nozzle 28 so as to provide an internal cooling function
by impingement and diffusion processes. Another part of the
compressor discharge air is directed along other paths to
cool interior regions of the turbine blades 24 of the high-
pressure turbine 18. These cooling flowpaths are generally
represented by the dark arrows in Figure 1~
~t is well-known in the art that during high-power,
high-temperature operating conditions a substantial amount
of cooling air is needed for these cooling processes.
Because of the limitations of size and manufacturing processes,
it is particularly difficult to cool tip sections 30 of the
turbine blades 24. These tip sections 30 are usually very
thin for aerodynamic performance reasons and this limits the
ability to efficiently duct cooling air into the tip sections.
In addition, the thin sections deteriorate due to oxidation
and corrosion causing substantial problems in engine per-
formance.
A prior art solution to the problem of cooling
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turbine blade tips is a ducting of a small portion of the
compressor discharge cooling air into the high-pressure
turbine 18 at an inlet location 32 just downstream of
the turbine nozzle 28. Cooling air ducted in this manner
would bypass the combustor and flow into the high-pressure
turbine 18 just upstream of the turbine blades 24.
Studies have indicated that this proposal reduces turbine
tip temperatures, but this approach also has a negative
impact on engine performance, both in terms of thrust
10 and fuel consumption. The de ~imental effect on engine
performance is caused because the cooling air enters the
gas flow stream behind the first-stage turbine nozzle
2~ and is, ~herefore, chargeable to the engin's thermody-
namic cycle. As a result, the amount of air burned
15 compared to the allowable turbine rotor inlet temperature
level is reduced and engine performance decreases.
Referring now to Fig 2, a portion of a gas turbine
engine 11 is shown that is generally similar to part of
the engine shown in Figure 1 but, this time, incorporating
20 an embodiment of the present invention. Again, as exp~ ined
in relation to the engine shown in Figure 1, a portion of
the cooling air discharged from the compressor does not
enter the cc~ustor 16, but, instead, flows downstream
around the combustor as indicated by dark arrows in
25 Figure 2. This cooling air does not undergo the mixing
and combustion processes occurring during engine operation
inside the combustor 16. Because the air does not
undergo combustion, it remains relatively cool and serves
as a source of high-pressure cooling air that can be
3~ utilized in the high-pressure turbine sections of the
engine. Any cooling air used in the high-pressure
turbine section must be highly pressure because the
internal gases flowing through the high-pressure turbine
area, as the name suggests, are at very high pressure.
8~
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The cooling air introduced into the high-pressure turbine
must be even higher in pressure than -those gases flowing
through the turbine so that the cooling air will be caused
by its own pressure forces to flow into the -turbine blades
and vanes and ~rom there into the combustion gas Elow
passage of the turbine section. If the cooling air that
was used for cooling in this region were lower in pressure
than the combustion gases flowing through the turbine
section, pressure forces would not permit the cooling air
to flow from interior regions of the turbine blades and
vanes out into the combustion gas flow passage.
Realizing that this compressor discharge air is the
best available source of cooling air flow that can be utili~ed
for cooling turbine blades, the problem becomes a matter of
utilizing this air in the best manner possible to cool the
turbine blades and the turbine blade tips. It is extremely
important that the volume of cooling air used be ]cept as
low as possible because the air has undergone a great deal
of work in the compressor section in order to compress that
air, and it is desirable to minimize the amount of air
used in o.rder to increase the efficiency of the engine.
It is also desirable to introduce this highly compressed
cooling air in a location that permits highly pressurized
air to be expanded and directed at the turbine blades in a
manner such that the cooling air will not only cool the
tips of the turbine blades but will also add to the
effecti~e gas foxces that cause the turbine blades 2~ to
rotate, thereby increasing the total power produced by the
engine 10.
If cooling air is introduced at an inlet location 32
immediately downstream of the turbine nozzle 28, the air
will tend to cool the turbine blade tips 20. However,
because the air has not been expanded and directed by the
turbine nozzle 28, it will not be useful for providing
appropriate gas forces~or causing the turbine blades 24 to
rotate.
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The present invention comprises means for introducing
cooling air forward or upstream of the first-stage
turbine nozzle 28 so that there is no associated engine
performance penalty. One embodiment of this means is
shown in Figure 2, and a part of the invention is shown in
larger scale i~ Figure 3. Referring initially to Figure 2,
a portion of the compressor discharge air that is flowing
outside of the combustor 16 is directed into combustor
inlet air holes 36 at a location just upstream of the
turbine nozzle 2g. The air is introduced at a location
just upstream of the turbine nozzle 28, partly to prevent
that cooling air from undergoing the normal combustion
processes inside the turbine 16, and also to lessen
heating of the cooling air from prolonged exposure to the
hot combustion gases. If this cooling air were to undergo
combustion, it would climb dramatically in temperature
and be rendered relatively useless for the purpose of
cooling tips 30 of turbine bladesO
Referring now to Figure 3, the inlet air holes 36
through which the cooling air is directed into a downstream
section of the combustor 16 are shown in greater detail.
A portion of a radially Guter wall section 38 of the
combustor 16 is shown in Figure 3. This portion of the
combustor wall section 38 is located just upstream of the
turbine nozzle 28 (not shown). In the cross-sectional
view shown, three inlet air holes 36 can be seen and their
relative configuration can be appreciated. It should
first be noted that the downstream portion of the combustor
wall section 38 is actually double-walled. An outer wall
section 40 connects to the turbine nozzle in a standard
manner as would be the normal practice in many gas turbine
engines. An inner combustor wall section 42 is provided
and is protected from hot combustion gases at its upstream
end by a flange 44. At its downstream end, the inner wall
section 42 extends almost to the turbine nozzle inl~-t.
Cooling air from the compressor discharge is blend into
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annulus regions 46 that are open in a downstream direction
and are generally protected from the combustion occurring
inside the com~ustor 16. Because the cooling air is blend
into these protected annulus regions 46, the cooling air
does not undergo combustion, and the air enters the turbine
nozzle at substantially compressor discharge temperature,
thereby forming a thick, low-temperature film along a
radially outer wall of the turbine flow path.
As stated earlier, there are three inlet air holes 36
visible in Figure 3. Each of the holes 36, as shown,
represents cne of a row of holes tha-t extend around the
entire circumference of the radially outer wall section 38
of the combustor 16. The total number of inlet air holes 35
could vary widely as could their general configuration.
A row of upstream inlet air holes ~8 is provided to
bleed cooling air into the annulus region be-tween the
flange 4~ and the inner wall 42. A row of intermediate
inlet air holes 50 is provided to bleed additional cooling
air into the annulus region between the inner wall 42
and the outer wall ~0. Finally, a row of downstream inlet
air holes 52 is provided to direct additional coo]ing
air into the annulus between the inner wall 42 and the
outer wall 40. It can be readily appreciated by those
skilled in the air that the size of these inlet air holes
36 can be varied for the purpose of introducing varying
amounts of cooling air. To serve as a guide, in one
embodiment of the pxesent invention, these holes are varied
from .026 inches ~ oe6 cm.) in diameter to .035 inches
(.089 cm.~ in diameter. These dimensions, however, are
simply a guideline and smaller or larger diameter holes
could easily be utilized without departing from the
scope of the present invention. Additionally, widely
varying inlet air hole configurations would also be
~ithin the scope of the invention.
Referring again to Figure 2, small black arrows are
shown entering the combustor 16 emanating from the annulus
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regions 46 within the combustor 16 and flowing downstream
along the radially outer turbine wall 34, past the turbine
nozzle 2~ to the region of the turbine blade tips 30.
This air tends to flow as a low temperature film in a
manner that is ideal for cooling the turbine blade tips 30
without using excessive amounts of compressor discharge air
thereby accomplishing the purpose of the present invention.
Referring now to Figure 4, a comparison of test results
is shown that graphically represents turbine blade tem-
peratures in a typical gas turbine engine andl additionally,represents turbîne blade temperatures in a second gas
turbine engine incorporating the present invention. The
X(horizontal~ coordinate in Figure 4 is marked off in
degrees Fahrenheit. The Y(vertical) coordinate in Figure
4 is a dimensionless representa-tion of turbine blade height,
beginning at a root of the turbine blade and ending at the
tip of the turbine blade. The lines shown on the graph of
Figure 4 designated 54 represent turbine blade temperatures
in two typical gas turbine engines, generally having an
engine configuration similar to that shown in Figure 2
but without incorporating the present invention. The line
designated 56 in Figure 4 represents turbine blade tem-
perature, again within an engine having generally the same
configuration as shown in Figure 2, but this time
incorporating the present invention. It can be readily
appreciated that turbine tip temperatures are significantly
decreased in the engine incorporating the present invention.
Because of thi~ temperature reduction at the turbine tip,
the present invention has been commonly referred to as a
"cool tip" ~ngine. It is important to note that this
reduction in turbine tip temperatures is achieved
generally without utilizing excessive amounts of compressor
discharge air and in a manner that directs the cooling
effect at the turbine blade tip~. It i5 desirahle to
obtain this "cool tip" effect in a localized manner as
sho~n graphically in Figure 4.
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Although the present invention has been described in
terms of its preferred embodiment, it will be apparent to
those skilled in the art the changes and modifications
thereof may be made without departing from the scope of
the appended claims which define the present invention.
