Gas turbine engine cooling system

A gas turbine engine cooling system. A first turbocompressor and a heat exchanger are fluidly interconnected and are each in fluid communication to receive air of differing pressures and temperatures. Typically, such air is received from various regions of the engine low pressure compressor and the engine high pressure compressor. The system delivers air through a duct to a portion of the engine for cooling, such as the engine high pressure turbine region, at lower temperatures and higher pressures than if cooling air were directly ducted from the engine compressor to the engine turbine.

BACKGROUND OF THE INVENTION 
The present invention relates generally to gas turbine engines, and more 
particularly to a system for cooling such an engine. 
Gas turbine engines (such as turbojet engines, bypass turbofan engines, 
turboprop engines, turboshaft engines, etc.) may be used to power flight 
vehicles (such as planes, helicopters, and missiles, etc.) and may also be 
used to power ships, tanks, electric power generators, pipeline pumping 
apparatus, etc. For purposes of illustration, the invention will be 
described with respect to an aircraft bypass turbofan gas turbine engine. 
However, it is understood that the invention is equally applicable to 
other types and/or uses of gas turbine engines. 
A gas turbine engine includes a core engine having, in serial flow 
relationship, high pressure compressor (also called a core compressor) to 
compress the airflow entering the core engine, a combustor in which a 
mixture of fuel and the compressed air is burned to generate a propulsive 
gas flow, and a high pressure turbine which is rotated by the propulsive 
gas flow and which is connected by a larger diameter shaft to drive the 
high pressure compressor. A typical aircraft bypass turbofan gas turbine 
engine adds a low pressure turbine (located aft of the high pressure 
turbine) which is connected by a smaller diameter coaxial shaft to drive a 
front fan (located forward of the high pressure compressor) which is 
surrounded by a nacelle and which may also drive a low pressure compressor 
(located between the front fan and the high pressure compressor). The low 
pressure compressor sometimes is called a booster compressor or simply a 
booster. It is understood that the term "compressor" includes, without 
limitation, high pressure compressors and low pressure compressors. A flow 
splitter, located between the fan and the first (usually the low pressure) 
compressor, separates the air which exits the fan into a core engine 
airflow and a surrounding bypass airflow. The bypass airflow from the fan 
exits the fan bypass duct to provide most of the engine thrust for the 
aircraft. Some of the engine thrust comes from the core engine airflow 
after it flows through the low and high pressure compressors to the 
combustor and is expanded through the high and low pressure turbines and 
accelerated out of the exhaust nozzle. 
Aircraft bypass turbofan gas turbine engines are designed to operate at 
high temperatures to maximize engine thrust. Cooling of engine hot section 
components (such as the combustor, the high pressure turbine, the low 
pressure turbine, and the like) is necessary because of the thermal 
"redline" limitations of the materials used in the construction of such 
components. Typically such cooling of a portion of the engine is 
accomplished by ducting (also called "bleeding") cooler air from the high 
and/or low pressure compressors to those engine components which require 
such cooling. Unfortunately the relatively low pressure and hot 
temperature of the compressor air limits its ability to be used to cool 
such engine components. 
SUMMARY OF THE INVENTION 
It is an object of the invention to provide a system for improved cooling 
of the hot section components and other portions of a gas turbine engine. 
In a first embodiment, the invention provides a system for cooling a 
portion of a gas turbine engine and includes a turbocompressor and a first 
heat exchanger. The first heat exchanger has an inlet and an outlet for a 
first fluid flow providing cooling to the first heat exchanger and has an 
inlet and an outlet for a second airflow receiving cooling from the first 
heat exchanger. The first fluid flow inlet is in fluid communication with 
lower temperature fluid (e.g., lower pressure and temperature discharge 
air from the booster compressor section of the engine compressor). The 
second airflow inlet is in fluid communication with higher temperature air 
from the engine compressor (e.g., higher pressure and temperature 
discharge air from the high pressure compressor section of the engine 
compressor). The second airflow outlet is in fluid communication with the 
inlet of the compressor section of the turbocompressor. The outlet of the 
compressor section of the turbocompressor is in fluid communication with 
the portion of the engine requiring the cooling. 
In a similar second embodiment of the invention, the first airflow inlet is 
in fluid communication with lower temperature fluid (e.g., lower pressure 
and temperature discharge air from the booster compressor section of the 
engine compressor). The second airflow inlet is in fluid communication 
with the outlet of the compressor section of the turbocompressor. The 
inlet of the compressor section of the turbocompressor is in fluid 
communication with higher temperature air from the engine compressor 
(e.g., higher pressure and temperature discharge air from the high 
pressure compressor section of the engine compressor). The second airflow 
outlet is in fluid communication with the portion of the engine requiring 
the cooling. 
Several benefits and advantages are derived from the gas turbine engine 
cooling system of the invention. Use of the turbocompressor and heat 
exchanger of the invention permit higher pressure and lower temperature 
air to be used for cooling portions of the engine such as engine hot 
section components. Maximum engine thrust, which is especially important 
during takeoff and climb, can be increased for a particular redline, 
temperature limit for the engine hot section components by using the 
higher pressure and lower temperature cooling air of the invention, as can 
be appreciated by those skilled in the art.

DETAILED DESCRIPTION OF THE INVENTION 
Referring now to FIG. 1, there is illustrated an aircraft bypass turbofan 
gas turbine engine 10 having a generally longitudinally extending axis or 
centerline 12 generally extending forward 14 and aft 16. The bypass 
turbofan engine 10 includes a core engine (also called a gas generator) 18 
which comprises a high pressure compressor or core compressor 20, a 
combustor 22, and a high pressure turbine 24, all arranged in a serial, 
axial flow relationship. A larger diameter annular drive shaft 26, 
disposed coaxially about the centerline 12 of the engine 10, fixedly 
interconnects the high pressure compressor 20 and the high pressure 
turbine 24. 
The core engine 18 is effective for generating combustion gases. 
Pressurized air from the high pressure compressor 20 is mixed with fuel in 
the combustor 22 and ignited, thereby generating combustion gases. Some 
work is extracted from these gases by the high pressure turbine 24 which 
drives the high pressure compressor 20. The combustion gases are 
discharged from the core engine 18 into a low pressure or power turbine 
28. The low pressure turbine 28 is fixedly attached to a smaller diameter 
annular drive shaft 30 which is disposed coaxially about the centerline 12 
of the engine 10 within the larger diameter annular drive shaft 26. The 
smaller diameter annular drive shaft 30 rotates a forward row of fan rotor 
blades 32. The smaller diameter annular drive shaft 30 also rotates a low 
pressure compressor 34 (also called a booster compressor or simply a 
booster). A flow splitter 36, located between the fan blades 32 and the 
low pressure compressor 34, separates the air which exits the fan into a 
core engine airflow which exits the exhaust nozzle (not shown) and a 
surrounding bypass airflow which exits the fan bypass duct 38. 
FIG. 1 shows a first application of the engine cooling system 110 of the 
invention used for cooling a first portion of the engine 10, wherein the 
first portion comprises the high pressure turbine 24. The cooling system 
110 receives air: from a duct 112 which bleeds air from the low pressure 
compressor discharge region 114; from a duct 116 which bleeds air from the 
high pressure compressor discharge region 118; and from a duct 120 which 
bleeds air from a region 122 in between such low and high pressure 
compressor discharge regions 114 and 118. The cooling system 110 
discharges air: to a duct 124 which routes air to the fan bypass duct 38; 
and to a duct 126 which routes air to the high pressure turbine 24 region. 
FIG. 2 shows a first embodiment 110' of the engine cooling system 110 
comprising a turbocompressor 128 and a first heat exchanger 130. The 
turbocompressor 128 has a compressor section 132 including an inlet 134 
and an outlet 136 and has a turbine section 138 including an inlet 140 and 
an outlet 142. Preferably the turbocompressor 128 has air bearings. The 
first heat exchanger 130 has an inlet 144 and an outlet 146 for a first 
airflow providing cooling to the first heat exchanger 130 and has an inlet 
148 and an outlet 150 for a second airflow receiving cooling from the 
first heat exchanger 130. The first airflow inlet 144 of the first heat 
exchanger 130 is in fluid communication with lower pressure and 
temperature air (such as with a portion of the air from the low pressure 
compressor discharge region 114 through duct 112/112a as shown in FIGS. 1 
and 2). The second airflow inlet 148 of the first heat exchanger 130 is in 
fluid communication with higher pressure and temperature air from the 
engine compressor (such as with a portion of the air from the high 
pressure compressor discharge region 118 through duct 116 as shown in 
FIGS. 1 and 2). The first airflow outlet 146 of the first heat exchanger 
130 preferably is in fluid communication with the fan bypass duct 38 
through duct 124a (and preferably discharges such air into the fan bypass 
duct 38 with an aft component of velocity). The second airflow outlet 150 
of the first heat exchanger 130 is in fluid communication with the inlet 
134 of the compressor section 132 of the turbocompressor 128 through duct 
152. The outlet 136 of the compressor section 132 of the turbocompressor 
128 is in fluid communication with the high pressure turbine 24 through 
duct 126 to cool at least a portion of the high pressure turbine 24. The 
inlet 140 of the turbine section 138 of the turbocompressor 128 is in 
fluid communication with intermediate pressure and temperature air from 
the engine compressor (such as with a portion of the air from the eighth 
stage high pressure compressor region 122 through duct 120 as shown in 
FIGS. 1 and 2). As can be appreciated by those skilled in the art, the 
higher pressure and temperature air has a higher pressure and temperature 
than that of the lower pressure and temperature air, and the intermediate 
pressure and temperature air has a pressure and temperature intermediate 
that of the lower pressure and temperature air and the higher pressure and 
temperature air. It is understood that the term "pressure" means total 
pressure (i.e., static pressure plus dynamic pressure). The outlet 142 of 
the turbine section 138 of the turbocompressor 128 preferably is in fluid 
communication with the fan bypass duct 38 through duct 124b (and 
preferably discharges such air into the fan bypass duct 38 with an aft 
component of velocity). 
In an alternate embodiment (not shown), the turbine section 138 of the 
turbocompressor 128 has its inlet 140 in fluid communication with the 
second airflow outlet 150 of the first heat exchanger 130 instead of being 
in fluid communication with the intermediate pressure and temperature air 
region 122. This embodiment drives the turbine section 138 of the 
turbocompressor 128 with higher pressure air. Preferably, in this 
embodiment, the outlet 142 of the turbine section 138 of the 
turbocompressor 128 is in fluid communication with the engine low pressure 
turbine instead of being in fluid communication (via duct 124b) with the 
fan bypass duct 38. This embodiment cools both the high and low pressure 
turbines of the engine with the use of a single heat exchanger. 
FIG. 3 shows a second embodiment 110" of the engine cooling system 110 
which is identical to the first embodiment 110' of FIG. 2 previously 
discussed, but with three differences. First, the second airflow outlet 
150 of the first heat exchanger 130 is in fluid communication with the 
high pressure turbine 24 through duct 126 to cool at least a portion of 
the high pressure turbine 24. Second, the inlet 134 of the compressor 
section 132 of the turbocompressor 128 is in fluid communication with 
higher pressure and temperature air from the engine compressor (such as 
with a portion of the air from the high pressure compressor discharge 
region 118 through duct 116 as shown in FIGS. 1 and 3). Third, the outlet 
136 of the compressor section 132 of the turbocompressor 128 is in fluid 
communication with the second airflow inlet 148 of the first heat 
exchanger 130. 
FIG. 4 shows a second application of the engine cooling system 210 of the 
invention used for cooling a first portion of the engine 10, wherein the 
first portion comprises the high pressure turbine 24 and also for cooling 
a second portion of the engine 10, wherein the second portion comprises 
the low pressure turbine 28. The cooling system 210 as shown in FIG. 4 is 
identical to the first application's cooling system 110, as shown in FIG. 
1, but with one addition. The cooling system 210 also discharges air to a 
duct 154 which routes air to the low pressure turbine 28 region. 
FIG. 5 shows a first embodiment 210' of the engine cooling system 210 which 
is identical to the first embodiment 110' of the engine cooling system 110 
of FIG. 2 previously discussed, but with one addition and one difference. 
Briefly, the addition is a second heat exchanger 156, and the difference 
is in the duct which is in fluid communication with the outlet 142 of the 
turbine section 138 of the turbocompressor 128. More particularly, the 
second heat exchanger 156 has an inlet 158 and an outlet 160 for a third 
airflow providing cooling to the second heat exchanger 156 and has an 
inlet 162 and an outlet 164 for a fourth airflow receiving cooling from 
the second heat exchanger 156. The first airflow inlet 158 of the second 
heat exchanger 156 is in fluid communication with lower pressure and 
temperature air (such as with a portion of the air from the low pressure 
compressor discharge region 114 through duct 112/112b as shown in FIGS. 4 
and 5). The second airflow inlet 162 of the second heat exchanger 156 is 
in fluid communication with the outlet 142 of the turbine section 138 of 
the turbocompressor 128 through duct 166. The first airflow outlet 160 of 
the second heat exchanger 156 preferably is in fluid communication with 
the fan bypass duct 38 through duct 124c (and preferably discharges such 
air into the fan bypass duct 38 with an aft component of velocity). The 
second airflow outlet 164 of the second heat exchanger 156 is in fluid 
communication with the low pressure turbine 28 through duct 154 to cool at 
least a portion of the low pressure turbine 28. 
FIG. 6 shows a second embodiment 210" of the engine cooling system 210 
which is identical to the second embodiment 110" of FIG. 3, but with the 
one addition (the second heat exchanger 156) and the one difference (the 
duct which provides fluid communication from the outlet 142 of the turbine 
section 138 of the turbocompressor 128) as previously discussed. 
Conventional engine cooling techniques duct compressor air directly to the 
high pressure turbine 24 region and the low pressure turbine 28 region of 
the engine 10. The cooling system of the invention can be used to augment 
such conventional engine cooling techniques or it can be used to 
substitute for such conventional techniques. 
The operation of the first embodiment 210' of the engine cooling system 210 
is typical of the other embodiments, and will be described with reference 
to a numerical example based on engineering analysis where the pressure P 
is measured in psia and the temperature T is measured in degrees R. 
Referring to FIGS. 4 and 5, it is seen that air (P=34.8, T=810) from the 
low pressure compressor discharge region 114 is carried by duct 112a to 
the first heat exchanger 130 to cool air (P=497, T=1689) from the high 
pressure compressor discharge region 118 carried by duct 116 entering the 
first heat exchanger 130 so that the airflow receiving cooling will exit 
the first heat exchanger 130 as air (P=462, T=1369) carried by duct 152 to 
the compressor section 132 of the turbocompressor 128. The compressor 
section 132 is driven by the turbine section 138 from air (P=277, T=1486) 
from the intermediate compressor region 122 carried by duct 120. Air 
(P=497, T=1416) leaves the compressor section 132 of the turbocompressor 
128 in duct 126 to cool the high pressure turbine 24. (Conventional 
cooling of the high pressure turbine directly with air from the high 
pressure compressor discharge region would deliver such air at P=464 and 
T=1647.) In a similar fashion, air (P=140, T=1109) leaves the second heat 
exchanger 156 in duct 154 to cool the low pressure turbine 28. 
(Conventional cooling of the low pressure turbine directly with air from 
the intermediate compressor region would deliver such air at P=130 and 
T=1186.) The greater engine cooling capacity of the system of the 
invention achieves a net thrust of 51878 pounds compared with 45139 pounds 
of net thrust using conventional cooling (with the high temperature 
turbine blade temperature at a "redline" limit of 1838 for both the 
cooling system of the invention and the conventional cooling). The 
improvement in net thrust for the engine cooling system of the invention 
is nearly fifteen percent. 
The foregoing description of a preferred embodiment of the invention has 
been presented for purposes of illustration. It is not intended to be 
exhaustive or to limit the invention to the precise form disclosed, and 
obviously many modifications and variations are possible in light of the 
above teachings. For example, it is understood that the phrase "engine 
compressor" includes any low, intermediate, and/or high pressure engine 
compressor. Also, various portions of the engine which may be cooled by 
the engine cooling system of the invention include those portions 
involving high pressure turbine cooling, low pressure turbine cooling, 
combustor cooling, compressor disc cooling, compressor discharge cooling, 
compressor and turbine case cooling, clearance control cooling, etc. 
Additionally, lower pressure and temperature air may be bled or ducted 
from the fan region, the fan bypass region, etc. as well as from the low 
pressure compressor region. Further, the invention is applicable to gas 
turbine engines having axial, radial, or other types of gas turbine engine 
compressors and/or turbines. It is likewise understood that in some 
applications, the engine cooling system of the invention may employ valves 
to control the airflow in the various ducts, and/or the engine cooling 
system of the invention may be employed in those engines having variable 
turbine nozzles where greater cooling is required when the turbine nozzle 
area is reduced. It is noted that the first airflow can be generalized as 
a first fluid flow having a higher temperature than that of the second 
airflow, and that such first fluid flow can be engine fuel wherein, for 
example, duct 112 would convey some fuel from the fuel tank to the heat 
exchanger(s) and duct 124 would convey the fuel from the heat exchanger(s) 
back to the fuel tank or to the combustor, etc. (such arrangement not 
shown in the drawings). Such modifications and variations, and other 
modifications and variations, are all within the scope of the claims 
appended hereto.