U.S. patent application number 17/519887 was filed with the patent office on 2022-02-24 for external mixing chamber for a gas turbine engine with cooled turbine cooling air.
The applicant listed for this patent is Raytheon Technologies Corporation. Invention is credited to James D. Hill, Frederick M. Schwarz.
Application Number | 20220056847 17/519887 |
Document ID | / |
Family ID | |
Filed Date | 2022-02-24 |
United States Patent
Application |
20220056847 |
Kind Code |
A1 |
Hill; James D. ; et
al. |
February 24, 2022 |
EXTERNAL MIXING CHAMBER FOR A GAS TURBINE ENGINE WITH COOLED
TURBINE COOLING AIR
Abstract
A gas turbine engine comprises a compressor section and a
turbine section, the compressor section having a last compressor
stage. High pressure cooling air is tapped from a location
downstream of the last compressor stage and passed through a heat
exchanger. Lower pressure air passes across the heat exchanger to
cool the high pressure cooling air. A housing surrounds the
compressor section and the turbine section and there being a space
radially outwardly of the housing, and a mixing chamber received in
the space radially outwardly of the housing, the mixing chamber
receiving the high pressure cooling air downstream of the heat
exchanger, and further receiving air at a temperature higher than a
temperature of the high pressure cooling air downstream of the heat
exchanger. Mixed air from the mixing chamber is returned into the
housing and utilized to cool at least the turbine section.
Inventors: |
Hill; James D.; (W. Abington
Twp., PA) ; Schwarz; Frederick M.; (Glastonbury,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Raytheon Technologies Corporation |
Farmington |
CT |
US |
|
|
Appl. No.: |
17/519887 |
Filed: |
November 5, 2021 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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15424927 |
Feb 6, 2017 |
11215120 |
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17519887 |
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International
Class: |
F02C 7/18 20060101
F02C007/18; F02C 3/107 20060101 F02C003/107; F01D 25/12 20060101
F01D025/12; F02C 6/08 20060101 F02C006/08 |
Claims
1. A gas turbine engine comprising: a propulsor, a compressor
section and a turbine section, said compressor section having a
last compressor stage; high pressure cooling air tapped from a
location downstream of said last compressor stage and passed
through a heat exchanger; lower pressure air passing across said
heat exchanger to cool said high pressure cooling air; a housing
surrounding said compressor section and said turbine section and
there being a space radially outwardly of said housing, and a
mixing chamber received in said space radially outwardly of said
housing, said mixing chamber connected to receive said high
pressure cooling air at a location downstream of said heat
exchanger, and further connected to receive air at a temperature
higher than a temperature of said high pressure cooling air taken
downstream of said heat exchanger; mixed air from said mixing
chamber returned into said housing and utilized to cool at least
one of said turbine section or said last compressor stage; wherein
said high pressure cooling air is tapped from a diffuser section
radially outwardly of a combustor located intermediate said
compressor section and said turbine section; and wherein said mixed
air passes through a compressor diffuser downstream of said last
compressor stage.
2. The gas turbine engine as set forth in claim 1, wherein said
mixing chamber has diverters to cause a change in a flow direction
from said higher pressure cooling air and said air at a higher
temperature.
3. The gas turbine engine as set forth in claim 2, wherein said
diverters cause a change in a flow direction of the high pressure
cooling air and a flow direction of said air at a higher
temperature.
4. The gas turbine engine as set forth in claim 3, wherein said
diverters cause said high pressure cooling air to swirl and said
air at a higher temperature to swirl.
5. The gas turbine engine as set forth in claim 4, wherein said
compressor section includes a low pressure compressor and a high
pressure compressor and said turbine section including at least a
high pressure turbine and a propulsor drive turbine, and said
propulsor drive turbine driving said propulsor through a gear
reduction.
6. The gas turbine engine as set forth in claim 5, wherein said
propulsor is a fan rotor.
7. The gas turbine engine as set forth in claim 4, wherein said
swirl caused by said diverters provide efficient and thorough
mixing of said high pressure cooling air and said air at a higher
temperature.
8. The gas turbine engine as set forth in claim 3, wherein said
compressor section includes a low pressure compressor and a high
pressure compressor and said turbine section including at least a
high pressure turbine and a propulsor drive turbine, and said
propulsor drive turbine driving said propulsor through a gear
reduction.
9. The gas turbine engine as set forth in claim 2, wherein said
compressor section includes a low pressure compressor and a high
pressure compressor and said turbine section including at least a
high pressure turbine and a propulsor drive turbine, and said
propulsor drive turbine driving said propulsor through a gear
reduction.
10. A gas turbine engine comprising: a propulsor, a compressor
section and a turbine section, said compressor section having a
last compressor stage; high pressure cooling air tapped from a
location downstream of said last compressor stage and passed
through a heat exchanger; lower pressure air passing across said
heat exchanger to cool said high pressure cooling air; a housing
surrounding said compressor section and said turbine section and
there being a space radially outwardly of said housing, and a
mixing chamber received in said space radially outwardly of said
housing, said mixing chamber connected to receive said high
pressure cooling air at a location downstream of said heat
exchanger, and further connected to receive air at a temperature
higher than a temperature of said high pressure cooling air taken
downstream of said heat exchanger; mixed air from said mixing
chamber returned into said housing and utilized to cool at least
one of said turbine section or said last compressor stage; wherein
said high pressure cooling air is tapped from a diffuser section
radially outwardly of a combustor located intermediate said
compressor section and said turbine section; wherein said mixed air
passes through a compressor diffuser downstream of said last
compressor stage; and wherein said compressor section includes a
low pressure compressor and a high pressure compressor and said
turbine section including at least a high pressure turbine and a
propulsor drive turbine, and said propulsor drive turbine driving
said propulsor through a gear reduction.
11. The gas turbine engine as set forth in claim 10, wherein said
propulsor is a fan rotor.
12. The gas turbine engine as set forth in claim 11, wherein the
fan rotor has a plurality of fan blades, and a low fan pressure
ratio is defined across the fan blades alone, and said low fan
pressure ratio being less than or equal 1.45.
13. The gas turbine engine as set forth in claim 12, further
comprising a low corrected fan tip speed of less than 1150
ft/second.
14. The gas turbine engine as set forth in claim 11, further
comprising a low corrected fan tip speed of less than 1150
ft/second.
15. The gas turbine engine as set forth in claim 11, wherein an
outer housing surrounds said fan rotor, and said fan delivering air
into a bypass duct defined between said outer housing and an inner
core engine, and further delivering air to said compressor section
with a bypass ratio defined as the volume of air delivered into the
bypass duct compared to the volume of air delivered to said
compressor section, and said bypass ratio being greater than or
equal to 10.0.
16. The gas turbine engine as set forth in claim 10, wherein said
propulsor drive turbine having a pressure ratio defined as a ratio
of a pressure measured prior to an inlet to said propulsor drive
turbine as related to a pressure aft an outlet of the propulsor
drive turbine, prior to an exhaust nozzle, and said pressure ratio
being greater than or equal to 5.0.
17. The gas turbine engine as set forth in claim 10, wherein said
gear reduction having a gear ratio greater than or equal to
2.3.
18. The gas turbine engine as set forth in claim 10, wherein said
mixing chamber has diverters to cause a change in a flow direction
from said higher pressure cooling air and said air at a higher
temperature.
19. The gas turbine engine as set forth in claim 18, wherein said
diverters cause a change in a flow direction of the high pressure
cooling air and a flow direction of said air at a higher
temperature.
20. The gas turbine engine as set forth in claim 19, wherein said
diverters cause said high pressure cooling air to swirl and air at
a higher temperature to swirl.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application is a continuation of U.S. patent
application Ser. No. 15/424,927 filed on Feb. 6, 2017.
BACKGROUND OF THE INVENTION
[0002] This application relates to a mixing chamber for mixing two
air sources in a gas turbine engine to utilize as turbine cooling
air.
[0003] Gas turbine engines are known and typically include a fan
delivering air into a bypass duct as propulsion air and also into a
core engine. The air in the core engine enters a compressor where
it is compressed and then delivered into a combustor. The air is
mixed with fuel and ignited and products of this combustion pass
downstream over turbine rotors driving them to rotate.
[0004] Historically, a fan drive turbine rotated at a single speed
with the fan. However, more recently, a gear reduction has been
placed between the fan drive turbine and the fan rotor. This allows
the fan to increase in diameter and rotate at slower speeds, which
has many beneficial effects.
[0005] The fan drive turbine is able to rotate at higher speeds.
Temperatures within the gas turbine engine increase with this
change for several reasons. Further, the pressure downstream of the
combustor also increases.
[0006] As can be appreciated, components of the gas turbine engine
and, in particular, those in the turbine section, see very high
temperatures. It is known to provide cooling air to cool those
components. However, due to the increased pressure, it becomes
desirable to use highly pressurized air as the cooling air such
that it is able to move into the turbine section. The most
pressurized air in the gas turbine engine is downstream of a high
pressure compressor and it is typically hot.
[0007] This high pressure cooling air must be cooled. After
cooling, it may be at a temperature too low for effective use under
all conditions. Thus, it would be desirable to mix this cooing air
with higher temperature air. However, available space for a
required mixing chamber becomes difficult to locate within the
compact space of the modern gas turbine engine.
SUMMARY OF THE INVENTION
[0008] In a featured embodiment, a gas turbine engine comprises a
compressor section and a turbine section, the compressor section
having a last compressor stage. High pressure cooling air is tapped
from a location downstream of the last compressor stage and passed
through a heat exchanger. Lower pressure air passes across the heat
exchanger to cool the high pressure cooling air. A housing
surrounds the compressor section and the turbine section and there
being a space radially outwardly of the housing, and a mixing
chamber received in the space radially outwardly of the housing,
the mixing chamber receiving the high pressure cooling air
downstream of the heat exchanger, and further receiving air at a
temperature higher than a temperature of the high pressure cooling
air downstream of the heat exchanger. Mixed air from the mixing
chamber is returned into the housing and utilized to cool at least
the turbine section.
[0009] In another embodiment according to the previous embodiment,
the air at a higher temperature is also the high pressure cooling
air which is returned into the mixing chamber.
[0010] In another embodiment according to any of the previous
embodiments, the high pressure cooling air is tapped from a
diffuser section radially outwardly of a combustor located
intermediate the compressor section and the turbine section.
[0011] In another embodiment according to any of the previous
embodiments, the mixed air passes through a compressor diffuser
downstream of the last compressor stage.
[0012] In another embodiment according to any of the previous
embodiments, the mixing chamber has diverters to cause a change in
a flow direction from the higher pressure cooling air and the air
at a higher temperature.
[0013] In another embodiment according to any of the previous
embodiments, the diverters cause the higher pressure cooling air to
flow in a first circumferential direction and the air at a higher
temperature to flow in an opposed circumferential direction.
[0014] In another embodiment according to any of the previous
embodiments, the mixing chamber has diverters to cause a change in
a flow direction from the higher pressure cooling air and the air
at a higher temperature.
[0015] In another embodiment according to any of the previous
embodiments, the diverters cause the higher pressure cooling air to
flow in a first circumferential direction and air at a higher
temperature to flow in an opposed circumferential direction.
[0016] In another embodiment according to any of the previous
embodiments, the high pressure cooling air is tapped from a
diffuser section radially outwardly of a combustor located
intermediate the compressor section and the turbine section.
[0017] In another embodiment according to any of the previous
embodiments, the mixed air passes through a compressor diffuser
downstream of the last compressor stage.
[0018] In another embodiment according to any of the previous
embodiments, the mixing chamber has diverters to cause a change in
a flow direction from the higher pressure cooling air and the air
at a higher temperature.
[0019] In another embodiment according to any of the previous
embodiments, the diverters cause the higher pressure cooling air to
flow in a first circumferential direction and the air at a higher
temperature to flow in an opposed circumferential direction.
[0020] In another embodiment according to any of the previous
embodiments, the mixing chamber has diverters to cause a change in
a flow direction from the higher pressure cooling air and the air
at a higher temperature.
[0021] In another embodiment according to any of the previous
embodiments, the diverters causing the higher pressure cooling air
to flow in a first circumferential direction and the air at a
higher temperature to flow in an opposed circumferential
direction.
[0022] In another embodiment according to any of the previous
embodiments, the mixed air passes through a compressor diffuser
downstream of the last compressor stage.
[0023] In another embodiment according to any of the previous
embodiments, the mixing chamber has diverters to cause a change in
a flow direction from the higher pressure cooling air and the air
at a higher temperature.
[0024] In another embodiment according to any of the previous
embodiments, the diverters cause the higher pressure cooling air to
flow in a first circumferential direction and the air at a higher
temperature to flow in an opposed circumferential direction.
[0025] In another embodiment according to any of the previous
embodiments, the mixing chamber has diverters to cause a change in
a flow direction from the higher pressure cooling air and the air
at a higher temperature.
[0026] In another embodiment according to any of the previous
embodiments, the diverters cause the higher pressure cooling air to
flow in a first circumferential direction and the air at a higher
temperature to flow in an opposed circumferential direction.
[0027] In another embodiment according to any of the previous
embodiments, the compressor section includes a low pressure
compressor and a high pressure compressor and the turbine section
including at least a high pressure turbine and a fan drive turbine,
and a fan rotor, and the fan drive turbine driving the fan rotor
through a gear reduction.
[0028] These and other features may be best understood from the
following drawings and specification.
BRIEF DESCRIPTION OF THE DRAWINGS
[0029] FIG. 1 schematically shows a gas turbine engine.
[0030] FIG. 2 shows a cooling air system for an engine such as the
FIG. 1 engine.
DETAILED DESCRIPTION
[0031] FIG. 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 augmentor 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.
[0032] 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 bearing systems 38
may be varied as appropriate to the application.
[0033] 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 bearing systems 38 about the engine
central longitudinal axis A which is collinear with their
longitudinal axes.
[0034] 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.
[0035] 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 geared architecture 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.
[0036] 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. 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 .degree.R)/(518.7 .degree.R)].sup.0.5. The "Low corrected
fan tip speed" as disclosed herein according to one non-limiting
embodiment is less than about 1150 ft/second.
[0037] FIG. 2 schematically shows a cooling air system 100 which
may be incorporated into a gas turbine engine such as that shown in
FIG. 1. A last compressor rotor stage 102 has a downstream end 104.
A compressor diffuser 106 is shown downstream of the stage 102. A
combustor 108 is shown downstream of the compressor diffuser 106.
Products of combustion from the combustor 108 pass over a turbine
vane 133 and a first turbine rotor 134.
[0038] A diffuser chamber 110 sits radially outwardly of the
combustor 108. High pressure cooling air is tapped at line 112 from
the diffuser chamber 110. This cooling air passes through a heat
exchanger 114, cooling the air from line 112. Downstream of heat
exchanger 114, air from line 112 heads into line 118, having been
cooled to relatively low temperatures. As an example, the air at
tap 112 may be at a high temperature, such as 1300.degree. F., or
higher. However, by the time the air reaches line 118, it has been
cooled down to temperatures on the order of 500.degree. F.
[0039] The air in line 118 passes into a mixing chamber 116. Mixing
chamber 116 is radially outwardly of a housing 117 enclosing the
compressor 102, diffuser 106, turbine components 133/134 and
combustor 108.
[0040] By providing the mixing chamber 116 outwardly of the
housing, there is more efficient use of the space within the
housing 117.
[0041] The mixing chamber 116 becomes desirable because the air in
line 118 may be at too low a temperature to be utilized to cool the
turbine components 133/134 under certain conditions. Thus hot high
pressure air 122 may be tapped from line 112, upstream of heat
exchanger 114. This air also passes into the mixing chamber 116. As
mentioned, the air in line 122 downstream may be at a high
temperature.
[0042] Now, the air in line 122 can mix within the mixing chamber
116, as explained below.
[0043] Air is shown tapped at 120 from a lower pressure location
121 in the compressor section to pass over the heat exchanger 114
to cool the cooled high pressure air.
[0044] In the past it has been proposed to mix the high pressure
cooling air with the hot air radially inwardly of the housing 117.
This can raise challenges with regard to the available space.
Moreover, having inlet lines bringing in higher temperature and
lower temperature air can cause thermal gradients.
[0045] As shown in FIG. 2, the lower pressure cooling air at a high
temperature downstream of the heat exchanger 114 passes into line
122 and back into the mixing chamber 116. While the air in line 122
moving into the mixing chamber is shown as that tapped from line
112, any number of other air sources can be utilized to pass back
into the mixing chamber 116. That is, it is not necessarily air
from line 112 which is sent into the mixing chamber as hot air.
[0046] Diverter 124 diverts the air from line 118 in a swirl
direction and diverter 126 diverts the air from line 122.
Essentially, the two airflows now ensure efficient and thorough
mixing.
[0047] Downstream of the chamber 116, the mixed air passes at 127
into a chamber 128 and then through the compressor diffuser 106. As
known, the compressor diffuser 106 may have guide vanes and the air
can pass through those guide vanes.
[0048] The mixed cooling air then passes into a conduit 130,
although some mixed air may pass at 131 to cool the compressor
section. The air in conduit 130 passes as shown at 132 to cool a
turbine vane 133 and a turbine rotor 134.
[0049] It may sometimes be undesirable to pass air as cool as
500.degree. F. through the housing 117 and into the engine. That
is, air in line 118 might be too cool compared to the temperature
of the turbine section and thus could raise thermal gradients. As
such, by thoroughly mixing the air from lines 122 with air from
line 118, a desired temperature reaches line 127.
[0050] Although an embodiment of this invention has been disclosed,
a worker of ordinary skill in this art would recognize that certain
modifications would come within the scope of this invention. For
that reason, the following claims should be studied to determine
the true scope and content of this invention.
* * * * *