U.S. patent application number 13/015109 was filed with the patent office on 2012-08-02 for gas turbine bleed ecs cooling.
This patent application is currently assigned to HAMILTON SUNDSTRAND CORPORATION. Invention is credited to Adam M. Finney.
Application Number | 20120192578 13/015109 |
Document ID | / |
Family ID | 46576187 |
Filed Date | 2012-08-02 |
United States Patent
Application |
20120192578 |
Kind Code |
A1 |
Finney; Adam M. |
August 2, 2012 |
GAS TURBINE BLEED ECS COOLING
Abstract
A method for managing heat in an airplane includes extracting
hot bleed air from a gas turbine engine, and cooling the hot bleed
air in a first heat exchanger thereby forming warm bleed air. The
method further includes expanding the warm bleed air with a turbine
thereby forming cool bleed air, and using the cool bleed air as a
heat sink in a second heat exchanger associated with an
environmental control system and thereby forming used bleed
air.
Inventors: |
Finney; Adam M.; (Rockford,
IL) |
Assignee: |
HAMILTON SUNDSTRAND
CORPORATION
Windsor Locks
CT
|
Family ID: |
46576187 |
Appl. No.: |
13/015109 |
Filed: |
January 27, 2011 |
Current U.S.
Class: |
62/87 ;
62/402 |
Current CPC
Class: |
B64D 13/06 20130101;
Y02T 50/50 20130101; B64D 2013/0618 20130101; Y02T 50/56
20130101 |
Class at
Publication: |
62/87 ;
62/402 |
International
Class: |
F25B 9/06 20060101
F25B009/06 |
Claims
1. A method for managing heat in an airplane, the method
comprising: extracting hot bleed air from a gas turbine engine;
cooling the hot bleed air in a first heat exchanger thereby forming
warm bleed air; expanding the warm bleed air with a turbine thereby
forming cool bleed air; and using the cool bleed air as a heat sink
in a second heat exchanger associated with an environmental control
system and thereby forming used bleed air.
2. The method of claim 1, further comprising: exhausting the used
bleed air out of the airplane.
3. The method of claim 1, wherein the hot bleed air is extracted
from one of a fan bleed port, an intermediate compressor bleed
port, and a high compressor bleed port of the gas turbine
engine.
4. The method of claim 1, wherein the hot bleed air is cooled in
the first heat exchanger by one of ram air, lower stage bleed air,
or fuel.
5. The method of claim 1, further comprising: operating one of a
pump, a compressor, a generator, or a gearbox with energy generated
by expanding the warm bleed air with the turbine.
6. The method of claim 1, wherein the cool bleed air cools a
working fluid in the second heat exchanger for one of an air cycle
system, a vapor cycle system, or a hybrid cooling system.
7. A heat management system for an airplane, the system comprising:
a first heat exchanger for cooling the hot bleed air produced by a
gas turbine engine, thereby forming warm bleed air; a turbine for
expanding the warm bleed air thereby forming cool bleed air; and a
second heat exchanger using the cool bleed air as a heat sink for a
fluid in an environmental control system.
8. The system of claim 7, wherein the hot bleed air is tapped from
the gas turbine engine at one of a fan section, an intermediate
compressor section, and a high compressor section.
9. The system of claim 7, wherein one of ram air, lower stage bleed
air, or fuel enter the first heat exchanger to provide the heat
sink for cooling the hot bleed air.
10. The system of claim 7, further comprising: a pump attached to
the turbine, the pump for using work generated by the turbine
expanding the warm bleed air.
11. The system of claim 7, further comprising: a compressor
attached to the turbine, the compressor for using work generated by
the turbine expanding the warm bleed air.
12. The system of claim 7, further comprising: a generator attached
to the turbine, the generator for using work generated by the
turbine expanding the warm bleed air.
13. The system of claim 7, further comprising: a gearbox attached
to the turbine, the gearbox for using work generated by the turbine
expanding the warm bleed air.
14. The system of claim 7, wherein the environmental control system
is one of an air cycle system, a vapor cycle system, or a hybrid
cooling system.
15. A heat management system for an airplane, the system
comprising: a gas turbine engine having a bleed port for extracting
hot bleed air; a first heat exchanger fluidly connected to the
bleed port, the heat first heat exchanger configured to cool the
hot bleed air into warm bleed air; a turbine fluidly connected to
the first heat exchanger, the turbine configured to expand the warm
bleed air into cool bleed air; and a second heat exchanger fluidly
connected to the turbine, the second heat exchanger configured to
cool a fluid for an environmental control system with the cool
bleed air.
16. The system of claim 15, wherein the bleed port is located at
one of a fan section, an intermediate compressor section, and a
high compressor section.
17. The system of claim 15, wherein one of ram air, lower stage
bleed air, or fuel enter the first heat exchanger to cool the hot
bleed air into warm bleed air.
18. The system of claim 15, wherein one of a pump, a compressor, a
generator, or a gearbox is attached to the turbine for using work
generated by expanding the warm bleed air.
19. The system of claim 15, wherein the environmental control
system is one of an air cycle system, a vapor cycle system, or a
hybrid cooling system.
Description
BACKGROUND
[0001] The present disclosure relates generally to airplanes and
more specifically, to environmental control systems in
airplanes.
[0002] Most modern military and commercial airplanes are powered by
a gas turbine engine and include an Environmental Control System
(ECS). Known ECS include air cooling and conditioning equipment
such as compressors, valves, and heat exchangers, which are
sometimes arranged into an ECS "pack". Hot pressurized air is
commonly taken from the gas turbine engine and ducted to the ECS
pack for cooling and conditioning. This conditioned air is then
provided to a cabin and/or flight deck for use by crew, passengers,
and/or equipment on the airplane.
SUMMARY
[0003] A method for managing heat in an airplane is disclosed. The
method includes extracting hot bleed air from a gas turbine engine,
and cooling the hot bleed air in a first heat exchanger thereby
forming warm bleed air. The method further includes expanding the
warm bleed air with a turbine thereby forming cool bleed air, and
using the cool bleed air as a heat sink in a second heat exchanger
associated with an environmental control system thereby forming
used bleed air.
[0004] A heat management system for an airplane is also disclosed.
The system includes a first heat exchanger, a turbine, and a second
heat exchanger. A gas turbine engine produces hot bleed air, and
the first heat exchanger cools the hot bleed air thereby forming
warm bleed air. The turbine expands the warm bleed air thereby
forming cool bleed air, and the second heat exchanger uses the cool
bleed air as a heat sink for a fluid in an environmental control
system.
[0005] In another embodiment of the heat management system, the gas
turbine engine has a bleed port for extracting hot bleed air. The
first heat exchanger is fluidly connected to the bleed port and
configured to cool the hot bleed air into warm bleed air. The
turbine is fluidly connected to the first heat exchanger and
configured to expand the warm bleed air into cool bleed air. The
second heat exchanger is fluidly connected to the turbine and
configured to cool a fluid for an environmental control system with
the cool bleed air.
BRIEF DESCRIPTION OF THE DRAWINGS
[0006] FIG. 1 is a schematic of an airplane having a gas turbine
engine.
[0007] FIG. 2 is a schematic of the gas turbine engine and a heat
management system in accordance with the present disclosure.
[0008] FIG. 3 is a flow chart of a method for managing heat in
accordance with the present disclosure.
DETAILED DESCRIPTION
[0009] As used herein, the term "airplane" includes any type of
aircraft having an environmental control system. FIG. 1 is a
perspective view of airplane 10 with cut-away to show a
cross-section of gas turbine engine 12. Engine 12 includes
compressor section 14, combustor section 16, and turbine section
18. Compressor section 14 further includes a fan 20, low pressure
compressor (LPC) and high pressure compressor (HPC) 22. Primary air
flow 24 travels through compressor section 14, combustor section
16, and turbine section 18 of engine 12 to generate power for
airplane 10.
[0010] Engine 12 is attached to a wing of airplane 10 and includes
compressor section 14, combustor section 16, and turbine section 18
in flow series. Fan 20, LPC and HPC 22 are located within
compressor section 14 in flow series. In operation, primary air
flow 24 enters Fan 20 of compressor section 14 and is compressed.
Next, partially compressed primary air flow 24 enters LPC and HPC
22 of compressor section 14 and experiences further compression.
Fully compressed primary air flow 24 then enters combustor section
16, which mixes fuel with fully compressed primary air flow 24 and
combusts the mixture. Combusted primary air flow 24 enters turbine
section 18 and experiences expansion, which forces one or more
turbines to rotate. The rotating turbine then drives Fan 20, LPC
and HPC 22.
[0011] FIG. 2 is a schematic of gas turbine engine 12 and heat
management system 26 in accordance with the present disclosure.
Engine 12 includes compressor section 14, combustor section 16,
turbine section 18, fan 20, LPC and HPC 22, primary air flow 24,
fan bleed port 28, low pressure or intermediate compressor (IC)
bleed port 30, high pressure (HP) bleed port 32, and secondary air
flow or hot bleed air 34 including fan bleed 34A, intermediate
compressor (IC) bleed 34B, and high compressor (HC) bleed 34C. Heat
management system 26 includes first heat exchanger 36, warm bleed
air 38, turbine 40, cool bleed air 42, second heat exchanger 44,
and used bleed air 46. Also depicted in FIG. 2 are first cool fluid
48, first warm fluid 50, and potential sources 52 for first cool
fluid 48 including ram air 52A, lower stage (LS) bleed air 52B, or
fuel 52C. Shaft 54 of turbine 40 is coupled to energy sink 56
including pump 56A, compressor 56B, generator 56C, or gearbox 56D.
Second warm fluid 58, second cool fluid 60, and portions of
environmental control system (ECS) 62 including vapor cycle system
62A, air cycle system 62B, and hybrid cooling system 62C are also
depicted in FIG. 2. Hot bleed air 34 is extracted from engine 12,
cooled, expanded, and used as a heat sink for a fluid in ECS
62.
[0012] As described above with reference to FIG. 1, engine 12
includes fan 20, LPC and HPC 22 of compressor section 14, combustor
section 16, and turbine section 18 in flow series. Shown in FIG. 2
are three engine bleed ports: fan bleed port 28, IC bleed port 30,
and HP bleed port 32. Fan bleed port 28 is located just downstream
of fan 20 and upstream of LPC. IC bleed port 30 is located just
downstream of LPC and upstream of HPC 22, downstream of fan bleed
port 28 and upstream of combustor section 16. HP bleed port 32 is
located at an end of HPC 22 just upstream of combustor section 16.
Primary air flow 24 flows through fan 20, LPC and HPC 22 of
compressor section 14, combustor section 16, and turbine section 18
of engine 12 to power airplane 10. A secondary air flow enters fan
20 along with primary air flow 24, but exits engine 12 at one or
more of fan bleed port 28, IC bleed port 30, and HP bleed port 32.
This secondary airflow or hot bleed air 34 can be extracted or
"bled" from fan bleed port 28 as fan bleed air 34A, from IC bleed
port 30 as IC bleed air 34B, and/or from HP bleed port 32 as HP
bleed air 34C. One or more of fan bleed port 28, IC bleed port 30,
and HP bleed port 32 are chosen depending on the type of airplane
10 and the application of hot bleed air 34. Once extracted from
engine 12, hot bleed air 34 enters heat management system 26.
[0013] Heat management system 26 includes three subsystems; each
subsystem is centered on one of first heat exchanger 36, turbine
40, and second heat exchanger 44. Each subsystem will be discussed
in turn, beginning with first subsystem 26A of heat management
system 26 centered upon first heat exchanger 36. Hot bleed air 34
enters an air inlet side of first heat exchanger 36 and warm bleed
air 38 exits an air outlet side of first heat exchanger 36. First
cool fluid 48 enters a fluid inlet side of first heat exchanger 36,
and first warm fluid 50 exits a fluid outlet side of first heat
exchanger 36. Source 52 of first cool fluid 48 can be ram air 52A,
LS bleed air 52B, fuel 52C, or any other first cool fluid 48 that
is cooler than hot bleed air 34. First cool fluid 48 is used as a
heat sink for hot bleed air 34 within heat exchanger 36. First cool
fluid 48 picks up heat from hot bleed air 34 such that hot bleed
air 34 becomes warm bleed air 38 and first cool fluid 48 becomes
first warm fluid 50.
[0014] After exiting first heat exchanger 36, warm bleed air 38
enters turbine 40, and second subsystem 26B_of heat management
system 26. Warm bleed air 38 enters an air inlet of turbine 40 and
cool bleed 42 air exits an air outlet of turbine 40. Shaft 54
attaches turbine 40 energy sink 56. Energy sink 56 can be a pump
56A, compressor 56B, generator 56C, gearbox 56D, or any other
component requiring energy input. After being cooled by first heat
exchanger 36, warm bleed air 38 is cooled further and expanded by
turbine 40 to become cool bleed air 42. The work created by the
expansion of warm bleed air 38 is transmitted through shaft 54 and
used by energy sink 56. In other words, turbine 40 can provide
energy to drive one or more of pump 56A, compressor 56B, generator
56C, or gearbox 56D.
[0015] After exiting turbine 40, cool bleed air 42 enters second
heat exchanger 44, and third subsystem 26C of heat management
system 26. Cool bleed air 42 enters an air inlet of second heat
exchanger 44 and warm bleed air 46 exits an air outlet of second
heat exchanger 44. Second warm fluid 58 enters a fluid inlet of
second heat exchanger 44 and second cool fluid 60 exits a fluid
outlet of second heat exchanger 44. Any portion of ECS 62 can
provide second warm fluid 58 to second heat exchanger 44 including
vapor cycle system 62A, air cycle system 62B, and/or hybrid cooling
system 62C, so long as second warm fluid 58 is warmer than cool
bleed air 42. Warm fluid 58 is used as a heat sink for cool bleed
air 42 within second heat exchanger 44. Cool bleed air 42 picks up
heat from second warm fluid 58 such that cold bleed air 42 becomes
warm bleed air 46 and second warm fluid 58 becomes second cool
fluid 60. In the depicted embodiment, second cool fluid 60 is
returned to ECS 62 for use as a heat sink in one of vapor cycle
system 62A, air cycle system 62B, or hybrid cooling system 62C.
Accordingly, hot bleed air 34 from engine 12 is cooled and expanded
into cool bleed air 42 and used as a heat sink for ECS 62.
[0016] FIG. 3 is a flow chart of the method for managing heat 64 in
accordance with the present disclosure. Shown in FIG. 3 are steps
of method 64: extract 66, cool 68, expand 70, heat exchange 72, and
exhaust 74. Method 64 shows how to use hot bleed air 34 from engine
12 to cool a second warm fluid 58 from ECS 62. First, hot bleed air
34 is extracted from gas turbine engine 12 (extract step 66). As
described with reference to FIG. 2, hot bleed air 34 can be
extracted from one or more of fan bleed port 28, IP bleed port 30,
or HP bleed port 32 as fan bleed air 34A, IP bleed air 34B, or HP
bleed air 34C, respectively. Second, hot bleed air 34 is cooled in
first heat exchanger 36 thereby forming warm bleed air 38 (cool
step 68). As described with reference to FIG. 2, cool fluid 48 is
obtained from any source 52 such as ram air 52A, LS bleed air 52B,
fuel 52C. Third, warm bleed air 38 is expanded with turbine 40
thereby forming cool bleed air 42 (expand step 70). As described
with reference to FIG. 2, turbine 40 is energetically attached to
energy sink 56 such as pump 56A, compressor 56B, generator 56C, or
gearbox 56D. Fourth, cool bleed air 42 is used as a heat sink in
second heat exchanger 44 associated with ECS 62 (heat exchange step
72). Cool bleed air 42 becomes used bleed air 46 when used as a
heat sink in second heat exchanger 44. Fifth, used bleed air 46 can
be used as a heat sink for another portion of ECS 26 and/or
exhausted outside of airplane (exhaust step 74).
[0017] Heat management system 26 and method for managing heat 64
are particularly advantageous for airplanes traveling at high mach
speeds. At supersonic speeds, ambient/ram air and fuel may not be
suitable for use as a heat sink because they may be simply too hot.
Heat management system 26 and method for managing heat 64 provide
an alternative to ambient/ram air and fuel whereby hot bleed air 34
from engine 12 is converted into cool bleed air 42 for use as a
heat sink for a fluid in ECS 62.
[0018] While the invention has been described with reference to an
exemplary embodiment(s), it will be understood by those skilled in
the art that various changes may be made and equivalents may be
substituted for elements thereof without departing from the scope
of the invention. In addition, many modifications may be made to
adapt a particular situation or material to the teachings of the
invention without departing from the essential scope thereof.
Therefore, it is intended that the invention not be limited to the
particular embodiment(s) disclosed, but that the invention will
include all embodiments falling within the scope of the appended
claims.
* * * * *