U.S. patent application number 09/952820 was filed with the patent office on 2003-03-20 for electric air conditioning system for an aircraft.
Invention is credited to Asfia, Julie F., Atkey, Warren A., Fiterman, Charles J., Loukusa, Steven M., Ng, Casey Y., Williams, Kenneth R..
Application Number | 20030051500 09/952820 |
Document ID | / |
Family ID | 25493265 |
Filed Date | 2003-03-20 |
United States Patent
Application |
20030051500 |
Kind Code |
A1 |
Asfia, Julie F. ; et
al. |
March 20, 2003 |
ELECTRIC AIR CONDITIONING SYSTEM FOR AN AIRCRAFT
Abstract
An all electric air conditioning system for an aircraft, wherein
the aircraft defines an interior volume having conditioned air at a
first pressure. A compressor is provided and is operable to
compress supply air to a second pressure. The compressor being
operated in response to an electrical drive motor. A passage
fluidly couples the compressor and the interior volume of the
aircraft. A heat dissipating device, such as a heat exchanger, is
positioned in the passage to extract heat energy from the supply
air. This arrangement permits conditioning of air within the
aircraft without using bleed air from the engines. The use of bleed
air results in a significant amount of fuel burn. An optional
conditioned air recovery system may be coupled to the interior
volume of the fuselage to direct at least a portion of the
conditioned air from the interior volume back for further
conditioning and use.
Inventors: |
Asfia, Julie F.; (Huntington
Beach, CA) ; Williams, Kenneth R.; (Huntington Beach,
CA) ; Atkey, Warren A.; (Bothell, WA) ;
Fiterman, Charles J.; (Mukilteo, WA) ; Loukusa,
Steven M.; (Bothell, WA) ; Ng, Casey Y.;
(Sammamish, WA) |
Correspondence
Address: |
HARNESS, DICKEY & PIERCE, P.L.C.
P.O. BOX 828
BLOOMFIELD HILLS
MI
48303
US
|
Family ID: |
25493265 |
Appl. No.: |
09/952820 |
Filed: |
September 14, 2001 |
Current U.S.
Class: |
62/402 |
Current CPC
Class: |
Y02T 50/50 20130101;
B64D 2013/0644 20130101; B64D 13/06 20130101; Y02T 50/44 20130101;
B64D 2013/0685 20130101; Y02T 50/56 20130101; Y02T 50/40 20130101;
Y02T 50/54 20130101; B64D 2013/0688 20130101 |
Class at
Publication: |
62/402 |
International
Class: |
F25D 009/00 |
Claims
What is claimed is:
1. An aircraft comprising: a fuselage defining an interior volume,
said interior volume having cabin air at a first pressure; a first
compressor operable to compress supply air to a second pressure; a
first electrical motor driving said first compressor; a passage
fluidly coupling said first compressor and said interior volume;
and a heat dissipating device positioned in said passage, said heat
dissipating device operable to extract heat energy from said supply
air.
2. The aircraft according to claim 1, further comprising: a vapor
system coupled downstream of said first compressor, said vapor
system being operable to remove water vapor from said supply
air.
3. The aircraft according to claim 2 wherein said vapor system
comprises: a reheater fluidly coupled to said heat dissipating
device; a condenser fluidly coupled to said reheater; and a water
collector fluidly coupled to said condenser.
4. The aircraft according to claim 2, further comprising: a vapor
system bypass valve positioned such that said vapor system bypass
valve is operable to selectively bypass said vapor system during
low vapor operation.
5. The aircraft according to claim 1, further comprising: a second
compressor fluidly coupled to said first compressor, said second
compressor being operable to compress said supply air from said
first compressor.
6. The aircraft according to claim 5, further comprising: a first
turbine operably coupled to one of said first compressor and said
second compressor.
7. The aircraft according to claim 5, further comprising: a first
turbine operably coupled to first compressor; and a second turbine
operably coupled to said second compressor.
8. The aircraft according to claim 5, further comprising: a second
electrical motor driving said second compressor.
9. The aircraft according to claim 1, further comprising: a cabin
air recovery system operably coupled to said first compressor, said
cabin air recovery system being operable to direct said cabin air
to one of said first compressor and said heat dissipating
device.
10. The aircraft according to claim 9 wherein said cabin air
recovery system comprises: a recovery turbine operably coupled to
said first compressor, said recovery turbine being operable to
direct said cabin air to one of said first compressor and said heat
dissipating device.
11. The aircraft according to claim 1 wherein said first pressure
and said second pressure are equal.
12. The aircraft according to claim 1 wherein said supply air is
ambient air outside of said fuselage.
13. The aircraft according to claim 1 wherein said supply air
includes at least a portion of said cabin air.
14. The aircraft according to claim 1 wherein said heat dissipating
device is a heat exchanger assembly, said heat exchanger assembly
including an inlet passage directing said supply air for use as a
heat sink.
15. The aircraft according to claim 1, further comprising: a ram
air fan rotatably coupled to said first compressor, said ram air
fan operable to supply a ram air to at least one of said first
compressor and heat dissipating device.
16. An aircraft comprising: a fuselage defining an interior volume;
an air conditioning system coupled to said fuselage, said air
conditioning system being operable to output conditioned air to
said interior volume; and a cabin air recovery system coupled to
said air conditioning system, said cabin air recovery system being
operable to direct at least a portion of said conditioned air from
said interior volume back to said air conditioning system for
further conditioning.
17. The aircraft according to claim 16 wherein said air
conditioning system comprises: a first compressor outputting
compressed air; a first electrical motor driving said first
compressor; a passage fluidly coupling said first compressor and
said interior volume; and a heat dissipating device positioned in
said passage, said heat dissipating device operable to extract heat
energy from said conditioned air.
18. The aircraft according to claim 17, further comprising: a vapor
system coupled downstream of said first compressor, said vapor
system being operable to remove water vapor from said conditioned
air.
19. The aircraft according to claim 18 wherein said vapor system
comprises: a reheater fluidly coupled to said heat dissipating
device; a condenser fluidly coupled to said reheater; and a water
collector fluidly coupled to said condenser.
20. The aircraft according to claim 19, further comprising: a first
turbine operably coupled to said first compressor, said first
turbine being fluidly positioned between said reheater and said
condenser.
21. The aircraft according to claim 20, further comprising: a
bypass valve positioned between said reheater and said condenser,
said bypass valve being operable to fluidly bypass said first
turbine.
22. The aircraft according to claim 18, further comprising: a vapor
system bypass valve positioned such that said vapor system bypass
valve is operable to selectively bypass said vapor system during
low vapor operation.
23. The aircraft according to claim 17, further comprising: a
second compressor fluidly coupled to said first compressor, said
second compressor being operable to compress said conditioned air
from said first compressor.
24. The aircraft according to claim 23, further comprising: a first
turbine operably coupled to one of said first compressor and said
second compressor.
25. The aircraft according to claim 23, further comprising: a first
turbine operably coupled to first compressor; and a second turbine
operably coupled to said second compressor.
26. The aircraft according to claim 23, further comprising: a
second electrical motor driving said second compressor.
27. The aircraft according to claim 17 wherein said cabin air
recovery system comprises: a recovery turbine operably coupled to
said first compressor, said recovery turbine being operable to
direct said conditioned air to one of said first compressor and
said heat dissipating device.
28. An aircraft comprising: a fuselage defining an interior volume,
said interior volume having conditioned air at a first pressure; a
first compressor operable to compress supply air to a second
pressure; a first electrical motor driving said first compressor; a
passage fluidly coupling said first compressor and said interior
volume; a heat dissipating device positioned in said passage, said
heat dissipating device operable to extract heat energy from said
supply air; and a conditioned air recovery system coupled to at
least one of said first compressor and said heat dissipating
device, said conditioned air recovery system being operable to
direct at least a portion of said conditioned air from said
interior volume to be mixed with said supply air.
29. The aircraft according to claim 28, further comprising: a vapor
system coupled downstream of said first compressor, said vapor
system being operable to remove water vapor from said supply air,
said vapor system having a reheater fluidly coupled to said heat
dissipating device, a condenser fluidly coupled to said reheater,
and a water collector fluidly coupled to said condenser.
30. The aircraft according to claim 29, further comprising: a vapor
system bypass valve positioned such that said vapor system bypass
valve is operable to selectively bypass said vapor system during
low vapor operation.
31. The aircraft according to claim 28, further comprising: a
second compressor fluidly coupled to said first compressor, said
second compressor being operable to compress said supply air from
said first compressor; and a first turbine operably coupled to one
of said first compressor and said second compressor.
32. The aircraft according to claim 31, further comprising: a
second turbine operably coupled to said second compressor.
33. The aircraft according to claim 31, further comprising: a
second electrical motor driving said second compressor.
34. The aircraft according to claim 28 wherein said conditioned air
recovery system comprises: a recovery turbine operably coupled to
said first compressor, said recovery turbine being operable to
direct said conditioned air to one of said first compressor and
said heat dissipating device.
35. The aircraft according to claim 28, further comprising: a ram
air fan rotatably coupled to said first compressor, said ram air
fan operable to supply a ram air to at least one of said first
compressor and heat dissipating device.
Description
FIELD OF THE INVENTION
[0001] The present invention generally relates to air conditioning
systems and, more particularly, to an electrically driven air
conditioning system for an aircraft that does not rely on engine
bleed air.
BACKGROUND OF THE INVENTION
[0002] Many air conditioning systems employed in modern commercial
aircraft utilize the air-to-air thermodynamic cycle to provide
cooling and/or heating air to the various compartments on the
aircraft, such as the passenger cabin, cargo holds, and the like.
Air from the compressor stages of the main jet propulsion engines,
also known as "bleed air," is generally output at high temperature
and pressure (i.e. 610OF and 60 psi). Conventionally, this bleed
air is then conditioned through conditioning packs before passing
into the pressurized fuselage for cabin temperature control,
ventilation, and pressurization. This conditioned air within the
fuselage is then discharged to the outside ambient air through
various overboard valves, overflow valves, and cabin leaks.
[0003] This known method of conditioning air for use with the
various aircraft systems is inefficient. That is, during a typical
steady state cruise operation, more energy than is necessary for
the primary requirements of the conditioning system (e.g. cabin
temperature control, ventilation, and pressurization) is added into
the conditioning system at the engines in the form of additional
fuel. Much of this excess energy is wasted in the form of heat and
pressure drop through ductwork, valves, and various other
components of the conditioning system. Moreover, extracting work
from the engines in the form of bleed air is inefficient relative
to other extraction methods. Consequently, the use of bleed air
from the engines reduces the efficiency of the engines and, thus,
increases the fuel consumption and load on the engines. By
eliminating or at least minimizing the use of bleed air in the
various aircraft systems, it is believed that more efficient jet
engines may be developed. Moreover, it is believed that alternative
air conditioning systems may lead to a reduction in aircraft
weight, assembly complexity, and fuel consumption.
[0004] Accordingly, there exists a need in the relevant art to
provide an air conditioning system for an aircraft that does not
rely on jet engine bleed air for operation. Furthermore, there
exists a need in the relevant art to provide an air conditioning
system for an aircraft that is capable of reducing the aircraft
weight, assembly complexity, and fuel consumption. Still further,
there exists a need in the relevant art to provide an air
conditioning system for an aircraft driven by electrical energy.
Moreover, there exists a need in the relevant art to provide an air
conditioning system for an aircraft that overcomes the
disadvantages of the prior art.
SUMMARY OF THE INVENTION
[0005] An all electric air conditioning system for an aircraft,
wherein the aircraft defines an interior volume having conditioned
air at a first pressure, is provided having an advantageous
construction. A compressor is provided and is operable to compress
supply air to a second pressure. The compressor being operated in
response to an electrical drive motor. A passage fluidly couples
the compressor and the interior volume of the aircraft. A
heat-dissipating device, such as a heat exchanger, is positioned in
the passage to extract heat energy from the supply air. This
arrangement permits conditioning of air within the aircraft without
using bleed air from the engines. The use of bleed air results in a
significant amount of fuel burn. An optional conditioned air
recovery system may be coupled to the interior volume of the
fuselage to direct at least a portion of the conditioned air from
the interior volume back for further conditioning and use.
[0006] Further areas of applicability of the present invention will
become apparent from the detailed description provided hereinafter.
It should be understood that the detailed description and specific
examples, while indicating the preferred embodiment of the
invention, are intended for purposes of illustration only and are
not intended to limit the scope of the invention.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] The present invention will become more fully understood from
the detailed description and the accompanying drawings,
wherein:
[0008] FIG. 1 is a circuit diagram illustrating a first embodiment
of the present invention in a ground or low altitude operation
configuration;
[0009] FIG. 2 is a circuit diagram illustrating the first
embodiment of the present invention in a cruise operation
configuration;
[0010] FIG. 3 is a circuit diagram illustrating a second embodiment
of the present invention in a ground or low altitude operation
configuration;
[0011] FIG. 4 is a circuit diagram illustrating the second
embodiment of the present invention in a cruise operation
configuration;
[0012] FIG. 5 is a circuit diagram illustrating a third embodiment
of the present invention in a ground or low altitude operation
configuration;
[0013] FIG. 6 is a circuit diagram illustrating the third
embodiment of the present invention in a cruise operation
configuration; and
[0014] FIG. 7 is a circuit diagram illustrating various alternative
modifications of the present invention.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
[0015] The following description of the preferred embodiments is
merely exemplary in nature and is in no way intended to limit the
invention, its application, or uses. That is, the electrical air
conditioning system of the present invention may find utility in
other applications, which commonly use bleed air to drive an air
conditioning system, such as in tanks and the like.
[0016] According to a first embodiment of the present invention, an
air conditioning system 10 is illustrated that is capable of
eliminating the use of bleed air. Air conditioning system 10
comprises a first compressor 12, a second compressor 14, a first
turbine 16, a second turbine 18, a heat exchanger assembly 20, a
reheater 22, a condenser 24, and a high-pressure water separator
26. With particular reference to FIG. 1, a ram air fan 28 provides
cold side air for heat exchanger assembly 20 during ground
operation. With particular reference to FIG. 2, a ram air inlet
scoop (not shown) provides cold side air for heat exchanger
assembly 20 during in-flight operation. Ram air, generally
indicated as 100, is provided to first compressor 12 via a passage
30 and further to heat exchanger assembly 20. A valve 62 controls
the amount of ram air 100 that is directed to heat exchanger
assembly 20. Alternatively, engine fan air could be used to provide
cool side air to first compressor 12 in place of ram air 100.
[0017] First compressor 12 is fluidly coupled to a primary heat
exchanger 32 of heat exchanger assembly 20 via a passage 34.
Primary heat exchanger 32 in turn is fluidly coupled to second
compressor 14 via a passage 36. Second compressor 14 in turn is
fluidly coupled to a secondary heat exchanger 38 via a passage 40.
Secondary heat exchanger 38 is fluidly separate from primary heat
exchanger 32. Hence, it should be understood that primary heat
exchanger 32 and secondary heat exchanger 38 might be configured as
separate units or a single unit having multiple discrete
chambers.
[0018] A compressor bypass valve 42 may fluidly interconnect
passage 36 and passage 40 so as to permit bypassing of second
compressor 14. Furthermore, an ozone converter 44 may be positioned
in series within passage 40 to permit proper conversion of ambient
air during a cruise phase of flight.
[0019] As seen in FIGS. 1 and 2, heat exchanger assembly 20 also
employs ram air 100 acting as a heat sink to remove excess heat
from the air upon exit from first compressor 12 and again upon exit
from second compressor 14. Trim air 46 may be extracted from
passage 40 for use in individual compartment temperature control or
for use in other aircraft systems.
[0020] Secondary heat exchanger 38 is fluidly coupled to reheater
22 via a passage 48. The cold outlet of reheater 22 is directed to
first turbine 16 through a passage 50 so as to be expanded and
reduced in temperature therein. This air is then directed into the
cold inlet of condenser 24 via a passage 52. It should be
appreciated that the cold inlet side of condenser 24 is maintained
above freezing to prevent ice formation. The air from condenser 24
is then directed to second turbine 18 via a passage 54 for final
expansion. Finally, air exits second turbine 18 via passage 56 and
is directed to a mix manifold (not shown) for distribution into the
aircraft cabin.
[0021] Still referring to FIGS. 1 and 2, air conditioning system 10
further includes a passage 58 fluidly interconnecting condenser 24
to water collector 26 and a passage 60 fluidly interconnecting
water collector 26 to reheater 22. A passage 61 is further provided
that fluidly interconnects reheater 22 and condenser 24. Air
conditioning system 10 still further includes a ram air modulator
valve/actuator 62 used for controlling the flow of ram air 100. A
turbine bypass valve 64 fluidly interconnects passage 54 and
passage 56 so as to permit bypassing of second turbine 18. A first
motor 66 is operably coupled between first compressor 12 and second
turbine 18 and a second motor 68 is operably coupled between second
compressor 14 and first turbine 16.
[0022] With particular reference to FIG. 2, it can be seen that air
conditioning system 10 further includes an altitude valve 70
fluidly interconnecting passage 48 and passage 54. Altitude valve
70 permits bypassing of a condensing loop 72 above a predetermined
altitude. Condensing loop 72 generally includes reheater 22,
passage 50, first turbine 16, passage 52, condenser 24, passage 58,
water collector 26, passage 60, and passage 61. The opening of
altitude valve 70 bypasses condensing loop 72 such that primary
cooling of the air occurs in heat exchanger assembly 20 and second
turbine 18. This arrangement at altitude enables the overall
pressure drop in the system to be minimized so as to provide
sufficient flow to the passenger cabin at lower power consumption
levels.
[0023] Referring to FIGS. 1 and 2, during operation, first
compressor 12 receives ambient air 100 from ram air fan 28. This
air is compressed within first compressor 12 and is passed through
primary heat exchanger 32 of heat exchanger assembly 20 to second
compressor 14. Primary heat exchanger 32 removes heat from the air
using ram air 100 as a heat sink. The air is then compressed within
second compressor 14 and passed through secondary heat exchanger 38
of heat exchanger assembly 20. Second compressor 14 may be bypassed
using compressor bypass valve 42.
[0024] During ground or low altitude operation, air then exits
secondary heat exchanger 38 and is directed to reheater 22. The
cold outlet of reheater 22 directs air to first turbine 16 where
the temperature and pressure are reduced. The air is then directed
to condenser 24 to remove excess water from the air. The cold
outlet of condenser 24 directs the air to second turbine 18 where
the temperature and pressure are further reduced. Lastly, the air
is then directed to the mixing manifold and distributed to the
aircraft cabin.
[0025] Air and water from condenser 24 flows to water collector 26
through passage 58, where water is collected by water collector
26.
[0026] During high altitude operation, air from secondary heat
exchanger 38 of heat exchanger assembly 20 is directed through
altitude valve 70 so as to completely bypass condensing loop 72.
Accordingly, air flows from secondary heat exchanger 38 directly to
second turbine 18 so as to minimize the pressure drop within system
10 during high altitude cruise. Therefore, power consumption is
minimized.
[0027] According to a second embodiment of the present invention,
an air conditioning system 10' is illustrated that is capable of
eliminating the use of bleed air and further capable of utilizing
the potential energy of pressurized air leaving the aircraft cabin
during high altitude flight.
[0028] With particular reference to FIGS. 3 and 4, in addition to
those elements described in reference to FIGS. 1 and 2, air
conditioning system 10' further includes an outflow turbine 110.
Outflow turbine 110 is illustrated as being operably coupled to
motor 68 and first turbine 16. However, it must be understood that
outflow turbine 110 may be alternatively coupled to first
compressor 12 or second turbine 18. Outflow turbine 110 receives
previously conditioned air from the cabin of the aircraft through a
cabin recovery valve 112. Cabin recovery valve 112 is actuated to
provide flow of conditioned air through outflow turbine 110. It
should be understood that cabin recovery valve 112 or outflow
turbine 110 might include an integral anti-depressurization valve
to guard against inadvertent depressurization of the aircraft
cabin. That is, should a duct burst or other failure to occur,
anti-depressurization valve will close to prevent further
depressurization of the aircraft cabin. The anti-depressurization
valve may be a conventional aerodynamic valve that closes upon
sensing too much air flow.
[0029] Cabin air 114 is directed through outflow turbine 110 where
it is quickly expanded. This expansion of cabin air 114 causes a
rapid temperature drop of cabin air 114, which is directed through
passage 116 to heat exchanger assembly 20. This cooled air serves
to supplement ram air 100, thereby reducing the drag associated
with the ram air system by not requiring as much outside ambient
air for heat exchanger assembly 20 cooling. Moreover, the power
generated by outflow turbine 110 serves to reduce the work required
by motor 68 when driving second compressor 14. A significant
electrical power and ram air drag saving is achieved as the cruise
phase is the majority of the entire flight.
[0030] Still referring to FIGS. 3 and 4, during ground or low
altitude operation, air conditioning system 10' works identically
to air conditioning system 10. However, during high altitude
operation, as described above, cabin air 114 is expanded and cooled
in outflow turbine 110 and is passed to heat exchanger assembly 20
for cooling. Like air conditioning system 10, condensing loop 72 is
bypassed using altitude valve 70. The opening of altitude valve 70
bypasses condensing loop 72 such that primary cooling of the air
occurs in heat exchanger assembly 20, supplementing with expanded
cabin air 114, and second turbine 18. This arrangement, at
altitude, enables the overall pressure drop in system 10' to be
minimized so as to provide sufficient flow to the passenger cabin
at lower power consumption levels.
[0031] According to a third embodiment of the present invention, an
air conditioning system 10" is illustrated that is capable of
eliminating the use of bleed air and further capable of utilizing
the potential energy of pressurized air leaving the aircraft cabin
during high altitude flight. However, unlike the second embodiment
of the present invention, air conditioning system 10" employs a
series of control valves such that first turbine 16 acts similar to
outflow turbine 110 of the second embodiment.
[0032] More particularly, as best seen in FIGS. 5 and 6, in
addition to those elements described in reference to FIGS. 1 and 2,
air conditioning system 10" further includes a first cabin recovery
valve 210. First turbine 16 receives previously conditioned air 114
from the cabin of the aircraft through first cabin recovery valve
210. First cabin recovery valve 210 is variably actuated to control
the preferred flow of conditioned air 114 into passage 50.
Conditioned air 114 joins air flow within passage 50 and is
directed to first turbine 16 where it is expanded and cooled. It
should be understood that first cabin recovery valve 210 or first
turbine 16 may include an integral anti-depressurization valve to
guard against inadvertent depressurization of the aircraft cabin.
That is, should a duct burst or other failure to occur,
anti-depressurization valve will close to prevent further
depressurization of the aircraft cabin. The anti-depressurization
valve may be a conventional aerodynamic valve that closes upon
sensing too much air flow.
[0033] Air conditioning system 10" further includes a second cabin
recovery valve 212 disposed within passage 50 upstream from the
inflow of cabin air 114. Second cabin recovery valve 212 is
selectively actuated to prohibit air flow from reheater 22 to first
turbine 16 and backflow of cabin air 114 to reheater 22. A third
cabin recovery valve 214 is disposed within a passage 216
interconnecting passage 52 and heat exchanger assembly 20. A check
valve 218 is further disposed in passage 52 downstream from the
interconnection with passage 216. Check valve 218 prevents backflow
of air from condenser 24 in the event of a failure of third cabin
recovery valve 214.
[0034] Still referring to FIGS. 5 and 6, during ground or low
altitude operation, air conditioning system 10" works identically
to air conditioning system 10. However, during high altitude
operation, cabin air 114 is expanded and cooled in first turbine 16
and is passed to heat exchanger assembly 20 for cooling. Like air
conditioning system 10, condensing loop 72 is bypassed using
altitude valve 70 and the bypass valves are actuated to direct
cabin air 114 to first turbine 16 and heat exchanger assembly 20.
Specifically, first cabin recovery valve 210 is opened to allow
flow of cabin air 114 into a passage 220. Cabin air 114 is then
directed to first turbine 16 via passage 50 by closing second cabin
recovery valve 212. Cabin air 114 is then expanded and cooled and
used to supplement ram air 100 in heat exchanger assembly 20. Check
valve 218 prevents flow through a failed-open valve 214 to the ram
system. The opening of altitude valve 70 bypasses condensing loop
72 such that primary cooling of the air occurs in heat exchanger
assembly 20, supplementing with expanded cabin air 114, and second
turbine 18. This arrangement, at altitude, enables the overall
pressure drop in system 10" to be minimized so as to provide
sufficient flow to the passenger cabin at lower power consumption
levels.
[0035] In addition to the above embodiments described in detail,
there are numerous modifications that are anticipated to further
tailor the air conditioning system of the present invention.
However, it must be understood that each of the following
modifications, although described together, is individually
applicable to the above described embodiments. That is, each
modification may be employed separately from the remaining
modifications, if desired. They are simply being described together
here in the interest of brevity.
[0036] Referring to FIG. 7, it should be understood that ram air
fan 28 may alternatively be coupled to second compressor 14,
generally indicated at 28'. Ram air fan 28' would thus supply ram
air to second compressor 14. Still referring to FIG. 7, primary
heat exchanger 32 may be eliminated if it is determined that a
two-stage heat exchanger system is not required, thereby generally
designated as 20'. Similarly, motor 68 may be eliminated if added
mechanical input is not required between second compressor 14 and
first turbine 16. Likewise, second turbine 18 may be eliminated if
the necessary temperature and pressure are achieved depending on
the equipment used and the aircraft requirements. However, it is
preferable that if second turbine 18 is eliminated, then turbine
bypass valve 64 be similarly eliminated since its use is now
defeated. Alternatively, turbine bypass valve 64 may be
repositioned between passage 50 and passage 52, thereby serving to
selectively bypass first turbine 16.
[0037] Existing aircraft require the use of bleed air to operate
the aircraft air conditioning system. However, bleed air requires a
significant amount of fuel burn where a significant amount of
energy is wasted by the processing of the bleed air. Hence, there
is a need in modern designs to alleviate the use of bleed air in
air conditioning systems. According to the principles of the
present invention, an all electrical air conditioning system is
provided that eliminates the need for bleed air. Moreover, the
present invention enables much of the energy of the conditioned air
within the cabin to be recovered, thereby reducing electrical power
consumption. The elimination of the use of bleed air enables
aircraft engines to be more efficiency designed, thereby reducing
the use of fuel. It should be appreciated that extracting
electricity from jet engines is much more efficient than extracting
bleed air. Still further, the present invention provides a method
of reducing the weight and maintenance requirements of the aircraft
since engine pneumatic ducting, APU ducting, and pneumatic
components are eliminated. Duct leaks may be eliminated or at least
reduce while overheat detection systems may no longer be necessary.
Additionally, air conditioning systems may be modularized, since
they no longer need to be sized relative to APU/Engine pneumatic
operation performance.
[0038] The description of the invention is merely exemplary in
nature and, thus, variations that do not depart from the gist of
the invention are intended to be within the scope of the invention.
Such variations are not to be regarded as a departure from the
spirit and scope of the invention.
* * * * *