U.S. patent application number 15/457457 was filed with the patent office on 2017-06-29 for system and method for operating a precooler in an aircraft.
The applicant listed for this patent is BOMBARDIER INC.. Invention is credited to Sigit AFRIANTO, Jean BROUSSEAU, Valerie DESILETS, Remi HAMEL.
Application Number | 20170184030 15/457457 |
Document ID | / |
Family ID | 45929030 |
Filed Date | 2017-06-29 |
United States Patent
Application |
20170184030 |
Kind Code |
A1 |
BROUSSEAU; Jean ; et
al. |
June 29, 2017 |
SYSTEM AND METHOD FOR OPERATING A PRECOOLER IN AN AIRCRAFT
Abstract
A method for controlling a pressure control mechanism in a
turbofan engine having a precooler permitting heat exchange between
ambient air and bleed air includes detecting at least one of an
engine failure or a bleed system failure, detecting at least one of
an ice condition or an activation of an anti-ice system, and
actuating the pressure control mechanism, thereby altering the heat
exchange between the ambient air and the bleed air.
Inventors: |
BROUSSEAU; Jean;
(Pierrefonds, CA) ; AFRIANTO; Sigit; (Pierrefonds,
CA) ; HAMEL; Remi; (Laval, CA) ; DESILETS;
Valerie; (Vaudreuil, CA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
BOMBARDIER INC. |
Dorval |
|
CA |
|
|
Family ID: |
45929030 |
Appl. No.: |
15/457457 |
Filed: |
March 13, 2017 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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14004974 |
Sep 13, 2013 |
9624831 |
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PCT/US2012/029368 |
Mar 16, 2012 |
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15457457 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
B64D 13/08 20130101;
Y02T 50/671 20130101; F02K 1/822 20130101; Y02T 50/675 20130101;
F02C 6/08 20130101; Y02T 50/44 20130101; Y02T 50/40 20130101; Y02T
50/60 20130101; F02K 3/075 20130101; B64D 2013/0618 20130101; F01D
25/14 20130101; F02C 7/00 20130101; Y02T 50/50 20130101; F02C 9/18
20130101; B64D 33/10 20130101; F02K 3/06 20130101; F01D 25/02
20130101; Y02T 50/56 20130101; F02C 7/14 20130101; F02K 3/115
20130101 |
International
Class: |
F02C 7/14 20060101
F02C007/14; F02K 3/115 20060101 F02K003/115; F01D 25/02 20060101
F01D025/02; F02C 6/08 20060101 F02C006/08; F02K 3/06 20060101
F02K003/06; B64D 13/08 20060101 B64D013/08 |
Claims
1. A method for controlling a pressure control mechanism in a
turbofan engine having a precooler permitting heat exchange between
ambient air and bleed air, the method comprising: detecting at
least one of an engine failure or a bleed system failure; detecting
at least one of an ice condition or an activation of an anti-ice
system; and actuating the pressure control mechanism, thereby
altering the heat exchange between the ambient air and the bleed
air.
2. The method of claim 1, wherein the pressure control mechanism
comprises a pressure relief door disposed proximate to an outlet of
the turbofan engine.
3. The method of claim 1, wherein the pressure control mechanism is
actuated upon detection of a combination of at least one of an
engine failure and a bleed system failure, and at least one of an
ice condition and an activation of the anti-ice system.
4. The method of claim 1, wherein actuating the pressure control
mechanism comprises opening a pressure relief door in the turbofan
engine.
5. The method of claim 2, wherein: the pressure relief door
comprises a plurality of pressure relief doors, and actuating the
plurality of pressure relief doors comprises opening the plurality
of pressure relief doors.
6. The method of claim 1, wherein the bleed air has a temperature
greater than the ambient air.
7. The method of claim 6, wherein the bleed air is cooled to a
predetermined temperature within the precooler via the heat
exchange with the ambient air.
8. The method of claim 7, wherein the predetermined temperature is
between about 200.degree. C. to 232.degree. C.
9. The method of claim 1, further comprising: providing an ambient
air passage for the ambient air; providing a bleed air passage for
the bleed air; and separating the ambient air passage and the bleed
air passage from one another such that the ambient air and the
bleed air are prevented from intermixing during the heat
exchange.
10. The method of claim 1, further comprising: providing an ambient
air passage for the ambient air; providing a bleed air passage for
the bleed air; and combining the ambient air passage and the bleed
air passage to permit intermixing of the ambient air and the bleed
air during the heat exchange.
11. The method of claim 1, wherein the bleed air is taken from
positions proximate to both a high pressure compressor and a low
pressure compressor in the turbofan engine.
12. The method of claim 1, wherein a demand for increased flow of
the ambient air is conditioned upon trigger variables, which
comprise signals indicative of an engine failure, a bleed air
system failure, an ice condition, or an activation of an anti-icing
system.
13. The method of claim 12, wherein the trigger variables further
comprise signals indicative of flow of the ambient air, flow of the
bleed air, pressure at the ambient air inlet, pressure at the
ambient air outlet, pressure at the bleed air inlet, pressure at
the bleed air outlet, outside air temperature, temperature at the
air inlet, temperature at the air outlet, temperature at the bleed
air inlet, temperature at the bleed air outlet, differential
pressure between the ambient air inlet and the ambient air outlet,
differential pressure between the bleed air inlet and the bleed air
outlet, weight on wheel off, altitude of the aircraft, and
time.
14. The method of claim 4, further comprising: closing the pressure
relief door based on a demand for decreased flow of the ambient
air.
15. The method of claim 1, wherein the pressure control mechanism
comprises a pressure relief door disposed proximate to an outlet of
the turbofan engine, downstream of the precooler, the method
further comprising: changing a degree of openness of the pressure
relief door in response to the detecting, thereby changing an
ambient air flow through the precooler.
16. The method of claim 15, wherein actuating the pressure control
mechanism includes opening the pressure relief door in the turbofan
engine at least in part to satisfy a demand for an increased flow
of the ambient air.
17. The method of claim 16, wherein actuating the pressure control
mechanism also includes closing the pressure relief door in the
turbofan engine at least in part to satisfy a demand for a
decreased flow of the ambient air.
18. An aircraft engine, comprising: a nacelle with an air inlet and
an air outlet; a cavity defined within the nacelle, the cavity
permitting ambient air to pass through the nacelle from the air
inlet to the air outlet; a compressor region defined within the
nacelle, the compressor region producing bleed air taken from at
least one position between the air inlet and the air outlet; a
precooler disposed within the nacelle, the precooler defining an
ambient air passage and a bleed air passage, wherein the ambient
air passes through the ambient air passage from an ambient air
inlet to an ambient air outlet, wherein the bleed air passes
through the bleed air passage from a bleed air inlet to a bleed air
outlet, and wherein heat is transferred between the ambient air and
the bleed air via heat exchange within the precooler; at least one
pressure relief door disposed within the cavity to create
additional engine venting proximate to the air outlet; and a
controller operatively connected to the pressure relief door,
wherein the controller at least opens the pressure relief door
based on a demand for increased flow of the ambient air through the
ambient air passage.
19. A method of operating an aircraft engine comprising a nacelle
with an air inlet and an air outlet, a cavity defined within the
nacelle, the cavity permitting ambient air to pass through the
nacelle from the air inlet to the air outlet, a compressor region
defined within the nacelle, the compressor region producing bleed
air taken from at least one position between the air inlet and the
air outlet, a precooler disposed within the nacelle, the precooler
defining an ambient air passage and a bleed air passage, wherein
the ambient air passes through the ambient air passage from an
ambient air inlet to an ambient air outlet, wherein the bleed air
passes through the bleed air passage from a bleed air inlet to a
bleed air outlet, and wherein heat is transferred between the
ambient air and the bleed air via heat exchange within the
precooler, at least one pressure relief door disposed within the
cavity to create additional engine venting proximate to the air
outlet, and a controller operatively connected to the pressure
relief door, wherein the controller at least opens the pressure
relief door based on a demand for increased flow of the ambient air
through the ambient air passage, the method comprising: receiving,
by the controller, trigger variables; evaluating the trigger
variables, by the controller, to determine if the ambient air
passing through the ambient air passage is cooling the bleed air
passing through the bleed air passage to a predetermined
temperature; and opening the pressure relief door, if the trigger
variables require an increased flow of the ambient air through the
ambient air passage.
20. A control system for controlling at least one pressure relief
door on an aircraft engine comprising a nacelle with an air inlet
and an air outlet, a cavity defined within the nacelle, the cavity
permitting ambient air to pass through the nacelle from the air
inlet to the air outlet, a compressor region defined within the
nacelle, producing bleed air taken from at least one position
between the air inlet and the air outlet, a precooler disposed
within the nacelle, the precooler defining an ambient air passage
and a bleed air passage, wherein the ambient air passes through the
ambient air passage from an ambient air inlet to an ambient air
outlet, wherein the bleed air passes through the bleed air passage
from a bleed air inlet to a bleed air outlet, wherein heat is
transferred between the ambient air and the bleed air via heat
exchange within the precooler, at least one pressure relief door
disposed within the cavity to create additional engine venting
proximate to the air outlet, and a controller operatively connected
to the pressure relief door, wherein the controller at least opens
the pressure relief door based on a demand for increased flow of
the ambient air through the ambient air passage, the system
comprising: a controller configured to receive trigger variables;
an actuator operatively connected to the at least one pressure
relief door, the actuator being configured at least to open the
pressure relief door; a communication line operatively connecting
the controller to the actuator; wherein, after evaluating the
trigger variables and determining if the ambient air passing
through the ambient air passage is cooling the bleed air passing
through the bleed air passage to a predetermined temperature, the
controller sends a signal to open the pressure relief door, if the
trigger variables require an increased flow of the ambient air
through the ambient air passage.
Description
CROSS-REFERENCE TO RELATED APPLICATION(S)
[0001] This is a Divisional patent application of U.S. patent
application Ser. No. 14/004,974, filed on Sep. 13, 2013, which is a
National Stage Entry patent application from PCT Patent Application
Serial No. PCT/US2012/029368, having an International filing date
of Mar. 16, 2012, and which relies for priority on U.S. Provisional
Patent Application Ser. No. 61/453,657, filed on Mar. 17, 2011, the
contents of all of which are incorporated herein by reference.
FIELD OF THE INVENTION
[0002] The present invention relates to a system and method for
operating a precooler in an aircraft. In particular, the present
invention encompasses an apparatus and method for temporarily
increasing the performance of a precooler in an aircraft.
DESCRIPTION OF RELATED ART
[0003] It is known to bleed hot, compressed air generated by an
aircraft engine and provide that compressed air to equipment on the
aircraft to perform certain onboard functions.
[0004] Specifically, it is known to siphon hot, compressed air
(also referred to as "bleed air") from an aircraft engine so that
the hot air may be used for aircraft functions outside of the
aircraft engine. For example, the bleed air may be used in an
aircraft's heating, ventilation, and air conditioning ("HVAC")
system, the aircraft's anti-icing system, and the aircraft's fuel
tank inerting system.
[0005] For an HVAC system installed in an aircraft with two
engines, bleed air typically is siphoned from each engine and is
sent to respective left and right side HVAC packs. The bleed air
may be mixed with recirculated air in the cockpit and passenger
cabin, where the bleed air conditions (i.e., heats) the cabin
temperature and pressurizes the aircraft's interior.
[0006] For the wing anti-icing system, the hot bleed air may be
used to heat areas of the aircraft which are prone to ice
accumulation, such as along a wing's leading edge.
[0007] With respect to the aircraft's fuel tank inerting system,
the bleed air may be used to reduce the oxygen content within the
aircraft's fuel tank(s), thereby minimizing the possibility of fuel
ignition within the fuel tank(s).
[0008] Depending upon the location where the bleed air is removed
from an engine, the bleed air may exit the engine at temperatures
up to 450.degree. C. or more. Specifically, bleed air taken from a
location near to the low pressure turbine may be at a temperature
of about 120.degree. C. Bleed air from a location near to the high
pressure turbine may be at a temperature of about 500.degree. C.
Since the temperature of the bleed air may be too hot to directly
circulate within the various systems of the aircraft, the hot bleed
air may have to be cooled prior to use with one or more of the
aircraft's other systems. As a result, it is known to provide a
cooling device, commonly referred to as a precooler, to cool the
hot bleed air down to temperature between about 200.degree. C. to
232.degree. C. depending on the usage. For engines such as turbofan
engines, which use a turbine driven fan to provide thrust, a
precooler is typically housed proximate to each engine, such as
within the engine's nacelle or above the engine's pylon.
[0009] A precooler typically utilizes outside (or ambient) that is
air drawn into by the fan to cool the hot, bleed air. Once inside
the nacelle, the ambient air may be between 70.degree. C., at low
altitude on a hot day, and -20.degree. C., at high altitude on a
cold day. The precooler typically includes a cross flow air-to-air
heat exchanger, which transfers heat energy from the streams of the
hot, bleed air to the stream of cold, ambient air, while the two
streams remain separated from one another. As should be apparent to
those skilled in the art, a stream of cooled, bleed air exits from
the precooler for use within the aircraft. Consequently, a stream
of heated, ambient air also exits from the precooler. The heated,
ambient air may be discharged into or around the engine nacelle or
outside the aircraft, for example, above the pylon
installation.
[0010] Precoolers are known in the art. Examples of known
precoolers are described in the following references: U.S. Patent
Application Publication No. 2008/0230651, U.S. Patent Application
Publication No. 2009/0000305, and U.S. Patent Application
Publication No. 2009/0007567. In addition, with respect to the art
of turbofan aircraft engines, the following references describe
engine precoolers: European Patent No. EP 0 953 506, U.S. Pat. No.
5,137,230, and U.S. Patent Application Publication No.
2009/0064658. A brief discussion of these references is provided
below.
[0011] U.S. Patent Application Publication No. 2008/0230651
(hereinafter "the '651 application") describes a turbofan provided
with a precooler. In a first embodiment, the precooler 18 is
positioned within an upper part 13S of a fan duct 13. (The '651
application at paragraph [0033].) In a second embodiment, the
precooler 31 includes a scoop 32 to capture some of the cold air
stream 9 (i.e., the ambient air) and carry it to the precooler 31.
(The '651 application at paragraph [0037].) In a third embodiment,
the precooler 38 is disposed in an upper part 12S of a chamber 12
surrounding the generator. (The '651 application at paragraph
[0040].)
[0012] U.S. Patent Application Publication No. 2009/0000305
(hereinafter "the '305 application") describes a dual flow turbine
engine equipped with a precooler. FIG. 1 in the '305 application
illustrates a prior art turbine engine, where the precooler 19 is
positioned within an upper part of the fan duct 13. (The '305
application at paragraph [0039].) FIG. 2 illustrates an embodiment
of the precooler 30, where the precooler 30 is not positioned
within the fan duct 13. (The '305 application at paragraph
[0042].)
[0013] U.S. Patent Application Publication No. 2009/0007567
(hereinafter "the '567 application") lists the same inventors as
the '305 application and includes drawings that are very similar to
the drawings provided in the '305 application. The '567 application
describes a dual flow turbine engine equipped with a precooler 30
positioned inside an intermediate chamber 12 of the engine, in
thermal contact with a rear part 10R of the fairing 10. (The '567
application at paragraph [0046].) An air intake 27 is positioned
upstream of the precooler 30 so that, when opened, a current of
cooling air 29 may be bled from the cold flow 9. (The '567
application at paragraph [0047].) The air intake 27 is provided
with a shut off means 28. (The '567 application at paragraph
[0047].)
[0014] European Patent No. EP 0 953 506 (hereinafter "the '506
patent") describes a flow control apparatus for gas turbine engine
installation pressure relief doors. Specifically, the '506 patent
describes pressure relief doors 12 that are operable during an
engine failure event. (The '506 patent at paragraph [0011].) The
relief doors 12 include side walls 42 that alter the flow of the
hot air passing through the relief doors 12. (The '506 patent at
paragraph [0011].) By altering the flow of the hot air, hot engine
gases do not stay in contact with the engine shroud 16 for as long
a period of time as with prior art designs, thereby minimizing the
possibility of damage to the engine. (The '506 patent at paragraph
[0011].) One type of engine failure mentioned in the '506 patent is
a bleed duct failure. (The '506 patent at paragraph [0004].)
[0015] U.S. Pat. No. 5,137,230 (hereinafter "the '230 patent")
describes an aircraft turbine engine bleed air recovery apparatus.
In particular, the '230 patent describes an environmental control
system ("ECS") with a boundary layer bleed compressor 64, driven by
a turbine 63, that draws boundary layer suction air into a
precooler heat exchanger 69. (The '230 patent at col. 6, lines
36-41.) The precooler heat exchanger 69 cools the bleed air, which
is used for the aircraft ECS. (The '230 patent at col. 6, lines
36-41.) In cases where the demand for bleed air exceeds the
capacity of the nozzle 65 or the turbine 63, bleed air may bypass
the turbine 63 via a bypass valve 72. (The '230 patent at col. 6,
lines 55-59.) A controller 82 controls the flow rate of bleed air
to the ECS pack 80. (The '230 patent at col. 7, lines 3-14.)
[0016] U.S. Patent Application Publication No. 2009/0064658
(hereinafter "the '658 application") describes a bypass
turbomachine capable of reducing jet noise. The turbomachine 10 is
designed to provide reverse thrust via the operation of covers 30
and openings that redirect the flow of the engine's output thrust.
(The '658 application at paragraph [0038].)
[0017] With respect to precoolers, it has been found that, under
certain operating conditions, the fan in the engine may not be able
to provide enough cold air pressure to the precooler to enable all
the onboard functions desired for the aircraft.
[0018] In the case on an engine failure or an engine bleed system
failure, it is contemplated that the remaining engine(s) and
respective precooler(s) will have a sufficient load capacity to
supply enough cooled bleed air to the HVAC system to make up for
output from the lost engine and precooler.
[0019] However, if one engine fails or one engine bleed system
fails and the aircraft enters an ice condition, it is possible that
the remaining precooler(s) may not be able to cool enough hot bleed
air to supply both the HVAC system and the anti-icing system at the
same time. In such rare circumstances, it may be necessary for the
aircraft to change course to avoid the ice condition, thereby
reducing the demand on the remaining precooler(s). Naturally, where
an aircraft changes course, the flight may be disrupted. Moreover,
the aircraft may be prevented from reaching its destination and may
be required to land for maintenance.
[0020] As should be apparent to those skilled in the art, this
problem may be addressed by providing a larger outlet to the
chamber or channel into which the precooler discharges, i.e., a
larger core ventilation outlet area. However, a larger outlet may
negatively affect aircraft performance, even during normal system
operation, since core exhaust recovery is reduced.
[0021] This problem also may be addressed by increasing the size of
the precooler. A larger precooler, however, will add mass to the
aircraft, will likely increase installation difficulty, and may
further disrupt the flow of air within or around the engine.
[0022] Thus, there is a need for a system which, by virtue of its
design and components, is able to overcome one or more of the
above-discussed problems with the prior art.
[0023] Additionally, there is a need for a system or method for
operating a precooler in an aircraft that temporarily improves the
performance of the precooler.
[0024] Moreover, there is a need for a system or method for
operating a precooler that enables operation of one or both the
aircraft's HVAC system and the aircraft's anti-icing system when
there is an engine failure or an engine bleed system failure,
without unduly affecting the aircraft's performance, during normal
operating conditions.
SUMMARY OF THE INVENTION
[0025] It is, therefore, one aspect of the present invention to
address at least one of the difficulties identified with respect to
the prior art.
[0026] Accordingly, the present invention provides a method for
controlling a pressure control mechanism in a turbofan engine
having a precooler permitting heat exchange between ambient air and
bleed air. The method includes detecting at least one of an engine
failure or a bleed system failure, detecting at least one of an ice
condition or an activation of an anti-ice system, and actuating the
pressure control mechanism, thereby altering the heat exchange
between the ambient air and the bleed air.
[0027] In one contemplated embodiment, the pressure control
mechanism includes a pressure relief door disposed proximate to an
outlet of the turbofan engine.
[0028] In another contemplated embodiment, the pressure control
mechanism is actuated upon detection of a combination of at least
one of an engine failure and a bleed system failure, and at least
one of an ice condition and an activation of the anti-ice
system.
[0029] Still further, it is contemplated that actuating the
pressure control mechanism includes opening a pressure relief door
in the turbofan engine.
[0030] It is also contemplated that the pressure relief door
includes a plurality of pressure relief doors, and that actuating
the plurality of pressure relief doors includes opening the
plurality of pressure relief doors.
[0031] In one further contemplated embodiment, the bleed air has a
temperature greater than the ambient air.
[0032] The present invention also contemplates that the bleed air
is cooled to a predetermined temperature within the precooler via
the heat exchange with the ambient air.
[0033] It is contemplated that the predetermined temperature is
between about 200.degree. C. to 232.degree. C.
[0034] The method of the present invention also may include
providing an ambient air passage for the ambient air, providing a
bleed air passage for the bleed air, and separating the ambient air
passage and the bleed air passage from one another such that the
ambient air and the bleed air are prevented from intermixing during
the heat exchange.
[0035] In another variation, the method may include providing an
ambient air passage for the ambient air, providing a bleed air
passage for the bleed air, and combining the ambient air passage
and the bleed air passage to permit intermixing of the ambient air
and the bleed air during the heat exchange.
[0036] It is contemplated that the bleed air may be taken from
positions proximate to both a high pressure compressor and a low
pressure compressor in the turbofan engine.
[0037] It is also contemplated that a demand for increased flow of
the ambient air is conditioned upon trigger variables, which
comprise signals indicative of an engine failure, a bleed air
system failure, an ice condition, or an activation of an anti-icing
system. The trigger variables may include signals indicative of
flow of the ambient air, flow of the bleed air, pressure at the
ambient air inlet, pressure at the ambient air outlet, pressure at
the bleed air inlet, pressure at the bleed air outlet, outside air
temperature, temperature at the air inlet, temperature at the air
outlet, temperature at the bleed air inlet, temperature at the
bleed air outlet, differential pressure between the ambient air
inlet and the ambient air outlet, differential pressure between the
bleed air inlet and the bleed air outlet, weight on wheel off,
altitude of the aircraft, and time.
[0038] In another embodiment, the method may include closing the
pressure relief door based on a demand for decreased flow of the
ambient air.
[0039] In the method of the present invention, where the pressure
control mechanism has a pressure relief door disposed proximate to
an outlet of the turbofan engine, downstream of the precooler, the
method may include changing a degree of openness of the pressure
relief door in response to the detecting, thereby changing an
ambient air flow through the precooler.
[0040] In this regard, actuating the pressure control mechanism may
include opening the pressure relief door in the turbofan engine at
least in part to satisfy a demand for an increased flow of the
ambient air.
[0041] Still further, actuating the pressure control mechanism also
may include closing the pressure relief door in the turbofan engine
at least in part to satisfy a demand for a decreased flow of the
ambient air.
[0042] The present invention also provides an aircraft engine. The
aircraft engine includes a nacelle with an air inlet and an air
outlet, a cavity defined within the nacelle, the cavity permitting
ambient air to pass through the nacelle from the air inlet to the
air outlet, a compressor region defined within the nacelle, the
compressor region producing bleed air taken from at least one
position between the air inlet and the air outlet, a precooler
disposed within the nacelle, the precooler defining an ambient air
passage and a bleed air passage, wherein the ambient air passes
through the ambient air passage from an ambient air inlet to an
ambient air outlet, wherein the bleed air passes through the bleed
air passage from a bleed air inlet to a bleed air outlet, and
wherein heat is transferred between the ambient air and the bleed
air via heat exchange within the precooler, at least one pressure
relief door disposed within the cavity to create additional engine
venting proximate to the air outlet, and a controller operatively
connected to the pressure relief door, wherein the controller at
least opens the pressure relief door based on a demand for
increased flow of the ambient air through the ambient air
passage.
[0043] The present invention also provides for a method of
operating an aircraft engine with a nacelle with an air inlet and
an air outlet, a cavity defined within the nacelle, the cavity
permitting ambient air to pass through the nacelle from the air
inlet to the air outlet, a compressor region defined within the
nacelle, the compressor region producing bleed air taken from at
least one position between the air inlet and the air outlet, a
precooler disposed within the nacelle, the precooler defining an
ambient air passage and a bleed air passage, wherein the ambient
air passes through the ambient air passage from an ambient air
inlet to an ambient air outlet, wherein the bleed air passes
through the bleed air passage from a bleed air inlet to a bleed air
outlet, and wherein heat is transferred between the ambient air and
the bleed air via heat exchange within the precooler, at least one
pressure relief door disposed within the cavity to create
additional engine venting proximate to the air outlet, and a
controller operatively connected to the pressure relief door,
wherein the controller at least opens the pressure relief door
based on a demand for increased flow of the ambient air through the
ambient air passage. The method includes receiving, by the
controller, trigger variables, evaluating the trigger variables, by
the controller, to determine if the ambient air passing through the
ambient air passage is cooling the bleed air passing through the
bleed air passage to a predetermined temperature, and opening the
pressure relief door, if the trigger variables require an increased
flow of the ambient air through the ambient air passage.
[0044] The present invention also provides for a control system for
controlling at least one pressure relief door on an aircraft engine
with a nacelle with an air inlet and an air outlet, a cavity
defined within the nacelle, the cavity permitting ambient air to
pass through the nacelle from the air inlet to the air outlet, a
compressor region defined within the nacelle, producing bleed air
taken from at least one position between the air inlet and the air
outlet, a precooler disposed within the nacelle, the precooler
defining an ambient air passage and a bleed air passage, wherein
the ambient air passes through the ambient air passage from an
ambient air inlet to an ambient air outlet, wherein the bleed air
passes through the bleed air passage from a bleed air inlet to a
bleed air outlet, wherein heat is transferred between the ambient
air and the bleed air via heat exchange within the precooler, at
least one pressure relief door disposed within the cavity to create
additional engine venting proximate to the air outlet, and a
controller operatively connected to the pressure relief door,
wherein the controller at least opens the pressure relief door
based on a demand for increased flow of the ambient air through the
ambient air passage. The system includes a controller configured to
receive trigger variables, an actuator operatively connected to the
at least one pressure relief door, the actuator being configured at
least to open the pressure relief door, and a communication line
operatively connecting the controller to the actuator, where, after
evaluating the trigger variables and determining if the ambient air
passing through the ambient air passage is cooling the bleed air
passing through the bleed air passage to a predetermined
temperature, the controller sends a signal to open the pressure
relief door, if the trigger variables require an increased flow of
the ambient air through the ambient air passage.
[0045] Still other aspects of the present invention will be made
apparent from the discussion provided herein.
BRIEF DESCRIPTION OF THE DRAWINGS
[0046] The present invention will now be described in connection
with the following figures, in which:
[0047] FIG. 1 is a schematic cross-sectional side view of a turbine
engine and system according to an embodiment of the present
invention;
[0048] FIG. 2 is an enlarged section of a portion of the engine
illustrated in FIG. 1;
[0049] FIG. 3 is a cross-sectional front view of the engine
illustrated in FIG. 1, the cross-section being taken along line
III-III in FIG. 1;
[0050] FIG. 4 is a graphical, top view of an aircraft embodying one
contemplated embodiment of the system of the present invention;
[0051] FIG. 5 is a graphical, top view of an aircraft embodying
another contemplated embodiment of the system of the present
invention;
[0052] FIG. 6 is a flow chart illustrating steps contemplated for
one method of increasing performance of a precooler according to
one embodiment of the present invention; and
[0053] FIG. 7 is a flow chart illustrating steps contemplated for
another method of increasing performance of a precooler according
to another embodiment of the present invention.
DETAILED DESCRIPTION OF EMBODIMENT(S) OF THE INVENTION
[0054] While the invention will be described in conjunction with
specific embodiments. It should be understood that the discussion
of any one, particular embodiment is not intended to be limiting of
the scope of the present invention. To the contrary, the specific,
enumerated embodiments are intended to illustrate a wide variety of
alternatives, modifications, and equivalents that should be
apparent to those of ordinary skill in the art. The present
invention is intended to encompass any such alternatives,
modifications, and equivalents as if discussed herein.
[0055] In the following description, the same numerical references
are intended to refer to similar elements. The re-use of reference
numerals for different embodiments of the present invention is
intended to simplify the discussion of the present invention. It
should not be inferred, therefore, that the re-use of reference
numbers is intended to convey that the associated structure is
identical to any other described embodiment.
[0056] Although the preferred embodiments of the present invention
as illustrated in the accompanying drawings comprise various
components, and although the preferred embodiments of the system
and corresponding parts of the present invention as shown consist
of certain geometrical configurations as explained and illustrated
herein, not all of these components and geometries are essential to
the invention and, thus, should not be taken in their restrictive
sense, i.e., should not be taken as to limit the scope of the
present invention.
[0057] It is to be understood, as should be apparent to a person
skilled in the art, that other suitable components and cooperations
therebetween, as well as other suitable geometrical configurations
may be used for a system according to the present invention, as
will be briefly explained herein and as may be easily inferred
herefrom by a person skilled in the art, without departing from the
scope of the invention.
[0058] Additionally, it should be appreciated that positional
descriptions such as "front," "rear," and the like are, unless
otherwise indicated, to be taken in the context of the figures and
should not be considered limiting of the present invention.
[0059] It will be appreciated that the present invention may be
practiced without the specific details which have been set forth
herein below in order to provide a thorough understanding of the
invention.
[0060] The method of the present invention is described herein as a
series of steps. It will be understood that this method and the
associated steps may be performed in any logical order. Moreover,
the method may be performed alone, or in conjunction with other
procedures and methods before, during or after such methods and
steps set forth herein without departing from the scope of the
present invention.
[0061] FIG. 1 shows a schematic representation of a turbine engine
10 housed within a nacelle 12. The nacelle 12 has a tubular shape
and is suspended from a wing 14 by a pylon 16. The nacelle 12 forms
an annular cavity 18 around the turbine engine 10 through which air
can flow. In use, the turbine engine 10 drives a fan 20 positioned
in front of the turbine engine 10, which draws ambient air 25 into
the cavity 18 through a front air inlet 22. This fan air (the
ambient air 25) passes through the nacelle 12 and out of a rear air
outlet 24. As will be discussed in further detail herein below,
some of the ambient air 25 is mixed with fuel and combusted within
the turbine engine 10 and some of the ambient air 25 is drawn past
the turbine engine 10 to provide thrust. This arrangement is known
as a turbofan. In addition, some of the ambient air 25 can also be
used to ventilate the turbine engine 10, as will be discussed
below.
[0062] As is known in the art, the cold air (i.e., ambient air) 25
enters the turbine engine 10 via a turbine inlet 26 and passes
through a low pressure compressor 28 followed by a high pressure
compressor 30. The ambient air 25 is then mixed with fuel and
ignited in a combustion chamber 32. The combustion gasses 39 pass
through high and low pressure turbines 34 and 36, causing rotation.
The high and low pressure turbines 34, 36 are connected to the fan
20 and, due to their rotation, drive the fan 20. The combustion
gases 39 exit from the high and low pressure turbines 34, 36 and
exit through a nozzle 38 and a turbine outlet 40.
[0063] As illustrated in FIG. 1, the compressors 28 and 30,
combustion chamber 32, and turbines 34 and 36 are enclosed by an
engine casing 41, which forms the outer wall of the turbine engine
10.
[0064] The extraction and conditioning of bleed air 42 will now be
now discussed in connection with the enlarged section of the
turbine engine 10 illustrated in FIG. 2.
[0065] With reference to FIG. 2, a stream of hot, compressed air 42
is bled from the turbine engine 10. This hot bleed air 42, which is
preferably drawn from both the low and high pressure compressors 28
and 30 via conduits 44 and 46, respectively, is sent to the hot air
inlet 48 of a precooler 50.
[0066] As discussed above, the precooler 50 serves to cool the
stream of hot bleed air 42, producing a stream of cooled hot bleed
air 52 which can be used, inter alia, in an aircraft's HVAC system
and anti-icing system. It is also known to send the stream of the
cooled bleed air 52 to the engine starter valve of the opposite
engine. The precooler 50 is integrated into the aircraft by a
system 51 which, as will be discussed in further detail below, may
be advantageously used to improve its performance.
[0067] A stream of cold air 54, taken from the ambient air 25 drawn
in by the fan 20, is received at a cold air inlet 56. The streams
of cold air 54 and hot bleed air 42 pass through a heat exchanger
58 which, as is known in the art, allows heat energy from the
stream of hot bleed air 42 to be transferred to the stream of cold
air 54, thereby cooling the former and heating the latter.
[0068] The heat exchanger 58 may be embodied in a variety of ways.
For example, the hot bleed air 42 may be fed through a plurality of
tubes or fins made of a conductive material. The cold air 54, as it
passes through the precooler 50, flows past the tubes or fins,
thereby allowing the heat transfer to occur. In this way, the cold
air 54 is heated to produce heated cold air 64. Similarly, the hot
bleed air 42 is cooled to produce the cooled, hot bleed air 52.
[0069] To simplify the discussion that follows, and for purposes of
consistency, the following convention is employed with respect to
the two air streams entering and exiting the precooler 50. The
cold, ambient air 54 entering the precooler 50 is referred to as
the "ambient input air 54." The heated, cold air 64 exiting from
the precooler 50 is referred to as the "ambient output air 64."
Similarly, the hot bleed air 42 entering the precooler 50 is
referred to as the "hot bleed input air 42." The cooled, hot bleed
air 52 exiting from the precooler 50 is referred to as the "hot
bleed output air 52." In addition, the cold air 25 drawn into the
engine 10 by the fan 20 is referred to generally as the "ambient
air 25." The bleed air 42 extracted from the engine 10 is referred
to generally as the "bleed air 42."
[0070] As should be apparent from the foregoing discussion, and as
should be apparent from the discussion that follows, the ambient
air 25 may or may not intermix with the bleed air 42 in the
precooler 50. In the illustrated embodiment, the ambient air 25 and
the bleed air 42 flow through separate pathways and do not mix with
one another. However, it is possible that the two streams 25, 42
may be mixed, partially or wholly, within the precooler 50 without
departing from the scope of the present invention. In other words,
the exact construction of the heat exchange portion of the
precooler 50 is not critical for operation of the present
invention.
[0071] The stream of hot bleed output air 52 exits the precooler 50
from a cooled hot air outlet 60, which is connected to the
appropriate aircraft systems via a conduit 62. A stream of ambient
output air 64 is discharged from a heated cold air outlet 66 into
the cavity 18.
[0072] In the illustrated embodiment, a scoop 68 is provided at the
cold air inlet 56 so as to increase the volume and/or pressure of
the ambient air 25 captured and fed into the precooler 50. As will
be appreciated by those skilled in the art, the scoop 68 may be
designed to maximize inlet pressure and minimize the pressure loss
of the flow of ambient air 25 therein. A fan air valve 70 for
regulating the stream of the ambient input air 54, and thereby the
temperature of the hot bleed output air 52, is also provided
between the scoop 68 and the cold air inlet 56. The size and
location of the fan air valve 70 is also chosen to maximize inlet
pressure and minimize the pressure drop.
[0073] In FIG. 2, the valve 70 is illustrated as a butterfly valve.
As should be apparent to those skilled in the art, other
constructions for the valve 70 may be employed without departing
from the scope of the present invention. In other words, the
construction of the valve 70 is not critical to the present
invention.
[0074] The majority of the ambient air 25 drawn into the fan 20
passes through the cavity 18, i.e., around the turbine engine 10
and the precooler 50, for propulsion. Only a small amount,
typically no more than 1%, is fed into the precooler 50. It is
noted, however, that a larger or smaller percentage of ambient air
25 may enter the precooler 50 without departing from the scope of
the present invention. As such, the present invention is not
limited to a construction that extracts less than 1% of the ambient
air 25 from the total air stream.
[0075] Returning to FIG. 1, the nacelle 12 preferably includes two
distinct structures: an engine cowl 70 and an inner fixed structure
76. The cowl 70 defines inner fairings 72 and outer fairings 74.
The cowl 70 also forms the outermost housing for the engine 10 and
the precooler 50. The front air inlet 22 is bound by the inner
fairing 72, while the cavity 18 extends around the turbine engine
10, between the inner fairing 72 and the engine casing 41.
[0076] The inner fixed structure 76 is a thin member that extends
concentrically between the cowl 70 and the engine casing 41. This
inner fixed structure 76 effectively divides the cavity 18 into two
parts or segments.
[0077] The first part or segment defined by the inner fixed
structure 76 is an engine fan bypass channel 78. The engine bypass
channel 78 extends radially between the engine cowl 70 and the
inner fixed structure 76, and longitudinally between a front bypass
inlet 79 and a rear bypass outlet 81. It is the ambient air 25
drawn by the fan 20 into and through the fan bypass channel 78 that
generates thrust.
[0078] The second part or segment defined by the inner fixed
structure 76 is an additional channel, called the engine core
compartment 80. The engine core compartment 80 may be used to
ventilate the engine 10. In addition, various engine accessories
may be installed within the engine core compartment 80. The engine
core compartment 80 extends radially between the engine casing 41
and the inner fixed structure 76 and longitudinally between a front
ventilation inlet 82 and a rear ventilation outlet 84. As ambient
air 25 flows into the inlet 82, across the engine casing 41 and out
of the outlet 84, the ambient air 25 may be used to cool the engine
10 and the combustion chamber therewithin without affecting the
rest of the flow through the fan bypass channel 78.
[0079] As is known in the art, and as shown in FIG. 1, the rear
portion of the inner fixed structure 76, that is to say a portion
proximate the nozzle 38, extends beyond the engine cowl 70.
[0080] The precooler 50 may be disposed within the fan bypass
channel 78 or the engine core compartment 80, as both receive
ambient air 25 drawn in by the fan 20. As mentioned above, the
precooler 50 alternatively may be located within or above the pylon
16 without departing from the scope of the present invention. In
the illustrated embodiment, the precooler 50 is disposed wholly
within the engine core compartment 80.
[0081] With reference to FIG. 3, it is noted that the fan bypass
channel 78 and the engine core compartment 80 are not completely
concentric. Rather, the fan bypass channel 78 forms an annular
segment that is discontinuous at the location where it is
intersected by an upwardly projecting extension 80a of the engine
core compartment 80. The inner fixed structure 76 includes two
vertically extending sections 86 which intersect the inner fairing
72, thereby forming the extension 80a above the engine 10.
[0082] It should be appreciated, however, that the radial location
of the extension 80a is not critical to operation of the present
invention. The extension 80a may project downward from the engine
10, to the side of the engine 10, or at an angle from the engine 10
(between the upward and downward positions) without departing from
the scope of the present invention.
[0083] As illustrated, the precooler 50 is positioned within the
extension 80a such that both the cold air inlet 56 and the heated
cold air outlet 66 are in communication with the ambient air 25
flowing through the engine core compartment 80. The stream of
ambient output air 64 is thereby discharged into the engine core
compartment 80 prior to being exhausted at the rear ventilation
outlet 84. In this manner, the propulsive air driven through the
fan bypass channel 78 is isolated from the precooler 50 and the
stream of ambient output air 64 exhausted therefrom.
[0084] It should be appreciated that the pressure within the fan
bypass channel 78 may be higher than that in the engine core
compartment 80, which results in a smaller pressure difference
across the precooler 50 positioned within the fan bypass channel
78. Such an arrangement is feasible if the stream of ambient output
air 64 is discharged far to the rear of the engine cowl 70.
[0085] In use, the ambient air 25 is slightly compressed by the fan
20 as it enters the nacelle 12. Preferably, the pressure of the
ambient input air 54 entering the precooler 50 is maximized by the
scoop 68. It will be appreciated that the difference between the
air pressure at the cold air inlet 56 and the back pressure within
the engine core compartment 80 proximate to the heated cold air
outlet 66 will, in large part, determine the performance of the
precooler 50.
[0086] As discussed above, it is contemplated that a typical
commercial aircraft will include two turbofan engines 10, each of
which has its own precooler 50. In normal operation, each engine 10
and precooler 50 pair supplies half of the hot bleed output air 52
to the HVAC system on the aircraft.
[0087] FIG. 4 provides a graphical representation of a top view of
one embodiment of an aircraft 110 and selected features for that
aircraft 110. In the illustrated example, the aircraft 110 includes
a fuselage 112, two wings 14, two engines 10, and two precoolers
50. Hot bleed output air 52 from each of the precoolers 50 passes
through the conduits 62 to several components requiring the hot
bleed output air 52. For example, some of the hot bleed output air
52 passes through a conduit 114, 116 from one engine 10 to the
other engine 10. As noted above, this hot bleed output air 52 is
inputted into the engine starter valve (not shown) for the opposing
engine 10. Some of the hot bleed output air 52 is directed to the
two HVAC systems 118, 120 (also referred to in the singular as the
"HVAC system") on board the aircraft 110, via conduits 122, 124. As
noted above, the HVAC systems 118, 120 use the hot bleed output air
52 to condition the cabin air by pressurizing and heating the cabin
air, for example. In addition, the hot bleed output air 52 may pass
through conduits 126, 128 to respective anti-icing systems 130,
132. The anti-icing systems 130, 132 utilize the hot bleed output
air 52 to heat selected parts of the aircraft 110, i.e., the
leading edges 134, 136 of the wings 14 that may be prone to icing
during flight.
[0088] FIG. 5 provides a graphical top view of a second embodiment
of an aircraft 110. In this embodiment, the aircraft 110 does not
include a conduit 116. Instead, the engines 10 share the same
conduit 114. In this second embodiment, elimination of the conduit
116 simplifies the aircraft 110 and reduces the weight of the
aircraft 110, which is advantageous for many reasons, as should be
apparent to those skilled in the art.
[0089] While not illustrated for the two illustrated variants of
the aircraft 110, the various conduits 114, 116, 122, 124, 126, 128
may include one or more valves and sensors that help to control the
flow of the hot bleed output air 52 to the various components. In
addition, while the HVAC systems 118, 120 and anti-icing systems
130, 132 are shown, the hot bleed output air 52 may be provided to
any number of different systems without departing from the scope of
the present invention.
[0090] It is noted that the various conduits 114, 116, 122, 124,
126, 128 illustrated for the aircraft 110 are exemplary only. A
larger or a fewer number of conduits may be employed without
departing from the scope of the present invention. Moreover, the
conduits 114, 116, 122, 124, 126, 128 may be arranged in any
configuration suitable for the aircraft 110 without departing from
the scope of the present invention.
[0091] When the aircraft 110 is operating under conditions without
ice formation (or "icing conditions"), the hot bleed output air 52
typically is provided to the HVAC systems 118, 120 to pressurize
and heat (or cool) the air within the cabin. Under these
conditions, there is a first demand placed upon the precoolers 50
to provide a sufficient quantity of hot bleed output air 52 for
operation of the HVAC systems 118, 120.
[0092] When the aircraft 110 enters an ice (or "icing") condition,
the anti-icing systems 130, 132 are activated. Since the anti-icing
systems 130, 132 also require hot bleed output air 52 as a heating
input, when the anti-icing systems 130, 132 are activated, the need
for the additional hot bleed output air 52 places a second demand
on the precoolers 50. In other words, the second demand on the
precoolers 50 adds to the overall demand placed on these
components. Each precooler 50 is, therefore, required to generate
additional hot bleed output air 52 to meet the increased
demand.
[0093] As should be apparent to those skilled in the art, the
precoolers 50 are sized to meet the demand expected during normal
operating conditions for the aircraft 110. Despite being sized
appropriately, it is foreseeable that certain conditions may arise
that are outside of the "normal" operation of the aircraft 110.
These abnormal conditions, which may occur statistically in one out
of 10,000 flights or less, also must be taken into account when
sizing the precoolers 50 and the systems that the precoolers 50
support.
[0094] In certain abnormal operating conditions, it is possible
that the precooler 50 may not supply a sufficient amount of hot
bleed output air 52 for the aircraft 110. In particular, it is
possible that, in the event of an engine failure or engine bleed
failure (among other types of operational events), when a single
engine 10 or single bleed system is called upon to make up for any
lost capacity due to the inoperability of the failed engine or
bleed system, the remaining precooler 50 cannot generate sufficient
hot bleed output air 52. As a result, it is possible the air
conditioning systems 118, 120 and/or the anti-icing systems 130,
132 may not operate properly due to a lack of sufficient hot bleed
output air 52. Specifically, the remaining precooler 50 may not be
able to cope with the increased demand for hot bleed output air
52.
[0095] It also should be appreciated that there are other
operational circumstances that may arise during use of the aircraft
110 which require an above-normal supply of hot bleed output air 52
from a given engine 10 and precooler 50. As noted above, for
example, when the HVAC systems 118, 120 and the anti-icing systems
130, 132 are operational at the same time, there is a considerable
demand for hot bleed output air 52. If other systems are activated
that also require hot bleed output air 52 (or if there is an engine
failure or hot bleed system failure), the demand may exceed, at
least momentarily, the ability of one or more of the precoolers 50
to meet the demand.
[0096] The system 51 is provided in order to improve the
performance of the precooler 50 in situations which require the
precooler 50 to cool higher than normal volumes of hot bleed air
42, under circumstances such as those described above. To
accomplish this, a pressure control mechanism 88 is provided along
the nacelle 12. The pressure control mechanism 88 is provided to
temporarily increase the pressure difference across the precooler
50 from the cold air inlet 56 to the heated cold air outlet 66,
thereby increasing both the flow of the ambient air 25 through the
precooler 50. The pressure control mechanism 88, therefore,
increases the ability of the precooler 50 to cool the hot bleed
input air 42.
[0097] In the embodiment illustrated in FIG. 1, the pressure
control mechanism 88 includes a pressure relief door 90, which is
positioned along the inner fixed structure 76 between the precooler
50 and the nozzle 38. Preferably, the pressure relief door 90 is
positioned along the rear section of the inner fixed structure 76,
beyond the rear bypass outlet 81.
[0098] The pressure relief door 90 is mounted to the inner fixed
structure 76, preferably by a pivotable connection 92, disposed
along a front edge of the pressure relief door 90. In this
arrangement, the pressure relief door 90 opens at its rear, as
shown in FIG. 1.
[0099] It is noted that the pressure relief door 90 need not be
positioned exactly where indicated in FIG. 1. The pressure relief
door 90 may be located elsewhere without departing from the scope
of the present invention. In addition, while one pressure relief
door 90 is illustrated, it is contemplated that several pressure
relief doors 90 may be employed without departing from the scope of
the present invention.
[0100] It is noted that there are several mechanisms that may be
employed to open and/or close the pressure relief door 90, upon the
occurrence of circumstances that require opening of the pressure
relief door 90.
[0101] In one contemplated embodiment, the pressure relief door 90
is designed only to open. In this contemplated embodiment, if the
pressure relief door 90 opens during flight, the pressure relief
door will remain opened until the aircraft 110 lands. After
landing, the pressure relief door 90 may be closed by trained
personnel, such as an aircraft mechanic. In this embodiment, the
pressure relief door 90 may be biased to open using a solenoid, a
spring, or other suitable mechanism capable of opening the pressure
relief door 90 and keeping the pressure relief door 90 opened until
closed by trained personnel.
[0102] In a second contemplated embodiment, the pressure relief
door 90 may be opened and closed during flight by a suitable
actuator. A solenoid, motor, or other device may be connected to
the pressure relief door 90 so that the pressure relief door 90 may
be opened when additional hot bleed output air 52 is needed. The
pressure relief door 90 may then be closed when the demand for hot
bleed output air 52 returns to a normal level. The actuator may be
controlled by a switch, which may be electrically or pneumatically
activated, as required or as desired.
[0103] As noted above, under normal operating conditions, such as
when all of the engines 10 and the precoolers 50 are operating
within normal parameters, the pressure relief door 90 remains
closed. As discussed above, the air pressure at the cold air inlet
56 will be above the ambient pressure outside the nacelle 12 due to
the fan 20 and, preferably, the scoop 68. The pressure within the
engine core compartment 80 at the heated cold air outlet 66 will be
below the pressure at the inlet 56, but will generally be slightly
above that of the ambient outside air 25. When an increased
capacity of the precooler 50 is required, the pressure relief door
90 can be opened so as to create additional venting area at the
rear of the inner fixed structure 76.
[0104] The additional area increases the flow of air through the
precooler 50, which allows a greater volume of hot bleed input air
42 to be processed into hot bleed output air 52. The additional
area created when the pressure relief door is opened allows the
pressure within the core engine compartment 80 to be released to
the atmosphere. In other words, the resultant drop in pressure at
the heated cold air outlet 66 increases the pressure difference
across the precooler 50 and increases the flow rate of the ambient
air 25 passing therethrough.
[0105] It will be appreciated however that opening the doors 90 and
lowering the pressure at the rear of the engine core compartment
80, while improving the performance of the precooler 50, may
negatively affect other operational parameters of the engine 10 by,
for example, increasing drag and/or fuel consumption. Accordingly,
the pressure relief door 90 is preferably only opened temporarily
when maximum performance of the precooler 50 is required.
[0106] A control system 200 is preferably provided to control
opening and closing of the door 90. The control system 200 responds
to an increased demand on the precooler 50. For example, the
control system 200 responds to a failure of an engine, a failure of
a hot air bleed system, and/or activation and/or initiation of the
anti-icing systems 130, 132, among other variables.
[0107] As illustrated in FIG. 1, the control system 200 depicted
graphically as a controller 202 connected, via a communication line
204, to an actuator 206 that connects to the pressure relief door
90. It is contemplated that the controller 202 is configured to
send an appropriate signal to the actuator 206 to open or close the
pressure relief door 90. The actuator may be spring-operated,
electrically motorized, hydraulically triggered, or pneumatically
operated, as should be apparent to those skilled in the art. In
other words, the operative mechanism of the actuator 206 is not
critical to operation of the control system 200. While the control
system 200 is illustrated in this simplistic manner, it should be
understood that this illustration is not intended to be limiting of
the present invention, as should become apparent from the
discussion herein.
[0108] The present invention also provides a method of increasing
performance of the precooler 50.
[0109] With reference to FIG. 6, a pressure control mechanism 88
(which includes a suitable processor) may operate as follows.
First, the mechanism 88 may detect, at step 100, when one of two or
more engines 10 fails or when one of two or more bleed systems
fail, thereby disabling the corresponding precooler 50 and
increasing the amount of hot bleed air siphoned from the remaining
engine(s) 10. Second, the mechanism 88 may detect, at step 102,
when the aircraft enters an ice condition and therefore requires
activation of the anti-icing systems 130, 132. In this regard, the
mechanism may detect when ice protection is activated on the
aircraft 110, which also requires additional hot bleed output air
52. Third, the mechanism 88 may actuate 104 the pressure relief
doors 90 in response to the engine failure and/or ice condition,
thereby lowering the air pressure at the heated cold air outlet
66.
[0110] It will be appreciated that various parameters can be
monitored in order to detect 102 the need to boost precooler
performance. As noted above, basic input parameters for the control
system 200 include receipt of signals from various component on the
aircraft 110 that monitor if an engine failure has occurred, if
there has been a failure of the bleed air system for one of the
engines 10, if there is an ice condition being experienced by the
aircraft 110, if the anti-ice systems 130, 132 have been activated,
or if there is an increased demand for the production of hot bleed
output air 52. Preferably, the system 51 additionally monitors
bleed system performance between the detecting step 102 and the
actuating step 104.
[0111] It will therefore be appreciated that an aircraft which
includes a precooler 50 and system 51 in accordance with the
present invention may include modifications to the design of the
front ventilation inlet 82, the rear ventilation outlet 84 and the
rest of the inner fixed structure 76 to ensure optimal flow through
the engine core compartment 80 during normal operation, thereby
improving the efficiency and performance of the engine 10.
[0112] It will also be appreciated that the present invention
should not be considered limited to the preferred embodiment shown
in the figures. As noted above, although only one pressure relief
door 90 is shown, it may be desirable to provide a plurality of
pressure relief doors 90 distributed circumferentially around
and/or longitudinally along the internal fixed structure 76. In
addition, the pressure relief door 90, which is shown
schematically, may be shaped and/or sized in different ways. For
example, rather than a pivotally opening door, as illustrated in
FIG. 1, the pressure relief mechanism 88 could be embodied by a
sliding panel, sliding louvers or another means of increasing the
exhaust area associated with the pressure relief door 90.
[0113] It will further be appreciated that, for a precooler 50
whose heated cold air outlet 66 discharges to the fan bypass
channel 78 rather than the engine core compartment 80 as
illustrated, the pressure relief mechanism 88 will be positioned
along the engine cowl 70, preferably proximate the rear bypass
outlet 81. Where the engine excludes an inner fixed structure 76,
the pressure relief mechanism 88 may be positioned along the engine
cowl 70.
[0114] Returning to the method of the present invention, there are
a number of inputs that may be evaluated to provide a signal for
the pressure relief door 90 to open at the appropriate time. To
assess the various inputs that are available on the aircraft 110,
the controller 202, which operatively connects to the pressure
relief door 90, also is understood to operatively connect to one or
more of the systems present on the aircraft.
[0115] In one instance, the controller 202 may be connected to the
bleed air systems in the engines 10 to receive a signal indicating
if one of the bleed air systems ceases to function. For example,
sensors may be positioned on or near the conduits 44, 46. If one of
the conduits 44, 46 should rupture, the sensor may be configured to
indicate such a failure, which would be interpreted as a failure of
the bleed air system. That signal could then be used, in
combination with other signals, to determine if the pressure relief
door 90 is to be opened.
[0116] In another instance, as indicated above, the controller 202
may be connected to the wing anti-icing system ("WAIS"). So
connected, the controller 202 is anticipated to receive signals
indicating if the WAIS is operating or not. In addition, the WAIS
may be configured to indicate the amount of hot bleed output air 52
demanded by the WAIS at any given moment in time. If so, the
controller 202 may be configured to trigger the opening of the
pressure relief door 90 only after receipt of specific input
signals associated with the amount of hot bleed output air 52
demanded by the WAIS.
[0117] In a further contemplated embodiment, the controller 202 may
be configured to receive signals separately from the hot bleed air
produced from the high and low pressure compressors 28, 30. As
such, if a signal is received suggesting that too much or too
little hot bleed air is being received from either of the
compressors 28, 30, this may result in a trigger event where the
pressure relief door 90 is opened or closed.
[0118] As noted above, bleed air may be taken from several
locations within the engine 10. Specifically, bleed air may be
taken from a location near to the low pressure compressor 28. Bleed
air taken from this location may be at a temperature of up to about
120.degree. C. Bleed air also may be taken from a location near to
the high pressure compressor 30. If so, the bleed air may be at a
temperature of up to about 500.degree. C.
[0119] It is contemplated that, if the aircraft is in idle or in a
descent mode, the bleed air may need to be taken from a position
near to the high pressure compressor 30 to ensure sufficient bleed
air pressure. A should be apparent, when the bleed air is taken
from the high pressure compressor 30, since it is at a higher
temperature, the need for cooling, ambient air is greater. When in
idle or in a descent mode, the precooler 50 may require an
increased flow of ambient air through the ambient air passage,
which demand may be satisfied by opening the pressure relief door
90.
[0120] It is also noted that bleed air from both compressors 28, 30
may be combined in a controlled manner by the controller 202 or
another system. For example, bleed air from the two compressors 28,
30 may be adjusted, in terms of their respective pressures, to
produce hot bleed input air 42 having predetermined pressure and
temperature characteristics.
[0121] Another signal that may be taken into account is a signal
from the wheels of the aircraft 110. Specifically, aircraft 110
typically are equipped with sensors that provide a signal if there
is a "weight on wheel off" event. This signal indicates if the
aircraft 110 is on the ground or in flight. Specifically, the
signal indicates if there is weight on the wheels or not. The
controller 202 may be configured to accept this signal and, if the
aircraft 110 is on the ground, the controller may be configured not
to open the pressure relief door 90, for example.
[0122] In a further contemplated embodiment, the controller 202 may
be configured to receive signals concerning the pressure and/or
temperature of the hot bleed input air 42. If the pressure and/or
temperature exceeds or falls below predetermined pressure and/or
temperature thresholds, the controller 202 may be configured to
open or close the pressure relief doors 90.
[0123] It is also contemplated that time may be a factor taken into
account by the controller 202. For example, with respect to the
pressure and/or temperature of the hot bleed input air 42, if the
pressure and/or temperature are above or below a predetermined
threshold for a predetermined period of time, the controller 202
may be configured to open or close the pressure relief door 90. By
taking into account the time or duration of a pressure/temperature
spike/depression, the controller 202 may avoid opening or closing
the pressure relief door 90 during periods where there is a
transient spike or depression that corrects itself within the
window of the time period or duration.
[0124] Still further, while the operation of the controller 202 has
been described in connection with the pressure and/or temperature
of the hot bleed input air 42, similar triggering events may be
defined for the hot bleed output air 52, the ambient input air 54,
and/or the ambient output air 64. Sensors may be provided at
various locations with respect to the precooler 50 to generate
these signals for use by the controller 202.
[0125] Another embodiment of the aircraft 110 of the present
invention contemplates that the controller 202 also will receive
signals indicative of the altitude of the aircraft 110. The
altitude signals also may be employed as a factor for the
controller to issue a trigger signal that opens or closes the
pressure relief door 90. For example, the controller 202 may be
configured to keep the pressure relief door 90 closed in one
instance of low altitude but opened in a separate instance at a
higher altitude.
[0126] As should be apparent from the foregoing, pressure sensors
are contemplated to be provided within the stream of the ambient
input air 54 and also within the stream of the ambient output air
64. The pressure sensors provide information so that the controller
202 may determine what the pressure drop is across the precooler
50. If the pressure differential falls below a predetermined
threshold, such as 0.5 psig or 1.0 psig, for example, the
controller 202 may be configured to open the pressure relief door
90 to increase the pressure differential.
[0127] In the foregoing discussion, the operation of the pressure
relief door 90 has been described as opened or closed. This is
intended to convey that the pressure relief door 90 is opened or
closed completely. In a separate embodiment, the controller 202 may
be configured to open the pressure relief door 90 by only a
fractional amount, depending upon the various input signals
received thereby. In this configuration, it is contemplated that
the pressure relief door 90 may be opened only partially to more
accurately control the pressure differential across the precooler
50.
[0128] In view of the foregoing, and with reference to FIG. 7, one
embodiment of the method of the present invention contemplates a
series of steps that include receiving by the controller 202, at
step 150, trigger variables, evaluating the trigger variables, by
the controller 202, at step 152, to determine if the air is passing
through the ambient air passage to sufficiently cool the bleed air
passing through the bleed air passage, and opening the pressure
relief door 90, at step 154, if the trigger variables satisfy
predetermined conditions.
[0129] The trigger variables include, among other variables, a
signal indicative of an engine failure, a bleed air system failure,
an ice condition, or activation of an anti-icing system. Other
trigger variables include, but are not limited to a signal
indicative of flow of air, flow of bleed air, pressure at the
ambient air inlet, pressure at the ambient air outlet, pressure at
the bleed air inlet, pressure at the bleed air outlet, outside air
temperature, temperature at the air inlet, temperature at the air
outlet, temperature at the bleed air inlet, temperature at the
bleed air outlet, differential pressure between the air inlet and
the air outlet, differential pressure between the bleed air inlet
and the bleed air outlet, weight on wheel off, altitude of the
aircraft, and time.
[0130] It will also be appreciated that the present invention could
be advantageously employed in aircraft with more than two
engines.
[0131] As being now better appreciated, the present invention is an
improvement and presents several advantages over other related
devices and/or methods known in the prior art. As should be
apparent to those skilled in the art, numerous modifications could
be made to the above-described embodiments without departing from
the scope of the invention, as apparent to a person skilled in the
art.
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