(12) INTERNATIONAL APPLICATION PUBLISHED UNDER THE PATENT COOPERATION TREATY (PCT)
`
`(19) World Intellectual Property Organization
`International Bureau
`
`I lllll llllllll II llllll lllll llll I II Ill lllll lllll 111111111111111111111111111111111
`
`(43) International Publication Date
`28 February 2002 (28.02.2002)
`
`PCT
`
`(10) International Publication Number
`WO 02/16743 Al
`
`(51) International Patent Classification7: F02C 7/224, 6/08
`
`(21) International Application Number: PCT/USOl/24756
`
`(22) International Filing Date: 8 August 2001 (08.08.2001)
`
`(25) Filing Language:
`
`(26) Publication Language:
`
`English
`
`English
`
`(30) Priority Data:
`09/643,996
`
`22 August 2000 (22.08.2000) us
`
`(74) Agent: STEPHENSON, Gregory, R.; Hamilton Sund(cid:173)
`strand Corporation, One Hamilton Road, MS 1-1-BC18,
`Windsor Locks, CT 06096-1010 (US).
`
`(81) Designated States (national): AE, AG, AL, AM, AT, AU,
`AZ, BA, BB, BG, BR, BY, BZ, CA, CH, CN, CO, CR, CU,
`CZ, DE, DK, DM, DZ, EC, EE, ES, FI, GB, GD, GE, GH,
`GM, HR, HU, ID, IL, IN, IS, JP, KE, KG, KP, KR, KZ, LC,
`LK, LR, LS, LT, LU, LV, MA, MD, MG, MK, MN, MW,
`MX, MZ, NO, NZ, PL, PT, RO, RU, SD, SE, SG, SI, SK,
`SL, TJ, TM, TR, TT, TZ, UA, UG, UZ, VN, YU, ZA, ZW.
`
`(84) Designated States (regional): European patent (AT, BE,
`CH, CY, DE, DK, ES, FI, FR, GB, GR, IE, IT, LU, MC,
`NL, PT, SE, TR).
`
`(71) Applicant: HAMILTON SUNDSTRAND CORPORA-
`TION [US/US]; One Hamilton Road, Windsor Locks, CT Published:
`06096-1010 (US).
`with international search report
`
`(72) Inventors: WILMOT, George, E., Jr.; 4 Surrey Drive,
`East Granby, CT 06026 (US). OTT, Gregory, M.; 24 Jamie
`Lane, Feeding Hills, MA 01030 (US).
`
`For two-letter codes and other abbreviations, refer to the "Guid(cid:173)
`ance Notes on Codes and Abbreviations" appearing at the begin(cid:173)
`ning of each regular issue of the PCT Gazette.
`
`iiiiiiii
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`----iiiiiiii
`== --
`!!!!!!!! -
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`iiiiiiii
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`iiiiiiii ~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~
`!!!!!!!!
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`(54) Title: INTEGRATED THERMAL MANAGEMENT AND COOLANT SYSTEM FOR AN AIRCRAFT
`
`----10
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`!!!!!!!! -iiiiiiii
`iiiiiiii ----
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`~
`~ (57) Abstract: A thermal management system avoids problems associated with the recirculation of fuel through a fuel tank on an
`~ aircraft in a system that includes a fuel reservoir (70), a pump (72) for pumping fuel from the reservoir (70) to a first heat load (74)
`,....i and then to a check valve (76) to a junction (78). At least one second heat load (80), (82), (84) is connected to the junction (78) and
`-..... a fuel/fan air heat exchanger (104) having a fan air flow path in heat exchange relation with a fuel flow path having a fuel discharge
`~ end (106) connected to the junction (78). The fan air flow path is adapted to be connected to receive compressed air from the fan (30)
`of an engine (10). A bleed air/fuel heat exchanger (92) receives fuel from the second heat load and bleed air from the engine (10) and
`0 is connected to deliver fuel to the engine (10). A control valve (86) is interposed between the second heat load and the engine (10)
`> for proportioning the flow of fuel to the engine (10) and the fuel fan heat exchanger (104) so that excess fuel is recirculated without
`~ reintroduction into the fuel reservoir (70).
`
`GE-1006.001
`
`

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`WO 02/16743
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`PCT/USOl/24756
`
`INTEGRATED THERMAL MANAGEMENT AND
`COOLANT SYSTEM FOR AN AIRCRAFT
`
`Technical Field
`
`s
`
`This invention relates to an integrated thermal management and
`
`coolant system for an aircraft, particularly an aircraft having a turbo fan
`
`propulsion engine.
`
`Background Art
`
`Modern sophisticated aircraft require equally sophisticated systems
`
`10
`
`for thermal management and cooling. In typical modes of operation of the
`
`aircraft, lubricating oil for the engine and hydraulic fluid used in the various
`
`hydraulic systems as well as the aircraft mounted accessory drive (AMAD)
`
`require cooling. At the same time, the avionic systems of the aircraft will
`
`require cooling during operation, some by a liquid coolant and others by cool
`
`15
`
`air. Concurrently, at low altitude or on the ground or at other relatively high
`
`temperature operating environments, the aircraft cabin requires cooling
`
`while at relatively low temperature altitudes such as at cruise altitude for a
`
`jet aircraft, the cabin will require warming. Through all of this it is generally
`
`desirable to heat the fuel delivered to the main propulsion engines to
`
`26 maximize the efficiency of the engines.
`
`A common thread in many prior art systems is the use of the air in
`
`which the aircraft is traveling as a sink into which heat is rejected. Most
`
`typically, this air is both so-called "ram air" and "bleed air". Ram air is, of
`
`course, air that is literally rammed into an inlet on the aircraft as a result of
`
`25
`
`the aircraft's forward velocity through a body of air. A penalty paid for the
`
`1
`
`GE-1006.002
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`

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`WO 02/16743
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`
`use of ram air is the aerodynamic drag imposed on the aircraft as a result of
`
`the presence of one or more ram inlets.
`
`In addition, the installation of ram air circuits in an aircraft so as to
`
`convey the ram air to a point of use is difficult. Moreover, in the case of
`
`5
`
`aircraft intended for military use, ram air inlets all too often may provide an
`
`undesirable aircraft position indicating radar return because of their
`
`configuration.
`
`Bleed air is air taken from the compressor section of a gas turbine
`
`engine, whether a main propulsion engine of the aircraft or a so-called APU
`
`10
`
`or auxiliary power unit. A penalty paid for the use of bleed air is a reduction
`
`in operating efficiency of the engine from which the air is bled.
`
`Many of these systems utilize aircraft fuel as a coolant prior to its
`
`combustion in an engine. The aircraft fuel cannot be heated to such a
`
`degree that it begins to "coke" and consequently, excess fuel is circulated to
`
`15
`
`the components that it is to cool and that fuel not required by the engine is
`
`recirculated to the fuel reservoir. Not infrequently, this fuel, particularly in
`
`military usages, may be returned to another, larger storage reservoir, from
`
`which fuel is withdrawn to be put to uses other than that of driving the main
`
`propulsion engines of the aircraft. Because of that possibility, desirable
`
`20
`
`additives for aircraft operation cannot be utilized in the fuel. For example, it
`
`is desirable to use additives that raise the temperature of the fuel at which
`
`coking begins to occur and the presence of such additives may not be
`
`desirable for all uses to which the fuel is put and can be expensive.
`
`The present invention is directed to overcoming one or more of the
`
`25
`
`above problems.
`
`2
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`GE-1006.003
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`

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`WO 02/16743
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`PCT/USOl/24756
`
`Disclosure of Invention
`
`It is the principal object of the invention to provide a new and
`
`improved, integrated, thermal management and environmental cooling
`
`system. An exemplary embodiment of the invention contains a number of
`
`5
`
`facets which, in a highly preferred embodiment, are all used together.
`
`However, in some instances, components giving but a single advantage may
`
`be employed exclusively without resort to the others or in such combinations
`
`as to achieve those of the advantages specifically desired.
`
`According to one aspect of the invention, the system includes a fuel
`
`10
`
`reservoir with a pump pumping fuel from the reservoir to a first heat load. A
`
`check valve is located downstream of the first heat load and upstream of a
`
`second heat load. The fuel, after being passed through the heat loads, is
`
`further heated by engine bleed air with part being diverted to the main
`
`propulsion engine for combustion therein and the remainder being
`
`15
`
`recirculat.ed through a heat exchanger cooled by fan air from an early stage
`
`of the engine and returned to the fuel line downstream of the check valve.
`
`Consequently, recirculation is not through the fuel tank and the problems
`
`associated with the return of recirculated fuel to the fuel reservoir or tank are
`
`avoided.
`
`20
`
`According to another aspect of the invention, the thermal
`
`management and coolant system avoids the use of ram air altogether by
`
`employing, when required, engine fan air from the fan duct of a turbo fan
`
`engine and engine bleed air from the compressor of the main engine core as
`
`the air utilized throughout the system in exclusion to ram air entirely.-
`
`25
`
`According to the invention and another facet thereof, heat
`
`exchangers for cooling t~e air used as a working fluid in the system are
`
`located in the bypass or fan air duct of the engine to reduce the fuselage
`
`3
`
`GE-1006.004
`
`

`
`WO 02/16743
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`
`volume and to increase efficiency by rejecting heat to the fan or bypass air
`
`passing through the engine.
`
`In another facet of the invention, fan air utilized in a heat exchanger
`
`to cool fuel that is being recirculated to the heat loads is discharged to the
`
`5
`
`engine bay to provide positive bay ventilation and eliminate the need for
`
`separate engine bay ventilation circuits.
`
`In still another aspect of the invention, the system employs a turbo
`
`machine having at least one turbine stage. Engine fan air is expanded in
`
`the one turbine stage, and thereby cooled while undergoing expansion,
`
`10
`
`before being discharged to an early part of a thermal management network.
`
`The one turbine stage thereby lowers the temperature of the air being
`
`utilized as the sink for the system.
`
`Additionally, the system of the invention, because it does not require
`
`ram air, may be located extremely close to the main propulsion engine to
`
`15 minimize ducting.
`
`Other objects and advantages will become apparent from the
`
`following specification taken in connection with the accompanying drawings.
`
`Brief Description of Drawing
`
`The Fig. is a schematic of an integrated thermal management coolant
`
`20
`
`system made according to the invention.
`
`4
`
`GE-1006.005
`
`

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`WO 02/16743
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`PCT/USOl/24756
`
`Best Mode for Carrying Out the Invention
`
`An exemplary embodiment of an integrated thermal management
`
`system made according to the invention is illustrated in the Fig. and will be
`
`described in the context of an aircraft having one or more jet main
`
`5
`
`propulsion engines, specifically, turbo fan engines. However, in some
`
`instances, where engine bleed air may be substituted for engine fan air, or
`
`where only part of the system is employed and that does not require the use
`
`of fan air, any type of turbine based main propulsion engine could be
`
`employed.
`
`10
`
`As seen in the Fig., the system, including the main propulsion engine,
`
`includes four main components. Specifically, a turbo fan propulsion engine,
`
`generally designated 10, is illustrated. Also illustrated is what might be
`
`termed a fuel based system, generally designated 12. The fuel based
`
`thermal management system 12 utilizes fuel for the aircraft to cool various
`
`15
`
`heat loads forming part of the aircraft.
`
`A third component is a turbo machine, generally designated 14. The
`
`turbo machine expands the working fluid through one or more turbine stages
`
`to cool the same to a lower temperature so that it may more efficiently cool
`
`working fluid used elsewhere in the system. The turbo machine 14 also
`
`20
`
`compresses the working fluid which, as noted earlier, is air, for use by the
`
`aircraft environmental control system (ECS).
`
`A final major component is a thermal management network, generally
`
`designated 16. Within the network 16, the working fluid at various stages of
`
`expansion and compression are moved about in paths to be described to
`
`25
`
`achieve the desired control of cabin temperature, provide air cooling for
`
`avionics, provide cooling for a liquid coolant loop which may cool other
`
`avionic systems, etc., etc.
`
`5
`
`GE-1006.006
`
`

`
`WO 02/16743
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`
`Turning now the main propulsion engine 10, the same is seen to
`
`include a rotary shaft, schematically illustrated at 20 carrying several stages
`
`of compressor blades 22 and several stages of turbine blades 24. One or
`
`more combustors 26 are located in this engine 10 for com busting fuel
`
`5
`
`received on a line 28 to provide gases of combustion to drive the turbine 24.
`
`The shaft 20, forwardly of the compressor stages 22, also mounts
`
`several stages of fan blades 30. The arrangement is one of a conventional
`
`turbo fan jet propulsion engine. The compressor 22, the turbine 24 and the
`
`combustor 26 make up the main core of the engine 10 and gases of
`
`10
`
`combustion expanded through the turbine 24 are discharged through a
`
`nozzle 32. Surrounding the engine core is a bypass or fan duct 34 which
`
`also surrounds the fan blades 30. As is well known, air under pressure
`
`admitted at the engine inlet 36 upstream of the fan blades 30 is not only
`
`compressed by the compressor 22, but by the fan blades 30 as well. The air
`
`15
`
`that passes about the engine core is known as fan air or bypass air and is
`
`confined about the engine core by the duct 34. The same is discharged
`
`about the nozzle 32 at a nozzle designated 38.
`
`The engine 10 includes an outlet 40 for fan air. The outlet 40 is
`
`connected to the interior of the fan duct 34 at an early stage of the fan
`
`20
`
`section 30 of the engine 10 and to a modulating valve 42. It is desirable to
`
`locate the outlet 40 at an early stage since at that point, compression will be
`
`relatively minimal with a consequence that the temperature rise in the air
`
`being compressed by the fan 30 will also be relatively minimal. That is to
`
`say, the air will be cooler at the outlet 40 at the location shown than if it were
`
`25
`
`located more downstream on the fan duct 34.
`
`A further fan air outlet 44 is located downstream of the outlet 40 and
`
`likewise connects to a modulating valve 46 for purpose to be seen.
`
`6
`
`GE-1006.007
`
`

`
`WO 02/16743
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`
`The engine 10 includes a bleed air outlet 48 connected to a late or
`
`final stage of the compressor 22. The outlet 48 provides engine bleed air
`
`from the compressor 22 to a modulating valve 50 which in turn is connected
`
`to a check valve 52 and a further modulating valve 54. A line 56 extends
`
`5
`
`from the modulating valve 54 to a first fan duct heat exchanger 58 and
`
`specifically, to the downstream end of the heat exchanger 58 in relation to
`
`the fan duct 34. Thus, bleed air may be passed to the heat exchanger 58
`
`and flow therein countercurrent to fan air flowing within the duct 34 to
`
`emerge from the first fan duct heat exchanger 58 on a line 60 for purposes
`
`10
`
`to be seen.
`
`A second fan duct heat exchanger 62 is likewise located within the
`
`fan duct 34 about the nozzle 32 and includes an inlet 64 as well as an outlet
`
`66. Both of the heat exchangers 58 and 62 have internal flow paths that are
`
`in heat exchange relation with the air flowing in the fan duct 34 as will be
`
`15
`
`seen. While both of the heat exchangers 58 and 62 have been shown and
`
`described as providing for countercurrent flow of a working fluid within each
`
`heat exchanger in relation to the flow of bypass air within the fan duct 34,
`
`concurrent flow, or cross flow, or combinations of the same could be utilized
`
`as desired, depending upon the heat exchange efficiency required of a given
`
`20
`
`system. In any event, it can be seen that bypass air can be utilized as the
`
`sink for heat within the syst~m being rejected through the heat exchangers
`
`58 and 62. Additionally, the heat exchanger 62 need not be located
`
`downstream of the heat exchanger 58 as shown, if desired.
`
`Returning now to the fuel based side 12 of the system, the same is
`
`25
`
`seen to include a fuel reservoir or tank 70 for containing fuel to drive the
`
`engine 10. The fuel reservoir 70 is connected to a pump 72 which pumps
`
`the fuel to a first heat load 7 4 which is indicated as an electrical load but
`
`7
`
`GE-1006.008
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`

`
`WO 02/16743
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`PCT/USOl/24756
`
`could be any sort of heat load requiring cooling. After cooling the first heat
`
`load 7 4, the fuel passes through a check valve 76 which allows flow in the
`
`direction of the arrow illustrated but not the reverse. The check valve 76
`
`therefore discharges fuel to a junction 78.
`
`5
`
`From the junction, fuel passes through a series of heat exchangers,
`
`in the embodiment illustrated, a hydraulic fluid/fuel heat exchanger 80, an
`
`AMAD/fuel heat exchanger 82 and an engine oil/fuel heat exchanger 84. Of
`
`course, there may be greater or fewer of the heat exchangers 80, 82 and 84
`
`dependent upon the aircraft systems requiring cooling. In the usual case,
`
`10
`
`however, the hydraulic/fuel heat exchanger 80 will receive hydraulic fluid
`
`from the hydraulic systems of the aircraft to cool the same and then return
`
`the now cooled hydraulic fluid to those systems. Similarly, the AMAD/fuel
`
`heat exchanger 82 receives hydraulic fluid and/or lubricating oil from the
`
`AMAD, cools the same and returns it to the AMAD. In a like fashion, the
`
`15
`
`engine oil/fuel heat exchanger 84 receives lubricating oil from the engine 10,
`
`cools the same and then returns the oil to the engine.
`
`After passing through the heat exchangers 80, 82, 84, the fuel is
`
`discharged to a modulating three way valve 86 which divides the fuel into
`
`two streams, one placed on a line 88 and the other placed on a line 90. Of
`
`20
`
`particular note is that the valve 86, in response to control signals provided
`
`elsewhere, proportions the fuel flow to the lines 88 and 90 for purposes to
`
`be seen.
`
`The fuel side 12 also includes a bleed air/fuel heat exchanger,
`
`generally designated 92, made up of first and second bleed air/fuel heat
`
`25
`
`exchanger cores 94 and 96 respectively. Each of the cores 94 and 96
`
`include a fuel flow path in heat exchange relation with a bleed air flow path.
`
`The bleed air flow paths of the heat exchanger cores 94 and 96 are
`
`8
`
`GE-1006.009
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`

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`WO 02/16743
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`connected in series to discharge through a line 98. On their inlet side, the
`
`same are connected via a three way modulating valve 100 to the outlet 60 of
`
`the first fan duct heat exchanger 58.
`
`The first bleed air fuel heat exchanger 94 has its fuel flow path
`
`5
`
`connected to the line 90 and discharges to the line 28 to provide fuel to the
`
`combustors 26 for combustion within the engine 10. The second bleed air
`
`fuel heat exchanger 96 is connected to the line 88 and discharges to a fuel
`
`inlet 102 of a fuel/air heat exchanger 104. The fuel/air heat exchanger has
`
`an outlet 106 which is connected to the junction 78 as illustrated. In addition
`
`10
`
`to the fuel flow path extending between the inlet 102 and outlet 106, the
`
`fuel/air heat exchanger 104 includes a flow path connected to the
`
`modulating valve 42 on its inlet side and on its outlet side to an engine bay
`
`108 in which the engine 10 is located. As a consequence of this
`
`construction, cooling air is provided from an early fan stage of the engine 10
`
`15
`
`in quantities proportioned through operation of the valve 42 to the fuel/air
`
`heat exchanger 104 to cool fuel being recirculated through the second bleed
`
`air fuel heat exchanger 96 to the junction 78 as well as to provide air to the
`
`engine bay 108 at a pressure that is positive with respect to that area and
`
`provide for ventilation of the same. Thus, the need for a separate engine
`
`20
`
`bay ventilation system is avoided while the fuel, after being utilized to cool
`
`aircraft fluids in the heat exchangers 80, 82 and 84, is recirculated after
`
`being cooled without being recirculated through the fuel reservoir 70. Thus,
`
`additives may be provided in the fuel to raise its coking temperature and
`
`allow the fuel to be provided to the engine 1 O at a higher temperature to
`
`25 maximize engine efficiency. Further, there is no need to carry excess fuel
`
`as in some prior art systems. It will also be appreciated that the recirculation
`
`system described with the advantages just listed further provides the
`
`9
`
`GE-1006.010
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`

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`WO 02/16743
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`
`advantage of operation with no reliance whatsoever on ram air, thus
`
`reducing aerodynamic drag on the aircraft in which it is employed and
`
`allowing the use of lesser volume ram air inlets if ram air is required for other
`
`purposes with such lesser volume ram air inlets minimizing the effect on the
`
`5
`
`radar cross section of the aircraft and reducing complexity and removing the
`
`doors used in typical ram air systems.
`
`It will be observed that the ultimate temperature of the fuel applied to
`
`the engine 10 is controlled by the percentage of the total fuel flow directed
`
`through the first bleed air fuel heat exchanger core 94 as controlled by the
`
`10
`
`valve 86, the flow of bleed air to the heat exchanger 58 as modulated by
`
`operation of the valve 126, and the flow of bleed air from the heat exchanger
`
`58 to the bleed air fuel heat exchanger 92 as controlled by the valve 100. In
`
`this respect, it will be observed that the latter is a three way valve and can
`
`direct bleed air flow from the heat exchanger 58 to a bypass line 112 which
`
`15
`
`connects to the line 98 at a point 114.
`
`It should also be observed that the bleed air fuel heat exchanger 92
`
`adds heat to the fuel flowing therethrough on the lines 88 and 90 so as to
`
`assure that fuel flowing to the engine 10 on the line 28 is at the most
`
`efficient, non-coking, elevated temperature allowed for system operation. It
`
`20
`
`should be noted that with the addition of fuel additives for the purpose of
`
`raising the coking temperatures of fuels to higher temperatures than
`
`currently achievable with conventional technology, the heat exchanger 58
`
`may be eliminated, with all bleed air cooling being achieved in the heat
`
`exchanger 92.
`
`25
`
`Further, an alternate configuration may be employed, depending on
`
`the physical design and particular materials employed in fabricating the
`
`valve 86. For example, the valve 86 may be located on the fuel outlet side
`
`10
`
`GE-1006.011
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`

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`of the heat exchanger 92. This results in a less complex construction of the
`
`heat exchanger. In this alternative construction, the fuel outlet side of the
`
`heat exchanger 84 would be connected directly to the fuel inlet side of the
`
`heat exchanger 92 while the fuel outlet side of the heat exchanger 92 would
`
`5
`
`be connected to the valve 86. The valve 86 would be modulated to supply
`
`fuel to the combustor 26 and directly to the fuel/fan air heat exchanger 104.
`
`Turning now to the turbo machine 14, the same, in the exemplary
`
`embodiment illustrated, includes a single shaft 116 on which a single rotary
`
`compressor stage 118 is mounted. First and second turbine stages 120 and
`
`10
`
`122 respectively, are also mounted on the shaft 116. However, it is
`
`specifically to be noted that coaxial shafts or other non-single shaft
`
`arrangements known in the art could be employed if desired.
`
`The compressor 118 includes an inlet 124 connected via a
`
`modulating valve 126 to the junction 114. Thus, engine bleed air, after
`
`15
`
`being cooled in the first fan duct heat exchanger 58 and directed to the
`
`junction 114 directly via the bypass 112 or via the bleed air fuel heat
`
`exchanger 92 is directed to the compressor stage 118. The compressor 118
`
`includes an outlet line 128 which is connected to the inlet 64 of the second
`
`fan duct heat exchanger 62 to provide engine bleed air, after a further stage
`
`20
`
`of compression, thereto. After being cooled in the second fan duct heat
`
`exchanger 62, this bleed air serves as the working fluid for the thermal
`
`management network 16 which is connected to the outlet 66 for the heat
`
`exchanger 62 as illustrated.
`
`Turning to the thermal management network 16, the same includes
`
`25
`
`an intercooler 130 or gas to gas heat exchanger.
`
`Internally, the intercooler 130 includes two gas flow paths in heat
`
`exchange relation to each other. One flow path is connected to the outlet 66
`
`11
`
`GE-1006.012
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`
`of the second fan duct heat exchanger 62 and to a so-called ground heat
`
`exchanger 134. The other gas flow path within the intercooler 130 is
`
`connected via a line 136 to the outlet 138 of the first turbine stage 120 of the
`
`turbo machine 14. This same gas flow path discharges overboard as
`
`5
`
`illustrated at 140. The first turbine stage 120 includes an inlet 142
`
`connected via check valve 144 to the valve 46 and the bleed air outlet 44.
`
`Thus, bleed air from the main propulsion engine 10 is directed, at a flow rate
`
`modulated by the valve 46, to the first turbine stage 120. The turbine thus
`
`drives the compressor 118 as well as expands the bleed air. The expansion
`
`10
`
`of the bleed air results in much cooler air exiting the first turbine stage 120
`
`and being directed to the intercooler 130. Thus, working fluid entering the
`
`thermal management network as sink air is cooled by the intercooler 130
`
`through the use of the turbo machine 14 which expands the sink air through
`
`the first turbine stage 120 to lower the temperature of the sink air.
`
`15
`
`If desired, provision may be made for the introduction of a ground
`
`heat exchanger 134 into the flow stream within the thermal management
`
`network 16. Typically, such a heat exchanger would be mounted on a
`
`wheeled ground cart or the like and have ambient air directed through it via
`
`a fan 146 to be discharged again to the ambient. It would be connected into
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`20
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`the system only while the aircraft is on the ground by suitable flexible
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`conduits.
`
`From the ground heat exchanger 134, if used, working fluid is passed
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`to a regenerator 148 and exits the same on a line 150. Further cooling of
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`the working fluid occurs within the regenerator by reason of substantially
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`25
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`spent working fluid being passed through the regenerator on a line 152 to a
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`modulating valve 154 to be dumped overboard as illustrated at 156.
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`GE-1006.013
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`WO 02/16743
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`In any event, working fluid in the line 150 is directed to a junction 158.
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`A bypass valve 160 connected to the junction 158 is operable, when open,
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`to direct some part of the fluid in the line 150 directly to the inlet 162 of the
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`second turbine stage 122. When the valve 160 is closed, all of the working
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`5
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`fluid on the line 150 is directed to a heat exchanger in the form of a reheater
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`164. The reheater 164 is desirably included to increase system efficiency
`
`but is not absolutely necessary and may be omitted, if desired, for cost
`
`reduction.
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`In the reheater 164, the working fluid from the junction 158 passes
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`10
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`through a flow path that is in heat exchange with a flow path extending from
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`an inlet 166 to the reheater 164 to an outlet 168 thereof. Within the reheater
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`164, working fluid actually warms the fluid passing from the inlet 166 to the
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`outlet 168.
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`The working fluid exits the reheater 164 on a line 170 and passes to a
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`15
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`condenser 172. Within the condenser 172, the working fluid is cooled and
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`to the extent it is cooled below its dew point, condensation of moisture will
`
`occur. That condensate will -be collected by a water collector or condensate
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`collector 17 4 on the outlet side of the condenser 172. To the extent that
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`cold water might be usable on the aircraft as, for example, cooling
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`20
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`components requiring water cooling or the like, the water separated in the
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`water collector 17 4 may be utilized for the purpose. In some
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`applications/installations the reheater 164 may be made as part of the
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`condenser 172, thus eliminating duct 170.
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`The working fluid exiting the condenser 172 is returned on a line 176
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`25
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`to the inlet 166 of the reheater 164. Because the fluid was cooled
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`substantially within the condenser 172, it is now reheated somewhat within
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`the reheater 164 by the incoming working fluid and then applied to the inlet
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`GE-1006.014
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`WO 02/16743
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`162 of the second turbine stage 122. This working fluid will be relatively free
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`of condensate and above the dew point of the working fluid. In this regard,
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`the bypass valve 160 is utilized to bypass the condenser 172 only in low
`
`humidity conditions as, for example, when the aircraft is flying at high
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`5
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`altitude (e.g., 30,000 ft.).
`
`In all events, the working fluid is expanded in the second stage
`
`turbine 122 and directed therefrom via an outlet 180 at a very, very low
`
`temperature.
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`Within the thermal management network 16 is a liquid coolant loop,
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`10
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`generally designated 182. A liquid coolant is flowed through the loop 182
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`and typically, but not always, will be a PAO fluid, that is, a polyalphaolefin
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`heat transport fluid. Thus, a first coolant heat exchanger 184 has a liquid
`
`flow path designated by an inlet 186 and an outlet 188 and a liquid coolant
`
`flow path in heat exchange therewith. The liquid coolant flow path is defined
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`15
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`by an inlet 190 and an outlet 192. The outlet 192 is connected to a heat
`
`load that requires liquid cooling as, for example, liquid cooled avionic
`
`systems 194. From there, the loop extends to a pump 196 which directs
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`liquid coolant through a three way valve 198 whereby it may be returned
`
`directly to the outlet 192 of the heat exchanger 184 or provided to a coolant
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`20
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`inlet 200 for a coolant flow path through a second coolant heat exchanger
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`202 to an outlet 204. The outlet in turn is connected to the inlet 190 of the
`
`first liquid coolant heat exchanger 184 to complete the loop 182.
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`The second liquid coolant heat exchanger 202 has a second fluid flow
`
`path in heat exchange relation with that defined by the inlet 200 and the
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`25
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`outlet 204 which acts as part of the line 152 and connects to the regenerator
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`148 to provide the aforementioned cooling action of the working fluid therein.
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`GE-1006.015
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`Returning to the first liquid coolant heat exchanger 184, the outlet 188
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`thereof is connected to a flow path through the condenser defined by an
`
`inlet 206 and an outlet 208. Thus, cold expanded air from the turbine stage
`
`122, after being directed through the first liquid coolant heat exchanger 184,
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`5
`
`is passed through the condenser 172 to cool the incoming working fluid and
`
`cause condensation of moisture therein. After emerging from the condenser
`
`from the outlet 208, the relatively cooled air is directed to a line 21 O which is
`
`connected to the aircraft cabin 212 via a modulating valve 214 and a check
`
`valve 216. Thus, cold air is provided from the condenser 208 to the cabin
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`10
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`212 for cooling purposes.
`
`The cool air from the condenser 208 may also be directed to a
`
`modulating valve 218 which directs the flow of such air to air cooled avionics
`
`220 for cooling purposes. Finally, the air from the condenser 208 may be
`
`passed to a line 222 connected to the regenerator 148 via the second liquid
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`15
`
`coolant heat exchanger 202 to provide for further cooling of the liquid
`
`coolant used to cool the liquid cooled avionics 194.
`
`The system may also include, if desired, a line with a modulating
`
`valve 230 interconnecting the inlet 124 of the compressor 118 and the inlet
`
`186 of the first liquid coolant heat exchanger 184. The valve 230 may be
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`20
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`operated to cause the bypassing of the turbo machine 14 in selective
`
`amounts.
`
`Also included is a line 232 extending from the valve 126. The line
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`232 also connects to the modulating valve 214 and includes its own
`
`modulating valve 234.
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`25
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`Those skilled in the art will readily appreciate that the temperature of
`
`bleed air at the valve 126 will be considerably higher than that at the outlet
`
`208 of the condenser 172. Those skilled in the art will also recognize that
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`GE-1006.016
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`WO 02/16743
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`with environmental control systems operating at altitude, it is almost always
`
`necessary to warm the air flowing to the cabin or other components rather
`
`than cool it. Thus, by selective operation of the valve 234, relatively warm
`
`air may be fed to the cabin 212 to be mixed with cool air from the condenser
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`5
`
`172 to achieve a desired temperature balance.
`
`To make full effective use of the system, in some instances, it may be
`
`desirable to construct the same so as to be usable employing bleed air from
`
`an auxiliary power unit or APU. An APU is schematically illustrated at 240
`
`and includes a bleed air output 242. The bleed air at the outlet 242 may be
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`10
`
`directed through a line 244 through an off/on valve 246 and a check valve
`
`248 to the inlet 142 of the first turbine stage 120. It thus may be used to