US3528250A - Bypass engine with afterburning and compressor bleed air heat exchanger in bypass duct - Google Patents

Bypass engine with afterburning and compressor bleed air heat exchanger in bypass duct Download PDF

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US3528250A
US3528250A US3528250DA US3528250A US 3528250 A US3528250 A US 3528250A US 3528250D A US3528250D A US 3528250DA US 3528250 A US3528250 A US 3528250A
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duct
air
bypass
engine
burner
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Douglas Johnson
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Motors Liquidation Co
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Motors Liquidation Co
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F28HEAT EXCHANGE IN GENERAL
    • F28DHEAT-EXCHANGE APPARATUS, NOT PROVIDED FOR IN ANOTHER SUBCLASS, IN WHICH THE HEAT-EXCHANGE MEDIA DO NOT COME INTO DIRECT CONTACT
    • F28D7/00Heat-exchange apparatus having stationary tubular conduit assemblies for both heat-exchange media, the media being in contact with different sides of a conduit wall
    • F28D7/16Heat-exchange apparatus having stationary tubular conduit assemblies for both heat-exchange media, the media being in contact with different sides of a conduit wall the conduits being arranged in parallel spaced relation
    • F28D7/1615Heat-exchange apparatus having stationary tubular conduit assemblies for both heat-exchange media, the media being in contact with different sides of a conduit wall the conduits being arranged in parallel spaced relation the conduits being inside a casing and extending at an angle to the longitudinal axis of the casing; the conduits crossing the conduit for the other heat exchange medium
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12Cooling of plants
    • F02C7/16Cooling of plants characterised by cooling medium
    • F02C7/18Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • F02C7/185Cooling means for reducing the temperature of the cooling air or gas
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K3/00Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/08Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluid; Control thereof
    • F02K3/105Heating the by-pass flow
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/60Efficient propulsion technologies
    • Y02T50/67Relevant aircraft propulsion technologies
    • Y02T50/675Enabling an increased combustion temperature by cooling

Description

3,528,250 AFTERBURNING AND COMPRESSOR EXGHANGER IN BYPASS DUCT BLEED Sept. 7 15, 1970 D. JOHNSON BYPASS ENGINE WITH AIR HEAT 2 Sheets-Sheet 1 Filed April 16, 1969 IAVENTUR. @415 Mm TTORNEY D. JOHNSON 3,528,250 BYPASS ENGINE WITH AFTERBURNING AND COMPRESSOR BLEED Sept. 15, 1970 AIR HEAT EXCHANGER IN BYPASS DUCT 2 Sheets-Sheet 2 Filed April 16, 1969 ATTORNEY United States Patent US. Cl. 60-261 8 Claims ABSTRACT OF THE DISCLOSURE A bypass turbojet engine includes an afterburner for reheating the turbine exhaust and also a burner in the bypass duct. Air flowing to the bypass burner is warmed by heat exchangers mounted in the line of air flow to the burner through which air is circulated from a high pressure stage of the compressor to the turbine for cooling the turbine.

My invention is directed to an improvement of bypass turbojet engines having a duct burner or burners to increase the temperature of the air discharged from the bypass duct. Essentially, my invention provides for improving the combustion in such duct burners, and at the Same time improving the efiiciency of turbine cooling, by circulating cooling air from the compressor of the engine to the turbine through a heat exchanger or heat exchangers mounted in the path of air flow through the bypass duct to the duct burner or burners.

A typical bypass type jet engine of the sort in which my invention preferably is incorporated comprises a gas turbine engine which drives a fan or low pressure compressor which supplies part of its discharge to the gas turbine engine and the remainder to a bypass duct, which ordinarily is an annular duct surrounding the gas turbine engine. The flow through the bypass duct and the exhaust from the engine are combined and discharged through a jet propulsion nozzle. The thrust in such engines may be increased for takeoff or during high speed flight by afterburning equipment located downstream of the turbine and duct burners serving to heat the air flowing through the bypass duct. The burners operating in the turbine exhaust burn the fuel in quite hot gas which still contains a considerable part, ordinarily the major part, of its oxygen uncombined. The bypass burner, on the other hand, burns fuel in air which, while warmed by the work exerted by the fan, is far cooler than the engine exhaust. There are greater problems of stability and smoothness of combustion in this relatively cool bypass air. This means that a lower gas velocity, larger combustor volume and more turbulent mixing are required in the vicinity of the flameholding elements which results in a larger, heavier burner and increased pressure drop. The increase in pressure drop reduces engine output.

I have concluded that it is feasible to heat a core of air going to the duct burner by a heat exchanger so located as to concentrate the heat output in the line of flow to the burner. If such heat exchangers are employed to cool air which is bled from the engine compressor to cool the turbine, such cooling increases the effectiveness of the cooling air and thereby decreases the amount of air which must be so used, thereby increasing the efliciency of the engine.

In a typical bypass engine, I calculate that it is feasible to increase the temperature of the core of air reaching the duct burner by approximately 100 F. by transferring heat from the turbine cooling air. This increase in temperature can the expected to give approximately fifteen percent improvement in stability of combustion in the m we duct burners. In this case, the combustion stability index is related to ease of starting. In an engine of the reheat type, the duct burner is not used at all times. It is used at takeoff and at high speed or supersonic flight. During subsonic flight, such as holding during a landing, the reheat system is not used. Thus, it is extremely important that the system ignite reliably every time it is called upon.

The principal objects of my invention are to improve the efficiency and reliability of bypass turbojet engines incorporating thrust augmenting burners, to improve the efiiciency of cooling of gas turbine engines, and to improve the operation of bypass duct air heating burners of jet engines. A further object is to provide heat exchanger and burner structure particularly suitable to accomplish the objects above stated.

The nature of my invention and the advantages thereof will be apparent to those skilled in the art from the succeeding detailed description of the preferred embodiment of the invention. This description, it is to be understood, is for the purpose of explaining the principles of the invention and is not presented in a limiting sense, since many modifications of structure may be made within the scope of the invention.

The preferred embodiment of the invention is illustrated in the drawings, in which FIG. 1 is a sectional view of a bypass turbojet engine with aft rburner, illustrating particularly the air flow paths through the engine.

FIG. 2 is a fragmentary sectional view to a greater scale, taken on a plane transverse to the axis of the engine as indicated by the line 22 in FIG. 1.

FIG. 3 is an enlarged view of a portion of FIG. 1, illustrating the heat exchanger and duct burner more clearly.

Referring first to FIG. 1, the ducted fan turbojet engine includes an outer casing 5 extending from an annular air inlet 6 to a variable jet propulsion nozzle 7. A threestage low pressure compressor or fan 9 is mounted at the inlet end of the outer casing, part of the discharge of this compressor being supplied to a bypass duct 10 defined between the casing 5 and an inner wall 11. The remainder of the low pressure compressor output is fed into an annular engine duct 13 defined by the engine outer case 14 and the usual interior walls (not identified). The engine case 14- encloses a high pressure compressor 15, combustion apparatus 17, a high pressure turbine 18, and a low pressure turbine 19. The two turbines might be combined into a single turbine, but this is not preferred. The bypass duct terminates in an annular outlet 21 and the engine duct terminates in an annular outlet 22 Within the bypass outlet. The flows through the bypass duct and through the turbine are combined in an afterburner 23 in which combustion takes place and in which the gases are mixed before being exhausted through the nozzle 7.

The int rnal structure of the rotating machinery such as the rotors of the compressors and turbines and the bearings and shafting are not illustrated, since the structure of these is entirely immaterial to my invention. The engine as so far described may be considered to be conventional.

Referring now to FIGS. 2 and 3, the combustion apparatus 17 includes an annular combustion liner which includes a number of fuel nozzles supplied through lines partially shown as 26. Air for the combustion apparatus is supplied from the compressor 15 and the discharge from the combustion apparatus flows through the blading of turbines 18 and 19. Fuel is sprayed into the turbine exhaust from annular manifolds 27 and the flame is retained in the afterburner by vaporizing flameholders 29.

Heating of the bypass air is accomplished by an annular bypass burner 30 mounted in the duct 10 adjacent the outlet 21. This burner may be of any suitable type and, as illustrated, comprises an annular dome or inlet cone 31 providing an inlet for primary combustion air and generally cylindrical combustion space defining walls 32 and 33 having suitable air inlets. Fuel nozzles 34 inject fuel for combustion within the burner 30. It will be noted that this burner is mounted approximately at the mean radius of the bypass duct 10. The bypass burner and the afterburner combustion arrangements represented by manifolds 27 and vaporizing flameholders 29 may be of any suitable known type and detailed description thereof is unnecessary.

The engine also includes a heat exchanger means comprising an annular array of heat exchange units 35 each. including an arcuate inlet manifold 37 supported from the inner wall 11 by a hollow strut 38 and an arcuate outlet manifold 39 supported from the inner wall on a hollow strut 41. There are preferably six such heat exchargers, each occupying substantially 60 of the circumference of duct 10. Struts 38 are connected by com pressed air lines 42 to a bleed air manifold 43 downstream from compressor 15. Thus, compressed air is supplied from the compressor through line 42 and strut 38 to the manifold 37. Manifold 37 is connected to manifold 39 by a relatively large number (about 150 in this embodiment) of generally parallel heat exchange tubes 45 preferably arranged in two closely adjacent rows, the tubes being disposed approximately on the mean radius of the duct 10, which diverges rearwardly of the engine. Thus the hot compressed air is conducted through these tubes 45 to manifold 39 and strut 41, and from strut 41 through an air tube 46, a fitting 47, and an air tube 49 into the turbine where it is circulated through the hot parts of the turbine by any suitable cooling arrangement, many of which are known in the art, and which need not be described here since the cooling circuit within the turbine is immaterial to my invention.

The air flowing in the bypass duct 10 flows along the tubes 45 in countercurrent relation to the compressed air and a considerable heating of the bypass air adjacent the mid-radius of the duct takes place. In a typical installation, this heating may amount to about 100 F. The warmer core of air enters the forward entrance of the burner 30 thus providing a more favorable environment for the combustion of fuel sprayed from each nozzle 34. The improvement of combustion is considerably more marked when the heating of the air is concentrated as far as feasible in the air which enters the duct burner for primary combustion. If the entire body of air flowing through the bypass duct is heated, the temperature rise will be much less and, therefore, the beneficial effect on combustion will be less.

The burners would not necessarily be annular and, if separate can-type burners are distributed around the axis of the engine, a heat exchanger 35 should be located as far as feasible in the direct line of flow to the primary air entrance of each. such combustion can. In the example described, the turbine cooling air will be cooled about 300 F. by heat exchange with the air in the bypass duct. This gives a very substantial increase in the cooling power of the air and permits a substantial reduction in the amount of air bled from the compressor for cooling purposes.

References herein to the low pressure compressor are intended to be generic to compressors and fans of single or plural stages.

It will be clear to those skilled in the art from the foregoing that the principles of my invention are adapted to significantl improve the operation and economy of ducted fan engines. The detailed description of the pre- 4 ferred embodiment of the invention is illustrative only, and various modifications may be made within the scope of the invention.

I claim:

1. A bypass type jet engine comprising, in combination, a low pressure compressor, an engine duct and a bypass duct supplied by the low pressure compressor, a high pressure compressor, combustion apparatus, and turbine means coupled to drive the compressors disposed in the engine duct, a jet nozzle supplied through the said ducts, a duct burner supplied with air through the bypass duct, a heat exchanger mounted in the bypass duct in the line of air flow to the said duct burner, means for bleeding hot compressed air from the high pressure compressor and circulating it through the heat exchanger to the turbine means for cooling the turbine means, so that the cooling air is cooled and the bypass air flow to the said burner is heated, thus improving combustion stability in the said burner.

2. An engine as defined in claim 1 in which the bypass duct is annular and surrounds the engine duct.

3. An engine as defined in claim 2 in which the heat exchanger and duct burner are mounted substantially at the mid-radius of the bypass duct.

4. An engine as defined in claim 3 in which the heat exchanger comprises numerous tubes extending longitudinally of the bypass duct.

5. A bypass type jet engine comprising, in combination, a low pressure compressor, an engine duct and a bypass duct supplied by the loW pressure compressor, a high pressure compressor, combustion apparatus, and turbine means coupled to drive the compressors disposed in the engine duct, an afterburner and a jet nozzle supplied through the said ducts, the afterburner including a duct burner supplied with air through the bypass duct, a heat exchanger mounted in the bpass duct in the line of air flow to the said duct burner, and means for bleeding hot compressed air from the high pressure compressor and circulating it through the heat exchanger to the turbine means for cooling the turbine means, so that the cooling air is cooled and the bypass air flow to the said burner is heated thus improving combustion stability in the said burner.

6. An engine as defined in claim 5 in which the bypass duct is annular and surrounds the engine duct.

7. An engine as defined in claim 6 in which the heat exchanger and duct burner are mounted substantially at the mid-radius of the bypass duct.

8. An engine as defined in claim 7 in which the heat exchanger comprises numerous tubes extending longitudinally of the bypass duct.

References Cited UNITED STATES PATENTS 2,504,181 4/ 1950 Constant 60-266 3,273,340 9/1966 Hull 60-3966 3,475,906 11/ 1969 Madelung 60-266 FOREIGN PATENTS 1,081,277 5/1960 Germany.

DOUGLAS HART, Primary Examiner US. Cl. X.R. 60-262, 266, 39.66

US3528250A 1969-04-16 1969-04-16 Bypass engine with afterburning and compressor bleed air heat exchanger in bypass duct Expired - Lifetime US3528250A (en)

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Cited By (37)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3775975A (en) * 1972-09-05 1973-12-04 Gen Electric Fuel distribution system
US3893297A (en) * 1974-01-02 1975-07-08 Gen Electric Bypass augmentation burner arrangement for a gas turbine engine
US4007892A (en) * 1971-07-15 1977-02-15 Tabor Alga M Aircraft flight method and apparatus for boosting an aircraft to a very high altitude and thereafter boosting the aircraft to a high rate of forward speed
US4120150A (en) * 1977-05-17 1978-10-17 The United States Of America As Represented By The Secretary Of The Air Force Compact fuel-to-air heat exchanger for jet engine application
US4137705A (en) * 1977-07-25 1979-02-06 General Electric Company Cooling air cooler for a gas turbine engine
US4170109A (en) * 1977-11-09 1979-10-09 United Technologies Corporation Thrust augmentor having swirled flows for combustion stabilization
US4170111A (en) * 1977-11-09 1979-10-09 United Technologies Corporation Thrust augmentor
US4187675A (en) * 1977-10-14 1980-02-12 The United States Of America As Represented By The Secretary Of The Air Force Compact air-to-air heat exchanger for jet engine application
US4254618A (en) * 1977-08-18 1981-03-10 General Electric Company Cooling air cooler for a gas turbofan engine
US4527390A (en) * 1977-10-25 1985-07-09 General Motors Corporation Low loss duct burner
WO1986002406A1 (en) * 1984-10-10 1986-04-24 Paul Marius A Gas turbine engine
US4991394A (en) * 1989-04-03 1991-02-12 Allied-Signal Inc. High performance turbine engine
US5003766A (en) * 1984-10-10 1991-04-02 Paul Marius A Gas turbine engine
US5414992A (en) * 1993-08-06 1995-05-16 United Technologies Corporation Aircraft cooling method
US5826424A (en) * 1992-04-16 1998-10-27 Klees; Garry W. Turbine bypass engines
US6134880A (en) * 1997-12-31 2000-10-24 Concepts Eti, Inc. Turbine engine with intercooler in bypass air passage
FR2866675A1 (en) * 2004-02-24 2005-08-26 Snecma Moteurs Afterburner device ignition performance improving method for turbofan engine, involves guiding fraction of primary flow in ignition zone so that zone temperature is greater than secondary flow temperature to favor ignition of afterburner
US20080095611A1 (en) * 2006-10-19 2008-04-24 Michael Ralph Storage Method and apparatus for operating gas turbine engine heat exchangers
JP2009541657A (en) * 2006-06-27 2009-11-26 エアバス フランスAirbus France Aircraft turbo-jet engine
US20100139288A1 (en) * 2008-12-10 2010-06-10 Pratt & Whitney Canada Corp. Heat exchanger to cool turbine air cooling flow
WO2010103229A3 (en) * 2009-03-13 2010-12-02 Price Induction Device for spraying a fluid using the air blast effect
US20110067407A1 (en) * 2009-09-23 2011-03-24 Snecma Flame-holder device comprising an arm support and a heat-protection screen that are in one piece
US20120128467A1 (en) * 2010-11-22 2012-05-24 Michael Anthony Ruthemeyer Integrated variable geometry flow restrictor and heat exchanger
US8266888B2 (en) * 2010-06-24 2012-09-18 Pratt & Whitney Canada Corp. Cooler in nacelle with radial coolant
US20130186102A1 (en) * 2012-01-25 2013-07-25 Honeywell International Inc. Gas turbine engine in-board cooled cooling air system
US20140161593A1 (en) * 2012-12-11 2014-06-12 United Technologies Corporation Turbine engine alignment assembly
WO2014149100A1 (en) * 2013-03-15 2014-09-25 Rolls-Royce North American Technologies, Inc. Heat exchanger integrated with a gas turbine engine and adaptive flow control
US20150128561A1 (en) * 2013-03-13 2015-05-14 Rolls-Royce North American Technologies, Inc. Three stream, variable area, vectorable nozzle
US9045998B2 (en) 2011-12-12 2015-06-02 Honeywell International Inc. System for directing air flow to a plurality of plena
US20150377135A1 (en) * 2014-06-30 2015-12-31 General Electric Company Method and system for radial tubular duct heat exchangers
US9267390B2 (en) 2012-03-22 2016-02-23 Honeywell International Inc. Bi-metallic actuator for selectively controlling air flow between plena in a gas turbine engine
US9388739B2 (en) 2012-05-02 2016-07-12 Pratt & Whitney Canada Corp. Air cooler system for gas turbine engines
US9562475B2 (en) 2012-12-19 2017-02-07 Siemens Aktiengesellschaft Vane carrier temperature control system in a gas turbine engine
US20170184027A1 (en) * 2015-12-29 2017-06-29 General Electric Company Method and system for compressor and turbine cooling
US9777963B2 (en) 2014-06-30 2017-10-03 General Electric Company Method and system for radial tubular heat exchangers
US9835380B2 (en) 2015-03-13 2017-12-05 General Electric Company Tube in cross-flow conduit heat exchanger
US10006369B2 (en) * 2015-06-18 2018-06-26 General Electric Company Method and system for radial tubular duct heat exchangers

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US20170059161A1 (en) * 2015-08-26 2017-03-02 United Technologies Corporation Apparatus and method for air extraction in a gas turbine engine

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2504181A (en) * 1950-04-18 Double compound independent rotor
DE1081277B (en) * 1958-08-16 1960-05-05 Johann Endres Dr Ing Two current combustion chamber for gas turbines, in particular gas turbine jet engines
US3224194A (en) * 1963-06-26 1965-12-21 Curtiss Wright Corp Gas turbine engine
US3273340A (en) * 1963-11-22 1966-09-20 Gen Electric Gas turbine powerplant having an extremely high pressure ratio cycle
US3475906A (en) * 1966-04-01 1969-11-04 Entwicklungsring Sued Gmbh Heat exchanger for aircraft air conditioning systems

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2504181A (en) * 1950-04-18 Double compound independent rotor
DE1081277B (en) * 1958-08-16 1960-05-05 Johann Endres Dr Ing Two current combustion chamber for gas turbines, in particular gas turbine jet engines
US3224194A (en) * 1963-06-26 1965-12-21 Curtiss Wright Corp Gas turbine engine
US3273340A (en) * 1963-11-22 1966-09-20 Gen Electric Gas turbine powerplant having an extremely high pressure ratio cycle
US3475906A (en) * 1966-04-01 1969-11-04 Entwicklungsring Sued Gmbh Heat exchanger for aircraft air conditioning systems

Cited By (58)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4007892A (en) * 1971-07-15 1977-02-15 Tabor Alga M Aircraft flight method and apparatus for boosting an aircraft to a very high altitude and thereafter boosting the aircraft to a high rate of forward speed
US3775975A (en) * 1972-09-05 1973-12-04 Gen Electric Fuel distribution system
US3893297A (en) * 1974-01-02 1975-07-08 Gen Electric Bypass augmentation burner arrangement for a gas turbine engine
US4120150A (en) * 1977-05-17 1978-10-17 The United States Of America As Represented By The Secretary Of The Air Force Compact fuel-to-air heat exchanger for jet engine application
US4137705A (en) * 1977-07-25 1979-02-06 General Electric Company Cooling air cooler for a gas turbine engine
FR2398884A1 (en) * 1977-07-25 1979-02-23 Gen Electric cooling system for engine gas turbine engine and thus obtained
US4254618A (en) * 1977-08-18 1981-03-10 General Electric Company Cooling air cooler for a gas turbofan engine
US4187675A (en) * 1977-10-14 1980-02-12 The United States Of America As Represented By The Secretary Of The Air Force Compact air-to-air heat exchanger for jet engine application
US4527390A (en) * 1977-10-25 1985-07-09 General Motors Corporation Low loss duct burner
US4170109A (en) * 1977-11-09 1979-10-09 United Technologies Corporation Thrust augmentor having swirled flows for combustion stabilization
US4170111A (en) * 1977-11-09 1979-10-09 United Technologies Corporation Thrust augmentor
WO1986002406A1 (en) * 1984-10-10 1986-04-24 Paul Marius A Gas turbine engine
US5003766A (en) * 1984-10-10 1991-04-02 Paul Marius A Gas turbine engine
US4991394A (en) * 1989-04-03 1991-02-12 Allied-Signal Inc. High performance turbine engine
US5826424A (en) * 1992-04-16 1998-10-27 Klees; Garry W. Turbine bypass engines
US5414992A (en) * 1993-08-06 1995-05-16 United Technologies Corporation Aircraft cooling method
US6134880A (en) * 1997-12-31 2000-10-24 Concepts Eti, Inc. Turbine engine with intercooler in bypass air passage
FR2866675A1 (en) * 2004-02-24 2005-08-26 Snecma Moteurs Afterburner device ignition performance improving method for turbofan engine, involves guiding fraction of primary flow in ignition zone so that zone temperature is greater than secondary flow temperature to favor ignition of afterburner
EP1577530A1 (en) * 2004-02-24 2005-09-21 Snecma Device and method for providing ignition to a post-combustor for bypass gas turbine
US20050257527A1 (en) * 2004-02-24 2005-11-24 Snecma Moteurs Method of improving the ignition performance of an after-burner device for a bypass turbojet, and an after-burner device of improved ignition performance
US7584615B2 (en) 2004-02-24 2009-09-08 Snecma Method of improving the ignition performance of an after-burner device for a bypass turbojet, and an after-burner device of improved ignition performance
US8739516B2 (en) * 2006-06-27 2014-06-03 Airbus Operations Sas Turboreactor for aircraft
US20090301057A1 (en) * 2006-06-27 2009-12-10 Airbus France Turboreactor for aircraft
JP2009541657A (en) * 2006-06-27 2009-11-26 エアバス フランスAirbus France Aircraft turbo-jet engine
EP1916399A2 (en) * 2006-10-19 2008-04-30 General Electric Company Heat exchanger assembly for a gas turbine engine
US20080095611A1 (en) * 2006-10-19 2008-04-24 Michael Ralph Storage Method and apparatus for operating gas turbine engine heat exchangers
EP1916399A3 (en) * 2006-10-19 2011-09-28 General Electric Company Heat exchanger assembly for a gas turbine engine
US8387362B2 (en) * 2006-10-19 2013-03-05 Michael Ralph Storage Method and apparatus for operating gas turbine engine heat exchangers
US20100139288A1 (en) * 2008-12-10 2010-06-10 Pratt & Whitney Canada Corp. Heat exchanger to cool turbine air cooling flow
US8181443B2 (en) 2008-12-10 2012-05-22 Pratt & Whitney Canada Corp. Heat exchanger to cool turbine air cooling flow
WO2010103229A3 (en) * 2009-03-13 2010-12-02 Price Induction Device for spraying a fluid using the air blast effect
CN102348510B (en) 2009-03-13 2014-04-16 价美公司 Device for spraying a fluid using the air blast effect
JP2012520415A (en) * 2009-03-13 2012-09-06 プライス インダクションPrice Induction How to use the air blast effect blowing fluid
RU2510296C2 (en) * 2009-03-13 2014-03-27 Прайс Индакшн Device to eject fluid by blown airflow
CN102348510A (en) * 2009-03-13 2012-02-08 价美公司 Device for spraying a fluid using the air blast effect
US20110067407A1 (en) * 2009-09-23 2011-03-24 Snecma Flame-holder device comprising an arm support and a heat-protection screen that are in one piece
US8266888B2 (en) * 2010-06-24 2012-09-18 Pratt & Whitney Canada Corp. Cooler in nacelle with radial coolant
US8961114B2 (en) * 2010-11-22 2015-02-24 General Electric Company Integrated variable geometry flow restrictor and heat exchanger
US20120128467A1 (en) * 2010-11-22 2012-05-24 Michael Anthony Ruthemeyer Integrated variable geometry flow restrictor and heat exchanger
US9045998B2 (en) 2011-12-12 2015-06-02 Honeywell International Inc. System for directing air flow to a plurality of plena
US20130186102A1 (en) * 2012-01-25 2013-07-25 Honeywell International Inc. Gas turbine engine in-board cooled cooling air system
US9243563B2 (en) * 2012-01-25 2016-01-26 Honeywell International Inc. Gas turbine engine in-board cooled cooling air system
US9267390B2 (en) 2012-03-22 2016-02-23 Honeywell International Inc. Bi-metallic actuator for selectively controlling air flow between plena in a gas turbine engine
US9388739B2 (en) 2012-05-02 2016-07-12 Pratt & Whitney Canada Corp. Air cooler system for gas turbine engines
US20140161593A1 (en) * 2012-12-11 2014-06-12 United Technologies Corporation Turbine engine alignment assembly
US9416682B2 (en) * 2012-12-11 2016-08-16 United Technologies Corporation Turbine engine alignment assembly
US9562475B2 (en) 2012-12-19 2017-02-07 Siemens Aktiengesellschaft Vane carrier temperature control system in a gas turbine engine
US20150128561A1 (en) * 2013-03-13 2015-05-14 Rolls-Royce North American Technologies, Inc. Three stream, variable area, vectorable nozzle
US9845768B2 (en) * 2013-03-13 2017-12-19 Rolls-Royce North American Technologies, Inc. Three stream, variable area, vectorable nozzle
US9587561B2 (en) 2013-03-15 2017-03-07 Rolls-Royce North American Technologies, Inc. Heat exchanger integrated with a gas turbine engine and adaptive flow control
WO2014149100A1 (en) * 2013-03-15 2014-09-25 Rolls-Royce North American Technologies, Inc. Heat exchanger integrated with a gas turbine engine and adaptive flow control
US20150377135A1 (en) * 2014-06-30 2015-12-31 General Electric Company Method and system for radial tubular duct heat exchangers
US9777963B2 (en) 2014-06-30 2017-10-03 General Electric Company Method and system for radial tubular heat exchangers
CN105222616A (en) * 2014-06-30 2016-01-06 通用电气公司 Method and system for radial tubular duct heat exchangers
EP2963248A1 (en) * 2014-06-30 2016-01-06 General Electric Company Method and system for radial tubular duct heat exchangers
US9835380B2 (en) 2015-03-13 2017-12-05 General Electric Company Tube in cross-flow conduit heat exchanger
US10006369B2 (en) * 2015-06-18 2018-06-26 General Electric Company Method and system for radial tubular duct heat exchangers
US20170184027A1 (en) * 2015-12-29 2017-06-29 General Electric Company Method and system for compressor and turbine cooling

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