US20050150970A1 - Cooling system for hot parts of an aircraft engine, and aircraft engine equipped with such a cooling system - Google Patents
Cooling system for hot parts of an aircraft engine, and aircraft engine equipped with such a cooling system Download PDFInfo
- Publication number
- US20050150970A1 US20050150970A1 US11/006,635 US663504A US2005150970A1 US 20050150970 A1 US20050150970 A1 US 20050150970A1 US 663504 A US663504 A US 663504A US 2005150970 A1 US2005150970 A1 US 2005150970A1
- Authority
- US
- United States
- Prior art keywords
- cooling system
- engine
- heat exchanger
- channel
- nacelle
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, 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
- F02C7/12—Cooling of plants
- F02C7/14—Cooling of plants of fluids in the plant, e.g. lubricant or fuel
- F02C7/141—Cooling of plants of fluids in the plant, e.g. lubricant or fuel of working fluid
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F28—HEAT EXCHANGE IN GENERAL
- F28D—HEAT-EXCHANGE APPARATUS, NOT PROVIDED FOR IN ANOTHER SUBCLASS, IN WHICH THE HEAT-EXCHANGE MEDIA DO NOT COME INTO DIRECT CONTACT
- F28D21/00—Heat-exchange apparatus not covered by any of the groups F28D1/00 - F28D20/00
- F28D21/0001—Recuperative heat exchangers
- F28D21/0014—Recuperative heat exchangers the heat being recuperated from waste air or from vapors
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/208—Heat transfer, e.g. cooling using heat pipes
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F28—HEAT EXCHANGE IN GENERAL
- F28D—HEAT-EXCHANGE APPARATUS, NOT PROVIDED FOR IN ANOTHER SUBCLASS, IN WHICH THE HEAT-EXCHANGE MEDIA DO NOT COME INTO DIRECT CONTACT
- F28D21/00—Heat-exchange apparatus not covered by any of the groups F28D1/00 - F28D20/00
- F28D2021/0019—Other heat exchangers for particular applications; Heat exchange systems not otherwise provided for
- F28D2021/0021—Other heat exchangers for particular applications; Heat exchange systems not otherwise provided for for aircrafts or cosmonautics
-
- Y—GENERAL 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
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- This invention relates to the technical field of cooling systems for hot parts of aircraft engines.
- a cooling system comprising a set of heat exchangers, to cool hot parts of an aircraft engine such as high pressure turbine blades in this aircraft engine.
- It also relates to an aircraft engine equipped with such a cooling system.
- heat exchangers can be installed in an aircraft engine to cool hot parts of an aircraft engine.
- Document FR 2 400 618 discloses a turbo-fan type of aircraft engine using an air/air type cooling system, and an associated cooling method.
- Hot parts such as fixed and mobile blades of the high pressure turbine are cooled by cooling air originating from part of the primary air drawn off at the outlet from the compressor or between compressor stages.
- This air cooling hot parts is itself firstly cooled before passing over the parts to be cooled, by passing inside pipes in a heat exchanger itself installed in a colder air current.
- This colder air current originates from part of the fan dilution air, or secondary air. It is drawn off from the fan duct, and more precisely in an annular flow passage delimited on one side by the gas generator and on the other side by a casing that surrounds part of the length of the gas generator.
- This part of the dilution air drawn off through the casing enters a diffuser section in which the dynamic air pressure is largely recovered, and is then transferred through the heat exchanger where it absorbs heat from cooling air drawn off from the compressor. Then once it has cooled the air that cools hot parts, this past of the dilution air is returned to the fan duct, its static pressure being adjusted to the static pressure existing in the fan duct.
- a first disadvantage is related to engine maintenance, particularly of the gas generator. If there is no cooling system, maintenance is done by opening a cowling of the nacelle to access the engine directly, and particularly the fuel injectors. In the presence of a cooling system according to prior art, and particularly if the casing is in position around the gas generator or a channel fixed to the nacelle, it becomes difficult to access some parts of the engine for maintenance.
- a second disadvantage of systems according to prior art is due to the fact that colder air used to cool fluid circulating in the heat exchanger is restored on the downstream side inside the fan duct. Therefore its pressure is adjusted to the pressure inside the fan duct at this location. There is then a risk that air could circulate in the inverse direction or not circulate at all, which would make the heat exchanger inoperative.
- This invention is intended to provide a solution to the disadvantages of the systems of prior art.
- the cooling system for hot parts of an aircraft engine is applicable to an aircraft engine housed in a nacelle, a primary air flow passing inside the engine and a secondary air flow passing around the engine inside the nacelle.
- the cooling system comprises at least one channel that draws off cold air in the secondary air flow and at least one heat exchanger located in the channel and in which hot air from the primary air flow circulates to be cooled before reaching hot parts to cool them.
- Said at least, one channel comprises the following three parts:
- the intermediate box has a longitudinal section with an approximately rectangular shaped profile.
- the intermediate box has a longitudinal section with a profile approximately in the shape of a trapezium, in which the large base is facing the engine and the small base is facing the nacelle.
- the cooling system also comprises an upstream seal between the supply pipe and the intermediate box and a downstream seal between the intermediate box and the evacuation pipe.
- each heat exchanger is associated with:
- each inlet duct comprises an attachment flange at one of its ends for fixing it onto the engine.
- each return duct comprises an attachment flange at one of its ends for fixing it to the engine. All these attachment flanges make a mechanical attachment of the box onto the engine.
- At least one of the inlet ducts is provided with a valve.
- At least one of the return ducts is provided with a valve.
- each heat exchanger is associated with four inlet ducts and four return ducts.
- the outlet from the evacuation pipe of each channel opens up at the outlet from the nacelle ejection nozzle, or beyond it on the downstream side.
- air that exits from the evacuation pipe is at atmospheric pressure. Consequently, the risk of recirculation of air in the heat exchanger is eliminated, and pressure losses in the heat exchanger are increased.
- the outlet section of the supply pipe of each channel is larger than its inlet section.
- the outlet section of the evacuation pipe of each channel is smaller than its inlet section.
- the cooling system comprises at least two heat exchangers distributed circumferentially around the engine, each heat exchanger being placed in a separate channel. Even more preferably, there are four heat exchangers.
- the cooling system comprises a single heat exchanger extending around the entire circumference of the engine and is placed in a single corresponding annular channel.
- the invention relates to an aircraft engine equipped with a cooling system according to the first aspect of the invention.
- FIG. 1 shows a diagrammatic sectional view through an aircraft engine comprising a cooling system according to the invention
- FIG. 2 is an aft perspective view showing the aft side of an aircraft engine comprising a cooling system according to the invention.
- FIG. 3 shows another aft perspective view of the engine, an aft part having been removed at the location of the cooling system, showing a section through the layout of the heat exchangers.
- FIG. 1 diagrammatically shows an aircraft engine 10 with an axis of revolution 12 .
- the aircraft engine 10 comprises low pressure compressor stages 14 , medium pressure compressor stages 16 , high pressure compressor stages 18 , a combustion chamber 20 and turbines 22 .
- the aircraft engine 10 is surrounded by a nacelle 24 that terminates with an ejection nozzle 26 .
- a primary air flow represented by arrows 100 circulates inside the engine 10 . It is heated as it passes through the compressor stages 14 , 16 and 18 before arriving at the turbines 22 .
- a secondary air flow represented by the arrows 200 circulates in the nacelle 24 around the engine 10 .
- This secondary air 200 external to the engine 10 is colder than the primary air 100 inside the engine 10 .
- the engine 10 comprises a cooling system 30 designed to cool hot parts of the engine 10 , for example such as the turbine blades 22 .
- the principle of this cooling system consists of cooling the air that then flows on or in hot parts to be cooled.
- the cooling system 30 comprises at least one channel 32 inside which air circulates (arrows 300 , 400 ) drawn off from the secondary air flow of the nacelle 24 and at least one heat exchanger 34 placed in this channel 32 .
- this heat exchanger may for example be a tube exchanger or a plate exchanger.
- Each channel 32 comprises three successive parts:
- junction 342 between the supply pipe 322 and the intermediate box 324 is made by assuring continuity between the outlet section 323 of the supply pipe 322 and the inlet section 328 of the intermediate box 324 , which have approximately the same dimensions for this purpose.
- This junction 342 comprises a seal (not shown) on the upstream side that fits to said sections.
- junction 346 between the intermediate box 324 and the evacuation pipe 326 is made by assuring continuity between the outlet section 329 of the intermediate box 324 and the inlet section 325 of the evacuation pipe 326 , which have approximately the same dimensions for this purpose.
- This junction 346 comprises a seal (not shown) on the upstream side that fits to said sections.
- the end sections of the intermediate box 324 at the corresponding junction 342 , 346 have a profile approximately in the shape of a trapezium when viewing the longitudinal section, for which the large base is facing the engine 10 and the small base is facing the nacelle 24 .
- the inlet 321 to the channel 32 at the inlet to the supply pipe 322 may be a static air inlet or a dynamic air inlet.
- the outlet 327 from the channel 32 that is the outlet from the evacuation pipe 326 is arranged so that it coincides approximately with the free end of the ejection nozzle from the nacelle 24 .
- air evacuated through the evacuation pipe 326 is at atmospheric pressure.
- the heat exchanger 34 is located in the intermediate box 324 of the channel 32 .
- This intermediate box 324 is facing the engine 10 at a certain distance from it. It is connected to the engine 10 using at least one inlet duct 42 that is used to bring in (arrow 420 ) air drawn off at the outlet from the compressor 14 , 16 , 18 to the heat exchanger 34 , so that this air is cooled, and at least one return duct 44 that is used to transfer this cooled air to the turbines 22 (arrow 440 ).
- FIGS. 2 and 3 illustrate an example cooling system 30 according to the invention in more detail, including four channels 32 distributed around the periphery of the nacelle 24 .
- the engine 10 is not shown in these figures.
- the arrows 200 indicate the secondary air flow direction, and therefore indicate the upstream and downstream sides of the channels 32 .
- FIG. 2 illustrates an example cooling system according to the invention more particularly showing the channels 32 and their supply pipes 322 and evacuation pipes 326 .
- the inlet section 321 of the supply pipes 322 is smaller than their outlet section 323 .
- air (arrows 300 ) originating from the nacelle 24 is cooled in supply pipes 322 before reaching the heat exchangers 34 where it is heated on contact with them.
- the inlet section 325 of the evacuation pipes 326 is identical to their outlet section 327 .
- air (arrows 400 ) that was heated in contact with the heat exchangers 34 is not accelerated in the evacuation pipes, before exiting towards the outside, so as to limit pressure losses.
- FIG. 3 only shows the supply pipes 322 and the intermediate boxes 324 of the channels 32 .
- This figure shows intermediate boxes 324 in which the heat exchangers 34 are located in more detail, together with inlet ducts 42 and return ducts 44 .
- inlet ducts 42 there are four inlet ducts 42 per intermediate box 324 arranged on the upstream side of the intermediate boxes 324 .
- return ducts there are four return ducts per intermediate box 324 located on the downstream side of the intermediate boxes 324 .
- the inlet ducts 42 and the return ducts 44 terminate on the engine side with attachment flanges 43 that are provided for fixing together the intermediate boxes 324 to said engine 10 .
- the inlet ducts 42 and the return ducts 44 are also used as attachment means for the intermediate boxes 324 , and therefore for the associated heat exchangers 34 on the engine 10 .
- the inlet ducts 42 are fixed on each intermediate box 324 at a distributor 46 that is used to supply all tubes or all plates of the heat exchanger(s) 34 located in the intermediate box 324 . Air that has just been cooled in contact with the heat exchanger(s) 34 is collected by a header 48 on which the return ducts 44 are fixed.
- the inlet ducts 42 are approximately straight and draw air in directly at the exit from the compressor 14 , 16 , 18 to bring it to the heat exchangers 34 .
- the return pipes 44 are bent so as to return air that passed through the heat exchangers 34 , to a point further downstream at the inlet to turbines 22 .
- the cooling system that has just been described has a number of advantages.
- a first advantage is related to the maintenance of some engine parts 10 , for example such as fuel injectors (not shown).
- the supply pipe 322 and the evacuation pipe 326 for each channel 32 are fixed to the nacelle 24 , while the intermediate box 324 is fixed to the engine 10 .
- these two pipes 322 , 326 are lifted at the same time as the cowling, while the intermediate box 324 containing one or several heat exchangers 34 remains fixed to the engine 10 . Therefore, it is easy for an operator to access the engine even in the presence of a channel 32 , due to the fact that the channel is made from three separate parts 322 , 324 , 326 .
- the intermediate box 324 Since the intermediate box 324 is also held at a distance from the engine 10 due to the presence of the inlet ducts 42 and the return ducts 44 , it becomes easy for an operator to access parts of the engine, even in an area located under the intermediate box 324 itself, by passing between the inlet ducts 42 and the return ducts 44 . Consequently, all that is necessary to obtain easy access to fuel injectors for maintenance, is to put the intermediate box 324 into position facing the combustion chamber 20 .
- Another advantage is related to the trapezoidal shape of the profile of the intermediate box 324 .
- the result of this shape is that when the cowling of the nacelle 24 is closed, the outlet section 323 of the supply pipe, 322 of each channel 32 covers the inlet section 328 of the intermediate box 324 .
- the inlet section 325 of the evacuation pipe 326 covers the outlet section 329 of the intermediate box 324 .
- This arrangement makes it possible to make the supply pipe 322 coincide well with the evacuation pipe 326 , and with the intermediate box 324 when the cowling is closed, and therefore assure a good seal of the channel 32 at junctions 342 and 346 .
- This seal can be further improved by the presence of seals around the periphery of junctions 342 , 346 .
- the cooling system that has just been described is a passive cooling system, in other words the air flow drawn off at the high pressure compressor 18 (arrow 420 ) is proportional to the cooling air flow in the high pressure turbine 22 (arrows 440 ).
- an active cooling system that includes at least one valve (not shown) placed on the inlet ducts 42 or on the return ducts 44 . By controlling opening and/or closing of these valves, it would then be possible to improve engine flight performances at the detriment of an increase in its mass, as a function of the different flight phases.
- each heat exchanger is associated with four inlet ducts and four outlet ducts. Without departing from the scope of the invention, it would be possible that the number of inlet ducts and the number of outlet ducts is different from four, and/or different from each other.
- the outlet from the channel coincides with the free end of the nacelle ejection nozzle. It would be possible to envisage the channel going beyond the free end of the ejection nozzle, in the downstream direction.
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Thermal Sciences (AREA)
- Physics & Mathematics (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
- Heat-Exchange Devices With Radiators And Conduit Assemblies (AREA)
- Cooling, Air Intake And Gas Exhaust, And Fuel Tank Arrangements In Propulsion Units (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Exhaust Silencers (AREA)
- Supercharger (AREA)
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
FR0450075 | 2004-01-13 | ||
FR0450075A FR2864996B1 (fr) | 2004-01-13 | 2004-01-13 | Systeme de refroidissement de parties chaudes d'un moteur d'aeronef, et moteur d'aeronef equipe d'un tel systeme de refroidissement |
Publications (1)
Publication Number | Publication Date |
---|---|
US20050150970A1 true US20050150970A1 (en) | 2005-07-14 |
Family
ID=34610801
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US11/006,635 Abandoned US20050150970A1 (en) | 2004-01-13 | 2004-12-08 | Cooling system for hot parts of an aircraft engine, and aircraft engine equipped with such a cooling system |
Country Status (8)
Country | Link |
---|---|
US (1) | US20050150970A1 (ru) |
EP (1) | EP1555406A1 (ru) |
JP (1) | JP2005201264A (ru) |
CN (1) | CN1624307A (ru) |
CA (1) | CA2491935A1 (ru) |
FR (1) | FR2864996B1 (ru) |
RU (1) | RU2362895C2 (ru) |
UA (1) | UA86348C2 (ru) |
Cited By (30)
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US20080095611A1 (en) * | 2006-10-19 | 2008-04-24 | Michael Ralph Storage | Method and apparatus for operating gas turbine engine heat exchangers |
GB2445237A (en) * | 2006-12-27 | 2008-07-02 | Gen Electric | Gas turbine engine having a cooling-air nacelle-cowl duct |
EP2011988A2 (en) * | 2007-07-06 | 2009-01-07 | General Electric Company | Heat exchanger for a turbine engine |
US20090301057A1 (en) * | 2006-06-27 | 2009-12-10 | Airbus France | Turboreactor for aircraft |
US20150152789A1 (en) * | 2013-12-03 | 2015-06-04 | United Technologies Corporation | Multi-bypass stream gas turbine engine with enlarged bypass flow area |
US9267434B2 (en) | 2012-01-29 | 2016-02-23 | United Technologies Corporation | Heat exchanger |
US20160230595A1 (en) * | 2015-02-06 | 2016-08-11 | United Technologies Corporation | Heat exchanger system with spatially varied additively manufactured heat transfer surfaces |
CN106014685A (zh) * | 2016-05-30 | 2016-10-12 | 西北工业大学 | 一种双s弯发动机喷管结构 |
US9562477B2 (en) | 2013-06-18 | 2017-02-07 | Rolls-Royce Deutschland Ltd & Co Kg | Accessory mounting for a gas turbine engine |
US20170268426A1 (en) * | 2016-03-18 | 2017-09-21 | United Technologies Corporation | Heat exchanger suspension system with pipe-to-linkage spring rate ratio |
US9777963B2 (en) | 2014-06-30 | 2017-10-03 | General Electric Company | Method and system for radial tubular heat exchangers |
US20170328280A1 (en) * | 2016-05-16 | 2017-11-16 | Rolls-Royce Plc | Heat sink |
US9835380B2 (en) | 2015-03-13 | 2017-12-05 | General Electric Company | Tube in cross-flow conduit heat exchanger |
US9995175B2 (en) | 2016-06-29 | 2018-06-12 | General Electric Company | System and method for gas bearing support of turbine |
US10006369B2 (en) | 2014-06-30 | 2018-06-26 | General Electric Company | Method and system for radial tubular duct heat exchangers |
US10066550B2 (en) | 2014-05-15 | 2018-09-04 | Rolls-Royce North American Technologies, Inc. | Fan by-pass duct for intercooled turbo fan engines |
US10167740B2 (en) | 2013-08-09 | 2019-01-01 | Safran Aircraft Engines | Turbine engine having an element for deflecting foreign objects |
US10247017B2 (en) | 2016-06-29 | 2019-04-02 | General Electric Company | System and method for gas bearing support of turbine |
US10378835B2 (en) | 2016-03-25 | 2019-08-13 | Unison Industries, Llc | Heat exchanger with non-orthogonal perforations |
US10384793B2 (en) | 2014-03-13 | 2019-08-20 | Safran Aircraft Engines | Nacelle comprising an exchanger for cooling a stream of fluid |
US10508598B2 (en) * | 2014-01-15 | 2019-12-17 | United Technologies Corporation | Cooling systems for gas turbine engines |
US10676205B2 (en) | 2016-08-19 | 2020-06-09 | General Electric Company | Propulsion engine for an aircraft |
US10800539B2 (en) * | 2016-08-19 | 2020-10-13 | General Electric Company | Propulsion engine for an aircraft |
US11105340B2 (en) | 2016-08-19 | 2021-08-31 | General Electric Company | Thermal management system for an electric propulsion engine |
CN113544456A (zh) * | 2019-03-01 | 2021-10-22 | 利勃海尔-航空航天图卢兹有限公司 | 用于通过冷次级空气冷却热初级空气的交换器,和设有这样的交换器的空调系统 |
US11156128B2 (en) | 2018-08-22 | 2021-10-26 | General Electric Company | Embedded electric machine |
US11300368B2 (en) | 2013-11-18 | 2022-04-12 | General Electric Company | Monolithic tube-in matrix heat exchanger |
US11572834B2 (en) | 2020-02-11 | 2023-02-07 | Rolls-Royce Plc | Gas turbine engine cooling system |
US11684974B2 (en) | 2014-10-21 | 2023-06-27 | Raytheon Technologies Corporation | Additive manufactured ducted heat exchanger system |
US11781506B2 (en) | 2020-06-03 | 2023-10-10 | Rtx Corporation | Splitter and guide vane arrangement for gas turbine engines |
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US7861512B2 (en) * | 2006-08-29 | 2011-01-04 | Pratt & Whitney Canada Corp. | Turbofan bypass duct air cooled fluid cooler installation |
FR2913063B1 (fr) | 2007-02-27 | 2012-03-16 | Snecma | Moteur d'aeronef equipe de moyens d'echange thermiques. |
FR2914365B1 (fr) * | 2007-03-28 | 2012-05-18 | Airbus France | Systeme de refroidissement et de regulation en temperature d'equipements d'un ensemble propulsif d'aeronef. |
EP2527252A3 (en) * | 2011-05-27 | 2018-01-10 | General Electric Company | Adaptive power and thermal management system |
US8789376B2 (en) * | 2011-05-27 | 2014-07-29 | General Electric Company | Flade duct turbine cooling and power and thermal management |
CN103899364B (zh) * | 2012-12-26 | 2015-12-02 | 中航商用航空发动机有限责任公司 | 航空发动机高压涡轮的轮缘密封结构、高压涡轮及发动机 |
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CN117346561B (zh) * | 2023-09-12 | 2024-04-19 | 贵州永红航空机械有限责任公司 | 一种高效环形散热器及换热方法 |
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-
2004
- 2004-01-13 FR FR0450075A patent/FR2864996B1/fr not_active Expired - Lifetime
- 2004-12-08 US US11/006,635 patent/US20050150970A1/en not_active Abandoned
-
2005
- 2005-01-06 JP JP2005001317A patent/JP2005201264A/ja active Pending
- 2005-01-06 CA CA002491935A patent/CA2491935A1/en not_active Abandoned
- 2005-01-11 EP EP05100120A patent/EP1555406A1/fr not_active Withdrawn
- 2005-01-12 UA UAA200500294A patent/UA86348C2/ru unknown
- 2005-01-12 RU RU2005100708/06A patent/RU2362895C2/ru active
- 2005-01-12 CN CNA2005100020379A patent/CN1624307A/zh active Pending
Patent Citations (10)
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US3842597A (en) * | 1973-03-16 | 1974-10-22 | Gen Electric | Gas turbine engine with means for reducing the formation and emission of nitrogen oxides |
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Cited By (45)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
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Also Published As
Publication number | Publication date |
---|---|
CA2491935A1 (en) | 2005-07-13 |
RU2362895C2 (ru) | 2009-07-27 |
UA86348C2 (ru) | 2009-04-27 |
CN1624307A (zh) | 2005-06-08 |
EP1555406A1 (fr) | 2005-07-20 |
FR2864996A1 (fr) | 2005-07-15 |
FR2864996B1 (fr) | 2006-03-10 |
JP2005201264A (ja) | 2005-07-28 |
RU2005100708A (ru) | 2006-06-20 |
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