• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    A propulsion unit (10) for an aircraft, comprising a motor (16), a nacelle (18) surrounding the engine, and a fire extinguishing system that can appear in the engine and / or in the nacelle. extinguishing means (34) supplying extinguishing agent of at least one pipe (36) for dispensing said extinguishing agent, which opens into a cavity (32) of the engine and / or a cavity (26) of the nacelle , characterized in that it further comprises means (48) for supplying air to said at least one pipe for the purpose of ventilating the or each cavity.
    公开号:FR3015568A1
    申请号:FR1363544
    申请日:2013-12-24
    公开日:2015-06-26
    发明作者:Pierrick Charlemagne；Delphine Leroux
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    [0001] TECHNICAL FIELD The present invention relates to the field of ventilation of an aircraft propulsion unit.
    [0002] STATE OF THE ART An aircraft propulsion unit comprises a motor and a nacelle surrounding this engine which is generally a turbomachine. The nacelle comprises a housing of revolution delimiting an annular flow stream of a flow of air around the engine, this air flow being called secondary flow in the case of a turbomachine with a double flow. The nacelle defines around the housing a first annular cavity. Some equipment of the propulsion unit are mounted in the nacelle, that is to say in the aforementioned annular cavity, and can be fixed on its housing. The outer wall of the nacelle generally comprises removable covers to allow access to these devices during a maintenance operation. The engine comprises an internal annular vein of flow of an air flow called primary flow in the case of a turbomachine with a double flow. The engine comprises coaxial revolution housings which are surrounded by a wall of revolution which delimits internally the vein of the secondary flow. This wall extends away from the engine casings and defines around them a second annular cavity. Some equipment is mounted in this second annular cavity. The equipment mounted in the cavities of the propulsion unit are more or less sensitive to heat and are ventilated during operation. This is particularly the case of a computer or onboard computer type EEC, which allows in particular to control engine actuators to optimize the performance of the turbomachine (FR-B1-2 960 912). This computer is usually mounted with other equipment (Accessory Gear Box (AGB), exchangers, etc.) in the basket.
    [0003] In order to ventilate the internal cavity of the nacelle, the latter comprises an air intake scoop in flight, the air drawn is then evacuated by an air outlet grille of the nacelle. However, on the ground, this ventilation is almost non-existent and the natural convection in the annular space of the nacelle may be insufficient to ensure the ventilation of its equipment. The computer releases a significant thermal power in operation that must be dissipated, whether the engine is running or is stopped. In addition, even when the engine is stopped after operation, the hot parts of the engine continue to radiate and heat colder peripheral parts of the engine, which can thus, at the engine stop, reach temperatures close or even higher than their temperatures when the engine is running. There is therefore a real need for a system capable of ventilating this type of cavity of a propulsion unit, even when the engine is stopped. Furthermore, an aircraft propulsion unit is equipped with a fire extinguishing system that can appear in the engine and / or in the nacelle. This extinguishing system comprises means for supplying extinguishing agent of at least one extinguishing agent distribution pipe, which opens into a cavity of the engine and / or a cavity of the nacelle. This extinguishing system is generally associated with a fire detection system (of the FDU type, of the English Fire Detection Unit) which comprises sensors mounted on the engine and / or the nacelle and which is intended to emit a signal warning to the pilot of the aircraft when at least one of the sensors detects the appearance of a fire. In the current technique, the channeling of the extinguishing system is only for the distribution of the extinguishing agent. As the occurrence of a fire is rare, this pipe is usually never used. However, it is still present in an aircraft propulsion unit for reasons of safety and certification.
    [0004] The present invention provides a simple, effective and economical solution to the aforementioned need of the prior art. SUMMARY OF THE INVENTION The invention proposes an aircraft propulsion unit comprising an engine, a nacelle surrounding the engine, and a fire extinguishing system that can appear in the engine and / or in the nacelle. extinguishing system comprising means for supplying extinguishing agent of at least one distribution pipe for said extinguishing agent, which opens into a cavity of the engine and / or a cavity of the nacelle, characterized in that it further comprises means for supplying air to said at least one pipe for the purpose of ventilating the cavity or cavities. The invention therefore consists in assigning to a known means a new additional function. Indeed, the distribution channel of the extinguishing agent of the fire extinguishing system is, as in the prior art, used to distribute this agent in the cavity or cavities of the propulsion unit in the event that a fire appears. According to the invention, this pipe also serves to ventilate this or these cavities. The pipe is associated with air supply means that delivers a flow of air to the pipe, this air flow is then routed through the pipe to the cavity or cavities to ventilate. The invention thus makes it possible to use an existing means (ducting) for conveying ventilation air to a cavity of the nacelle and / or the engine, this existing means being used for a very different thing in the prior art. (extinguishing agent distribution). As indicated above, the extinguishing system of a propulsion system is rarely used. Its pipe may be to ventilate the cavities of the propulsion unit, even when the engine is stopped. Indeed, the operation of the air supply means of the pipe are advantageously independent of the engine and can therefore operate when the engine stops. The equipment mounted in the cavities of the propulsion system can be ventilated when the engine is stopped, which increases the life of this equipment. The air supply means preferably comprise a fan. This fan can be an electric fan.
    [0005] The supply means may be connected by a valve and / or a non-return valve to the at least one pipe. They can be connected to this pipe by a derivation Y for example. The feeding means can be housed in the nacelle. Alternatively, they can be housed in a connecting pylon of the propulsion system to the aircraft. The connection between a propulsion unit and its pylon generally comprises a firewall adapted to limit the transmission of a fire that would occur in the propulsion system. The supply means are thus protected from a fire that could appear in the nacelle.
    [0006] The supply means can be configured to deliver an air flow rate of between 1 and 500 g / s. Advantageously, the pipe may comprise at least one outlet which is located near an air inlet scoop of the nacelle or the engine.
    [0007] The present invention also relates to a method for ventilating a cavity of an engine and / or a cavity of a nacelle of an aircraft propulsion unit, this propulsion unit comprising a fire extinguishing system. which may appear in the engine and / or in the nacelle, this extinguishing system comprising means for supplying extinguishing agent with at least one distribution pipe for said extinguishing agent which opens into the engine cavity and / or the engine cavity. the nacelle, characterized in that it consists in supplying air to said at least one pipe. DESCRIPTION OF THE FIGURES The invention will be better understood and other details, characteristics and advantages of the invention will become apparent on reading the following description given by way of nonlimiting example and with reference to the appended drawings in which: FIG. 1 is a diagrammatic side view of an aircraft propulsion assembly; FIG. 2 is a very schematic front view of an aircraft propulsion assembly; FIG. 3 is a view corresponding to FIG. 2 and illustrating FIG. an embodiment of the invention; and FIG. 4 is a view corresponding to FIG. 2 and illustrating an alternative embodiment of the invention. DETAILED DESCRIPTION Reference is first made to FIG. 1 which represents a propulsion unit 10 of an aircraft, this propulsion unit comprising a pylon 12 connecting to a structural part of the aircraft, which here is a wing 14 of the aircraft. aircraft. The propulsion unit 10 comprises a motor 16 of the turbomachine type surrounded by a nacelle 18, the nacelle 18 defining a first annular stream 20 of flow of a secondary flow around the engine which comprises a second internal vein (not shown) of flow of a primary flow. The motor 16 typically comprises, from upstream to downstream, in the flow flow direction, a fan, at least one compression module, a combustion chamber, at least one turbine module, and a gas ejection nozzle. of combustion.
    [0008] The blower motor 16 is surrounded by a housing 22 of the nacelle 18, which is itself surrounded by a revolution wall 24 of the nacelle. Equipment is mounted in the annular cavity 26 delimited by the outer wall 24 and the casing 22 of the nacelle 18. The compression module, the combustion chamber, and the turbine module 30 comprise outer casings 28 which are surrounded by a wall 30. The wall 30 internally defines the vein 20 of the secondary flow and extends away from the outer casings 28 to define therewith an annular cavity 32 in which are also housed equipment. The propulsion unit 10 further comprises a fire extinguishing system that can appear in the engine 16 and / or in the nacelle 18. This extinguishing system comprises means 34 for supplying extinguishing agent of at least a pipe 36 for distributing the extinguishing agent. As can be seen in the drawing, this duct 36 may comprise several inputs (here two in number) which are each connected to supply means 34, and several outputs 38 (here two in number). In the example shown, the duct 36 comprises a first outlet 38 in the cavity 26 of the nacelle 18 and a second outlet 40 in the cavity 32 of the motor 16. Each outlet 38, 40 here has a Y shape and comprises two orifices. discharge outlet for the projection of extinguishing agent in directions substantially tangential to the housing 22 or housings 28. The outlet 38 is preferably located near an air inlet scoop of the nacelle 16 (for the ventilation of the cavity 26 in operation of the engine) and the outlet 40 is preferably located near an air intake scoop of the engine (for ventilation of the cavity 32 in operation of the engine). The outlets 38 and 40 are oriented to cooperate respectively with the air intake scoops of the nacelle 18 and the motor 16 to generate flows in similar directions. This prevents the extinguishing agent from leaving the areas to extinguish before filling. In addition, with the invention, the aeration by the scoop or the invention is similar and simplifies the management of ventilation flows. The pipe 36 is made of a fire-resistant material and in particular at very high temperatures (for example higher than 1000 ° C.).
    [0009] The supply means 34 may comprise a reservoir of extinguishing agent (such as Halon for example) under pressure, an outlet of which is connected to an inlet of the pipe 36 by a priming system such as a pyrotechnic initiator. This priming system is controlled remotely by the pilot of the aircraft, from the cockpit of the aircraft. The extinguishing system is associated with a fire detection system 42 (for example of the FDU type) which is connected to sensors 44 mounted on the engine 16 and the nacelle 18 and which are each intended to emit an alert signal. for the attention of the pilot of the aircraft when at least one of the sensors 44 detects the appearance of a fire. The sensors 44 comprise for example thermocouples. As is schematically represented in FIG. 2, the supply means 34 are generally mounted in the perimeter of the aircraft, this perimeter being here schematically delimited by the dashed lines 46 and including the pylon 12. The duct 36 extends from the supply means 34 to the cavities 26, 32. The duct 36 must therefore pass through the vein of the secondary flow 20 and may for this be housed in a tubular arm of servitudes passage of an intermediate casing of the engine. The dashed lines 46 symbolize a firewall to be crossed, it is therefore necessary at this point the minimum channel to limit the holes in this wall. Referring now to Figure 3 which shows a first embodiment of the invention.
    [0010] According to the invention, means are provided for supplying air to the duct 36, this air being intended to be conveyed through the duct 36 to the cavities 26, 32 for ventilation. In the example shown, the air supply means comprise an electric fan 48 whose air outlet 50 is connected to the pipe 36, in the vicinity of its inlet 52 connected to the supply means 34. by means of a Y-branch (preferably upstream of the fire-resistant wall) of which one of the lateral branches is connected to the fan 48, the other lateral branch is connected to the supply means 34, and the middle branch is connected to the cavities 26, 32. The connection between the fan 48 and the pipe 36 may comprise a solenoid valve or a non-return valve. The fan 48 can be configured to deliver an air flow rate of between 1 and 500 g / s. The fan 48 is preferably electrically controlled by independent control means of the motor 16 so that it can operate at engine stop. These control means are for example integrated into the electronic network of the aircraft. In the example shown in FIG. 3, the fan 48 is housed in the pylon 12. The variant embodiment of the invention shown in FIG. 4 differs from the embodiment described in the foregoing essentially in that the fan 48 is housed in the nacelle 18, that is to say in the annular cavity 26 of the nacelle. The fan 48 may be electrically controlled by engine control means which are configured to operate and remain operational at the engine stop.
    [0011] As shown in FIG. 1, the invention can be applied to a propulsion unit 10 fixed to a wing 14 of an aircraft, the pylon 12 then being situated at 12 o'clock (twelve hours) by analogy with the dial of a clock . Alternatively and as shown in Figures 3 and 4, the invention can be applied to a propulsion unit 10 attached to the fuselage of the aircraft, the pylon 12 then being for example located at 3 o'clock or 9 o'clock. It is also conceivable that the invention is applied to other types of propulsion units such as those at least partially buried in the fuselage of the aircraft.
    权利要求:
    Claims (9)
    [0001]
    REVENDICATIONS1. A propulsion unit (10) for an aircraft, comprising a motor (16), a nacelle (18) surrounding the engine, and a fire extinguishing system that can appear in the engine and / or in the nacelle. extinguishing means (34) supplying extinguishing agent of at least one pipe (36) for dispensing said extinguishing agent, which opens into a cavity (32) of the engine and / or a cavity (26) of the nacelle , characterized in that it further comprises means (48) for supplying air to said at least one pipe for the purpose of ventilating the cavity or cavities.
    [0002]
    2. propulsion unit (10) according to claim 1, characterized in that the air supply means comprise a fan (48).
    [0003]
    3. Propulsion unit (10) according to claim 2, characterized in that the fan (48) is an electric fan.
    [0004]
    4. propulsion unit (10) according to one of the preceding claims, characterized in that the means (48) for supplying air are connected by a valve and / or a non-return valve to said at least one pipe (36). ).
    [0005]
    5. Propulsion unit (10) according to one of the preceding claims, characterized in that the means (48) for supplying air are housed in the nacelle (18).
    [0006]
    6. propulsion unit (10) according to one of claims 1 to 4, characterized in that it further comprises a pylon (12) for connecting the propulsion unit to the aircraft, said means (48) of air supply being housed in the pylon.
    [0007]
    7. Propulsion unit (10) according to one of the preceding claims, characterized in that the means (48) for supplying air are configured to deliver an air flow rate of between 1 and 500g / s.
    [0008]
    8. Propulsion unit (10) according to one of the preceding claims, characterized in that the pipe (36) comprises at least one outlet (38, 40) which is located near an air intake scoop of the nacelle (18) or the engine (16).
    [0009]
    9. A method of ventilating a cavity (32, 26) of an engine (16) and / or a cavity of a nacelle (18) of an aircraft propulsion unit (10), this propulsion unit including a fire extinguishing system which may appear in the engine and / or in the nacelle, this extinguishing system comprising means (34) for supplying extinguishing agent of at least one distribution pipe (36) said extinguishing agent, which opens into the engine cavity and / or the nacelle cavity, characterized in that it comprises supplying air to said at least one pipe.
    类似技术:
    公开号 | 公开日 | 专利标题
    EP3087265B1|2018-02-07|Aircraft propulsion assembly with fire extinguishing system and ventilation method
    CA2581540C|2014-07-08|Turbine engine inlet cone deicing system for aircraft
    CA2546685C|2012-09-11|Aircraft fluid cooling system and an aircraft provided with said system
    FR3039133A1|2017-01-27|AIRCRAFT WITH A PROPULSIVE ASSEMBLY COMPRISING A BLOWER AT THE REAR OF THE FUSELAGE
    FR3015573A1|2015-06-26|AIRCRAFT TURBOMACHINE COMPRISING A HEAT EXCHANGER OF THE PRE-COOLING TYPE
    EP2979980B1|2016-08-24|Method and device for detecting when an aircraft flies in icing conditions
    CA2875044A1|2015-06-19|Casing for a propulsion assembly
    FR2960912A1|2011-12-09|Controller protecting device for turbomachine of aircraft, has lid covering main case and comprising air inlet zone and air outlet zone, and air circulation system assuring circulation of air between air inlet and air outlet zones
    WO2019077251A1|2019-04-25|Device and method for cooling a low pressure turbine in a turbomachine
    CA2818705C|2018-06-12|Oil discharge device and turbomachine comprising such a device
    EP3084187B1|2020-02-26|Fluid-draining device for an aircraft engine
    EP2941382B1|2018-12-26|Electric de-icing device for turbojet engine nacelle
    CA2929951C|2022-03-08|Aircraft propulsion assembly with fire extinguishing system
    EP2697493B1|2018-01-03|Fuel pipes with means for controlling a pump
    EP3039341B1|2018-12-05|Turbomachine combustion chamber comprising air deflection means for reducing wake created by igniter
    CA2944336A1|2015-10-08|Drive chain for a helicopter incorporating a pyrotechnic assistance drive module and helicopter comprising same
    FR3075875B1|2019-11-15|OIL CIRCUIT FOR AUXILIARY OIL RING TURBOMACHINE
    WO2015118274A1|2015-08-13|System for the emergency starting of a turbomachine
    FR3081514A1|2019-11-29|PROPELLANT AIRCRAFT ASSEMBLY AND METHOD FOR REDUCING VENTILATION AIR FLOW IN THE AIRCRAFT PROPELLANT ASSEMBLY
    FR3001525A1|2014-08-01|Method for managing fuel consumption of twin-engine assembly of helicopter, involves operating engine by supplying fuel throughout start-up injector and by interrupting power supply of main injector in super-idle mode
    EP3041742B1|2019-01-09|Engine casing having pivotable access doors
    FR3070418A1|2019-03-01|AIRCRAFT TURBOMACHINE
    FR3044865A1|2017-06-09|ELECTRONIC CALCULATOR OF AN AIRCRAFT TURBOMACHINE
    FR3005486A1|2014-11-14|LUBRICATION IN A TURBOMACHINE
    FR3101613A1|2021-04-09|PROTECTION DEVICE FOR AN ELEMENT OF AN AIR CONDITIONING SYSTEM AGAINST CONTACT WITH AN ADVERSE OBJECT
    同族专利:
    公开号 | 公开日
    RU2016121154A3|2018-09-27|
    RU2016121154A|2018-01-30|
    US10336464B2|2019-07-02|
    EP3087265B1|2018-02-07|
    US20160368618A1|2016-12-22|
    CA2929951A1|2015-07-02|
    FR3015568B1|2016-01-01|
    RU2672197C2|2018-11-12|
    WO2015097391A1|2015-07-02|
    EP3087265A1|2016-11-02|
    CN105793542A|2016-07-20|
    CN105793542B|2017-11-28|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    GB2057574A|1979-08-31|1981-04-01|Gen Electric|Variable clearance control for a gas turbine engine|
    US6082464A|1997-07-22|2000-07-04|Primex Technologies, Inc.|Dual stage fire extinguisher|
    WO2006076649A2|2005-01-12|2006-07-20|Eclipse Aviation Corporation|Fire suppression systems|
    EP2624353A1|2012-02-03|2013-08-07|Airbus Operations GmbH|Emergency supply system for a transportation means, method for supplying electrical power and for inhibiting fire and aircraft with an emergency supply system|FR3041936A1|2015-10-02|2017-04-07|Airbus Operations Sas|PROPULSIVE AIRCRAFT SYSTEM COMPRISING AT LEAST ONE FIRE RESISTANT|
    EP3933183A1|2020-07-02|2022-01-05|Airbus Operations |Propulsion assembly of an aircraft comprising a fire protection system with a line for distribution of fire extinguishing agent|US4351394A|1979-12-28|1982-09-28|Enk William A|Method and system for aircraft fire protection|
    US4441314A|1980-09-26|1984-04-10|United Technologies Corporation|Combined turbine power plant blade tip clearance and nacelle ventilation system|
    JPS61132728A|1984-12-03|1986-06-20|Hitachi Ltd|Internal ventilation equipment of gas turbine package|
    US5239817A|1991-11-04|1993-08-31|General Electric Company|Fire zone ventilation shut-off system|
    US6935433B2|2002-07-31|2005-08-30|The Boeing Company|Helium gas total flood fire suppression system|
    RU2324064C1|2006-08-14|2008-05-10|Николай Борисович Болотин|Energy gas-turbine power plant|
    FR2905358B1|2006-09-06|2008-10-17|Eurocopter France|FIREWALL FLAME OF AIRCRAFT.|
    FR2960912B1|2010-06-08|2012-06-08|Snecma|DEVICE FOR PROTECTING AN AIRCRAFT TURBO-MACHINE CALCULATOR IN CASES OF FIRE|FR3039591B1|2015-07-31|2017-08-25|Snecma|INSULATION OF AN AIRCRAFT TURBOMACHINE TANK IN THE EVENT OF FIRE BY CLOSURE OF A VALVE SENSITIVE TO THE RELEASE OF AN EXTINGUISHING AGENT|
    US10369393B2|2017-11-01|2019-08-06|The Boeing Company|Aircraft fire extinguishing with heated tubing|
    US20200298036A1|2017-11-03|2020-09-24|Bombardier Inc.|Aircraft fire suppression system|
    DE102018222162A1|2018-12-18|2020-06-18|Rolls-Royce Deutschland Ltd & Co Kg|Ventilation and extinguishing device for a gas turbine engine|
    US11230975B2|2019-03-20|2022-01-25|Raytheon Technologies Corporation|Modulated fire extinguishing vent for a gas turbine engine|
    CN111997760B|2019-05-27|2022-01-21|中国航发商用航空发动机有限责任公司|Aircraft engine|
    FR3110141A1|2020-05-13|2021-11-19|Airbus Operations |Aircraft comprising a propulsion unit and a propulsion unit fire-fighting system.|
    法律状态:
    2015-12-15| PLFP| Fee payment|Year of fee payment: 3 |
    2016-05-06| RM| Correction of a material error|Effective date: 20160404 |
    2016-12-05| PLFP| Fee payment|Year of fee payment: 4 |
    2017-11-21| PLFP| Fee payment|Year of fee payment: 5 |
    2018-02-02| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20170719 |
    2018-11-27| PLFP| Fee payment|Year of fee payment: 6 |
    2019-11-20| PLFP| Fee payment|Year of fee payment: 7 |
    2020-11-20| PLFP| Fee payment|Year of fee payment: 8 |
    2021-11-18| PLFP| Fee payment|Year of fee payment: 9 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1363544A|FR3015568B1|2013-12-24|2013-12-24|VENTILATION OF A PROPELLANT AIRCRAFT ASSEMBLY|FR1363544A| FR3015568B1|2013-12-24|2013-12-24|VENTILATION OF A PROPELLANT AIRCRAFT ASSEMBLY|
    PCT/FR2014/053497| WO2015097391A1|2013-12-24|2014-12-22|Aircraft propulsion assembly with fire extinguishing system|
    US15/100,246| US10336464B2|2013-12-24|2014-12-22|Aircraft propulsion assembly with fire extinguishing system|
    CA2929951A| CA2929951A1|2013-12-24|2014-12-22|Aircraft propulsion assembly with fire extinguishing system|
    EP14833513.6A| EP3087265B1|2013-12-24|2014-12-22|Aircraft propulsion assembly with fire extinguishing system and ventilation method|
    CN201480065639.6A| CN105793542B|2013-12-24|2014-12-22|Aircraft propulsion assembly with fire extinguishing system|
    RU2016121154A| RU2672197C2|2013-12-24|2014-12-22|Aircraft propulsion assembly with fire extinguishing system|
    [返回顶部]