• 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:
    Forward lip of turbojet nacelle, comprising an annular defrosting volume (4) receiving a supply of hot air (40) for defrosting its outer walls, this lip further comprising acoustic panels (42) disposed on its wall substantially turned radially towards the axis of the nacelle, characterized in that it comprises hot air outlet bores (50) which are arranged between the front end of the lip, and the acoustic panels (42).
    公开号:FR3023538A1
    申请号:FR1456743
    申请日:2014-07-11
    公开日:2016-01-15
    发明作者:Marc Versaevel；Arnaud Delehouze；Laurence Lemains；Pierre-Francois Behaghel
    申请人:Aircelle SA；
    IPC主号:
    专利说明:

    [0001] The present invention relates to a front lip of a turbojet engine nacelle comprising a de-icing system, as well as a method for supplying hot air to this lip, and a turbojet nacelle equipped with such a front lip.
    [0002] The leading edges of the aircraft, in particular the front lips of the air intakes of the turbojet pods forming forward-facing curved flanges, can under certain climatic conditions, such as the crossing of clouds with a low temperature, present the formation frost that ends up as ice blocks.
    [0003] This results in a modification of the aerodynamic profile of the nacelle, which can disrupt the air supply and the proper functioning of the engine, and the performance of this engine is reduced. In addition, it is possible to obtain a detachment of ice blocks which enter the turbojet engine and damage fan blades. Flight authorizations in icing conditions require the presence of a de-icing system. A known type of deicing or anti-icing system, presented in particular by the documents EP-B1-0913326 or US-A1-20020179773, comprises a circular tube around the nacelle, which supplies hot air taken from the turbojet engine, the interior volume of the front lip of this nacelle to warm its outer walls. Furthermore, to reduce the acoustic emissions of turbojet engines, some interior walls of the nacelle are lined with sandwich panels comprising a central core having cells in the form of a honeycomb, which is covered by a sealed inner back skin, and by a outer front skin facing the sound source, which is perforated or porous. The open cells then constitute a device of the Helmholtz resonator type, which contributes to greatly reducing the acoustic emissions. The central core of the sandwich panel may comprise a single thickness of cells, or two thicknesses separated by a micro-perforated median skin, to improve the acoustic performance of the panel. This type of acoustical panel is particularly arranged on the interior walls of the vein. annular cold air, in the case of a turbofan engine, and on the inner wall of the upstream air inlet. In particular in the prior art documents set forth above, the radially inner surfaces of the leading edge lip, facing the axis of the nacelle, are packed from the front of this lip, of this type of acoustic panels. It should be noted that the lining as far upstream as possible from the air intake of the nacelle, in particular in the substantially cylindrical interior volume, gives the best acoustic performance of this nacelle. However, the acoustic panels constituting a thermal insulation, then small perforations are made on the inner skin of these panels, to obtain a hot air flow through the panels to warm the outer wall from the air flowing to inside this lip. However, this method poses various problems, in particular the hot air flow required is hardly compatible with the acceptable perforation of the inner skin, and the acoustic performance of these panels are reduced. In addition, we obtain a loss of aerodynamic efficiency of the nacelle, with in particular a scoop effect where the outside air enters the first upstream cells through the acoustic panel, to come into the interior volume of the lip, then emerges from the last cells downstream crossing this panel again. The present invention is intended to avoid these disadvantages of the prior art. To this end, it proposes a front lip of a turbojet engine nacelle, comprising an annular defrosting volume receiving a supply of hot air to effect a defrosting of its outer walls, this lip further comprising acoustic panels disposed on its wall substantially turned towards the axis of the nacelle, remarkable in that it comprises hot air outlet holes which are arranged between the front end of the lip, and the acoustic panels.
    [0004] An advantage of this front lip is that one obtains a layer of hot air coming out of the holes which covers the acoustic panels disposed just downstream of these holes, effectively performing a defrosting of the panels. We can then have such acoustic panels as far upstream of the lip, almost to its front end, which gives good acoustic performance, with a minimum loss of aerodynamic qualities of the nacelle. The front lip of the nacelle according to the invention may further comprise one or more of the following features, which may be combined with one another. Advantageously, the bores are arranged just upstream of the acoustic panels. This produces a hot air flow that leaves the annular volume closer to these acoustic panels. Advantageously, the acoustic panels cover a major part of the length along the axis of the nacelle, the outer surface of the annular volume facing this axis. In this way, a large surface coverage is achieved by these acoustic panels, giving a better attenuation of the noises. Advantageously, the surface of the front lip having no acoustic panel, is heated by direct contact with the hot air circulating in the annular volume. In this way, a conventional deicing process is used for this part. Advantageously, the annular volume is separated by internal radial partitions in several angular sectors. These angular sectors 25 can each optimize the air flow of its sector, to obtain a substantially uniform distribution of the hot air layer on the inner periphery of the front lip. In particular, each angular sector may comprise a separate feed tube. The invention also relates to a method of supplying hot air with an annular volume of a front lip of nacelle comprising any one of the preceding characteristics, which adjusts the hot air flow rate delivered in this volume by according to the operating conditions of the turbojet engine. Advantageously, during take-off of the aircraft, the method delivers a lower hot air injection rate.
    [0005] Advantageously, downhill from the aircraft, the method delivers a larger hot air injection rate. The invention furthermore relates to a turbojet engine nacelle comprising a front lip having an annular defrosting volume receiving a supply of hot air, this front lip comprising any of the preceding characteristics. The invention will be better understood and other features and advantages will appear more clearly on reading the following description given by way of example, with reference to the accompanying drawings in which: - Figure 1 is an axial sectional diagram a turbojet engine nacelle 15 comprising acoustic panels arranged in the annular vein; FIGS. 2a and 2b are respectively general and detailed views of a front lip comprising a first de-icing system, produced according to the prior art; FIGS. 3a and 3b are respectively general and detailed views of a front lip comprising a second de-icing system, produced according to the prior art; FIG. 4 is a diagram in axial section of a front lip comprising a de-icing system, produced according to the invention; and FIGS. 5 and 6 are front diagrams of two types of tubes for feeding this front lip. FIG. 1 shows the nacelle having, on the front side indicated by the arrow "AV", upstream of the substantially cylindrical front shell 14, an annular volume 4 located inside the front lip 2, constituting a deicing compartment which is closed by a rear wall 6. The nacelle comprises a fan which sends a secondary stream of cold air into an annular vein 10, disposed downstream.
    [0006] The inner surface 8 of the front ferrule 14 as well as the external and internal structures 12 of the cold air annular vein 10, receive acoustic panels comprising a honeycomb core having a closed rear wall and a front wall disposed in surface that is porous or pierced, to absorb the noise emitted by the engine. Figures 2a and 2b show the annular volume 4 comprising a circular tube 20 disposed substantially in the center of this volume, which is supported by a plurality of radial plates 22 fixed to the rear wall 6 to maintain it in this position.
    [0007] The circular tube 20 has at one end a hot air inlet 24, and at its opposite end an outlet 26 of the air which has cooled after heating the annular internal volume 4 by contact with this tube serving as radiator. The front lip 2 comprising its wall heated from the inside, thus melts frost or ice that could be deposited on its outer surface. Figures 3a and 3b show alternatively another type of hot air supply of the inner annular volume 4, comprising an air inlet bore 30 through the rear wall 6, and a tangentially disposed tube 32 having an orifice output 34, which rotates the air in this annular volume. This gives a good distribution of the hot air throughout the annular volume 4. FIG. 4 shows the internal annular volume 4 comprising a system for distributing the hot air throughout the volume 40, which can be following a known type presented above. The inner surface of the shell 14 turned towards the center of the nacelle, comprises acoustic panels 42 which extend upstream on the front lip 2, covering a major part of the length along the axis of the nacelle, the outer surface of the annular volume 4 facing this axis.
    [0008] Each acoustic panel 42 comprises a honeycomb core 44 having a sealed inner skin 46, and a porous or pierced outer skin 48 which is disposed in the extension of the front part of the lip 2 in order to obtain aerodynamic continuity. . The front lip 2 comprises just upstream of the acoustic panels 42 a succession of bores 50 distributed around the annular periphery, which allow a flow of hot air forming a substantially regular film drawn downstream, and coming to cover by a boundary layer. air 52 these acoustic panels to warm them to prevent the formation of frost, or perform a defrost. In particular, the hot air film causes a deflection of the upstream air droplets, which removes them from the acoustic panel 42, as well as an evaporation of the droplets having passed through this film, which are deposited on this panel. The pattern of the bores 50 and the shape of these bores, in particular the diameter, the distribution, the taper or the inclination of these bores, are adjusted so as to optimize the thickness of the hot air boundary layer 52, and to promote the deviation of the trajectory of the drops relative to the wall of the nacelle. It will be noted that the cells 44 of the acoustic panels 42 having only outward bores, can not constitute a scoop 20 recirculating the flow of hot air coming from the outside, towards the inside of the annular volume 4. The the upstream part of the front lip 2 and the part turned radially outwards, having no acoustic panels 42, are heated in the usual way by the circulation of hot air in the annular volume 4.
    [0009] A compromise is thus obtained allowing as much as possible upstream placement of acoustic panels 42 on the lip of nacelle 2, ensuring good acoustic performance, with an efficient de-icing system consuming a limited flow of hot air, and with aerodynamic losses that remain low.
    [0010] In order to obtain a boundary layer of hot air 52 having a suitable flow rate, enabling it to remain stuck permanently on the acoustic panel 42, it is advantageous to adjust the hot air flow rate as a function of the operating conditions of the turbojet. Indeed if the boundary layer 52 is detached from the acoustic panel 42, there is no effective defrost this panel. In particular at take-off from the aircraft, a high pressure of hot air supplied by the turbojet compressor is obtained, and the depression at the lip 2 is important, then a low hot air injection rate will be achieved. . On descent a lower hot air pressure supplied by the compressor is obtained, and the depression on lip 2 is also low, then a large hot air injection rate will be achieved. FIG. 5 shows a single annular perforated tube 40, distributing, in a regular manner by different holes, the hot air in a succession of angular sectors 62 delimited by radial partitions 60 separating the annular volume 4, which makes it possible to regulate the flow rate in each of these 15 sectors independently. Figure 6 shows the annular volume 4 separated into sectors 62 by radial partitions 60, which are not traversed by pierced tubes 40 having several parts each located in one of the sectors. For these two last versions comprising the annular volume 4 separated in sectors 62, it is possible in particular to adjust the number or the diameter of the bores of the tube 40 opening into each of the sectors in order to obtain a homogeneous distribution of the air boundary layer. hot 52.
    权利要求:
    Claims (10)
    [0001]
    CLAIMS1 - Front lip turbojet engine nacelle, comprising an annular defrosting volume (4) receiving a supply of hot air (40) to perform a defrosting of its outer walls, this lip further comprising acoustic panels (42) arranged on its wall substantially rotated radially towards the axis of the nacelle, characterized in that it comprises hot air outlet bores (50) which are arranged between the front end of the lip, and the acoustic panels (42).
    [0002]
    2 - lip before nacelle according to claim 1, characterized in that the holes (50) are arranged just upstream of the acoustic panels (42).
    [0003]
    3 - front lip nacelle according to claim 1 or 2, characterized in that the acoustic panels (42) cover a major part of the length along the axis of the nacelle, the outer surface of the annular volume (4) rotated towards this axis.
    [0004]
    4 - front lip nacelle according to any one of the preceding claims, characterized in that the surface of the front lip (2) having no acoustic panel (42), is heated by direct contact with hot air 20 flowing in the annular volume (4).
    [0005]
    5 - front lip nacelle according to any one of the preceding claims, characterized in that the annular volume (4) is separated by internal radial partitions (60) in several angular sectors (62).
    [0006]
    6 - front lip of nacelle according to claim 5, characterized in that each angular sector (62) comprises a separate feed tube (40).
    [0007]
    7 - Hot air supply method of an annular volume (4) of a nacelle front lip (2) according to any one of the preceding claims, characterized in that it adjusts the flow of hot air delivered in this volume depending on the operating conditions of the turbojet engine.
    [0008]
    8 - Hot air supply method according to claim 7, characterized in that the takeoff of the aircraft it delivers a lower hot air injection rate.
    [0009]
    9 - Hot air supply method according to claim 7 or 8, characterized in that downhill from the aircraft it delivers a larger hot air injection rate.
    [0010]
    10 - turbojet engine nacelle comprising a front lip having an annular defrosting volume (4) receiving a supply of hot air (40), characterized in that said front lip is made according to any one of claims 1 to 6.
    类似技术:
    公开号 | 公开日 | 专利标题
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    WO2021105022A1|2021-06-03|Air inlet and method for de-icing an air inlet into a nacelle of an aircraft turbojet engine
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    同族专利:
    公开号 | 公开日
    FR3023538B1|2016-07-15|
    US20170122204A1|2017-05-04|
    US10487738B2|2019-11-26|
    WO2016005711A1|2016-01-14|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US4738416A|1986-09-26|1988-04-19|Quiet Nacelle Corporation|Nacelle anti-icing system|
    FR2637251A1|1989-09-27|1990-04-06|Gen Electric|Anti-icing device for the intake cowl of an aircraft engine|
    US20020179773A1|1999-11-23|2002-12-05|Breer Marlin D.|Method and apparatus for aircraft inlet ice protection|
    FR2840879A1|2002-05-22|2003-12-19|Short Brothers Plc|Device for protection against ice accumulation on aircraft nacelle air intake comprises leading pressurized hot gases into structure internal compartment and passing them outside to detach water film from external surface|
    FR2917067A1|2007-06-08|2008-12-12|Airbus France Sas|COATING FOR ACOUSTIC TREATMENT INCORPORATING THE FUNCTION OF TREATING FROST WITH HOT AIR|
    FR2952032A1|2009-11-05|2011-05-06|Airbus Operations Sas|Aircraft nacelle, has heat-conducting material element provided between lip and duct to ensure continuity of aerodynamic surfaces and heat propagation, and comprising heat resistant material coating for acoustic treatment|
    FR2980776A1|2011-10-03|2013-04-05|Airbus Operations Sas|AIRCRAFT NACELLE COMPRISING A PANEL FOR ACOUSTIC TREATMENT INTEGRATING HOT AIR CHANNELS AND AT LEAST ONE ANNULAR CHANNEL|FR3089252A1|2018-12-04|2020-06-05|Safran Nacelles|Device and method for defrosting an air intake of a nacelle of an aircraft turbojet engine|AU581684B2|1984-10-08|1989-03-02|Short Brothers Plc|Duct for hot air|
    US5088277A|1988-10-03|1992-02-18|General Electric Company|Aircraft engine inlet cowl anti-icing system|
    US5841079A|1997-11-03|1998-11-24|Northrop Grumman Corporation|Combined acoustic and anti-ice engine inlet liner|
    FR2820716B1|2001-02-15|2003-05-30|Eads Airbus Sa|PROCESS FOR DEFROSTING BY FORCED CIRCULATION OF A FLUID, OF A REACTION ENGINE AIR INLET COVER AND DEVICE FOR ITS IMPLEMENTATION|
    FR2924409B1|2007-12-03|2010-05-14|Airbus France|AIRCRAFT NACELLE COMPRISING MEANS FOR HOT AIR EXHAUST|
    FR2924408B1|2007-12-03|2010-05-07|Airbus France|TURBOREACTOR NACELLE AND METHOD FOR CONTROLLING DECOLUTION IN A TURBOREACTEUR NACELLE|
    WO2010086560A2|2009-02-02|2010-08-05|Airbus Operations Sas|Aircraft nacelle including an optimised acoustic processing system|US10533497B2|2016-04-18|2020-01-14|United Technologies Corporation|Short inlet with integrated liner anti-icing|
    US10221765B2|2016-08-26|2019-03-05|Honeywell International Inc.|Anti-icing exhaust system|
    FR3062880A1|2017-02-10|2018-08-17|Airbus|AIR INTAKE STRUCTURE FOR AN AIRCRAFT NACELLE|
    FR3077800B1|2018-02-12|2020-09-25|Safran Nacelles|DEFROSTING AND ACOUSTIC TREATMENT DEVICE FOR AN AIR INLET LIP OF A TURBOREACTOR NACELLE|
    FR3095420A1|2019-04-26|2020-10-30|Safran Nacelles|Nacelle air inlet and nacelle having such an air inlet|
    FR3096662A1|2019-05-27|2020-12-04|Safran Nacelles|Turbomachine nacelle air inlet lip comprising an acoustic device and method of manufacturing such a lip|
    法律状态:
    2015-06-30| PLFP| Fee payment|Year of fee payment: 2 |
    2016-01-15| PLSC| Search report ready|Effective date: 20160115 |
    2016-06-30| PLFP| Fee payment|Year of fee payment: 3 |
    2017-06-29| PLFP| Fee payment|Year of fee payment: 4 |
    2018-03-02| CD| Change of name or company name|Owner name: SAFRAN NACELLES, FR Effective date: 20180125 |
    2018-06-28| PLFP| Fee payment|Year of fee payment: 5 |
    2020-06-23| PLFP| Fee payment|Year of fee payment: 7 |
    2021-06-23| PLFP| Fee payment|Year of fee payment: 8 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1456743A|FR3023538B1|2014-07-11|2014-07-11|TURBOJET NACELLE FRONT LEVER HAVING HOT AIR DRILLING UPSTREAM ACOUSTIC PANELS|FR1456743A| FR3023538B1|2014-07-11|2014-07-11|TURBOJET NACELLE FRONT LEVER HAVING HOT AIR DRILLING UPSTREAM ACOUSTIC PANELS|
    PCT/FR2015/051925| WO2016005711A1|2014-07-11|2015-07-10|Front lip of a turbofan engine nacelle comprising hot-air bores upstream from acoustic panels|
    US15/403,658| US10487738B2|2014-07-11|2017-01-11|Front lip of a turbofan engine nacelle comprising hot-air bores upstream from acoustic panels|
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