• 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:
    The invention relates to an aircraft nacelle comprising an air inlet (46) configured to channel an air flow (48) towards a motor which comprises a fan with an axis of rotation (44), said air inlet (46) comprising, in a flow direction of the air flow (48), a lip (50) with a leading edge and an inner conduit (52) extending said lip (50) to the blower, characterized in that the air inlet (46) comprises a plurality of bumps (70) distributed on the circumference of the air inlet (46) and after each bump (70) a concave shape ( 74) configured to obtain a compression of the airflow.
    公开号:FR3022218A1
    申请号:FR1455341
    申请日:2014-06-12
    公开日:2015-12-18
    发明作者:Francois Rouyre
    申请人:Airbus Operations SAS；
    IPC主号:
    专利说明:

    [0001] BACKGROUND OF THE INVENTION The present invention relates to an aircraft nacelle comprising an improved air inlet. As represented in FIG. 1, an aircraft nacelle 10 comprises, at the front, an air inlet 12 making it possible to channel an air flow towards a motor, a first part of the incoming air flow, called primary flow, passing through a turbojet to participate in the combustion, the second part of the air flow, called secondary flow, being driven by a blower 14 and flowing in an annular duct delimited by the inner wall of the nacelle and the wall exterior of the turbojet. For the following description, the longitudinal direction corresponds to the direction of the axis of rotation 16 of the fan 14. A longitudinal plane is a plane which passes through the axis of rotation 16. A transverse plane is a plane perpendicular to the axis of rotation 16. As illustrated in Figure 2, the air inlet 10 comprises a lip 18 whose surface in contact with the aerodynamic flows is extended inside the nacelle by a substantially circular inner conduit 20 and outside the nacelle by an outer wall 22 substantially circular. The lip 18 makes it possible to channel an air flow 24 inside the inner conduit 20. In a longitudinal plane, the air inlet 12 has a section with a substantially rectilinear portion at the inner conduit 20, a portion substantially rectilinear at the outer wall 22 and a curved intermediate portion with a small radius of curvature at the lip 18. This curved intermediate portion has a vertex A which corresponds to the most forward point of the lip 18. This vertex Has described an approximately circular profile 26 along the circumference of the nacelle 10, as illustrated in FIG. 1. This profile 26 corresponds to the leading edge of the air inlet.
    [0002] According to the prior art, this profile 26 is inscribed in a plane 28. Preferably, this plane 28 is not perpendicular to the axis of rotation 16 of the fan 14 and forms an angle of the order of 75 to 105 ° with this axis of rotation 16 so that the top of the air inlet located at the top of the nacelle which corresponds to an angular position at 00 is more forward than the top of the air inlet located at the bottom the nacelle which corresponds to a 180 ° angular position. FIG. 3A shows the expanded profile 26 described by the vertex A with respect to a reference plane perpendicular to the axis of rotation 16 of the fan 14.
    [0003] Since the plane 28 is inclined with respect to the axis of rotation 16, the profile 26 describes approximately a single sinusoid with an amplitude E which corresponds to the difference in the longitudinal direction between the top of the air inlet located at the top the basket and the one at the bottom. In Figures 3B and 3C, there is shown respectively a cross section of the inner conduit 20 at a point B of the air inlet and at a point C of the air inlet. The points B and C each approximately describe a circle respectively of radius R and radius R ', as illustrated in FIGS. 3B and 3C. When the air flow enters the air inlet 12, depending on the flight phase of the aircraft (climb, cruise, descent), airflow can be detached at the level of the zone junction between the lip 18 and the inner conduit 20, causing an increase in the thickness of the boundary layer 32 at the surface of the inner conduit 20, as shown in Figure 2. The thickness of the boundary layer 32 being greater than the game provided between the inner conduit 20 and the ends of the blades 34, the latter are not in a longitudinal laminar flow but interfere with the boundary layer 32. The interaction of the ends of the blades 34 and the turbulent flows of the boundary layer 32 generates a source of noise. To limit the impact of noise pollution, techniques have been developed to reduce noise, in particular by arranging, at the walls of the inner conduit 20, panels or coatings intended to absorb some of the sound energy, in particular by using the principle of Helmholtz resonators.
    [0004] Even if these panels or coatings are efficient, they only attenuate the noise emitted by some sources of motorization and are not intended to limit their appearances. The present application aims to improve the aerodynamic characteristics of the air inlet to limit the appearance of noises. For this purpose, the subject of the invention is an aircraft nacelle comprising an air inlet configured to channel an air flow towards an engine which comprises a fan with an axis of rotation, said air intake comprising, in a direction of flow of the air flow, a lip with a leading edge and an inner conduit which extends said lip towards the blower, characterized in that the air inlet comprises a plurality of distributed bumps on the circumference of the air inlet and after each bump a concave shape configured to obtain a compression of the air flow. This compression causes an acceleration of the air flow in contact with the inner duct and therefore a reduction of the thickness of the boundary layer at the surface of the inner duct. In this way, the ends of the blades no longer interfere with the turbulent flow of the boundary layer which limits the appearance of noise. Preferably, the air inlet comprises at least thirty-two bumps regularly distributed on its circumference.
    [0005] According to other features: the bumps are distributed in a plane parallel to the leading edge; the bumps are arranged at a junction zone between the lip and the inner duct; the bumps form in a plane parallel to the leading edge a succession of hollow shapes and projecting shapes with a sinusoidal profile; each bump has a height of between 1 and 5 cm; each bump comprises a vertex and an upstream foot separated by a distance in a direction parallel to the axis of rotation less than or equal to twice the height. Advantageously, the leading edge has a profile which alternately comprises recessed shapes and projecting shapes, a protruding shape being separated from the fan a greater distance than the adjacent recessed shapes. Preferably, the profile of the leading edge describes a succession of recessed shapes and projecting shapes distributed on either side of a plane which forms an angle of 75 to 1050 relative to the axis of rotation of the plane. the blower.
    [0006] According to one embodiment, the projecting shapes have a radius of curvature less than the radius of curvature of the recessed shapes. Other features and advantages will become apparent from the following description of the invention, a description given by way of example only, with reference to the appended drawings, in which: FIG. 1 is a perspective view of a nacelle; FIG. 2 is a longitudinal section of a portion of an air inlet which illustrates the prior art. FIG. 3A is a curve of an extended profile of the leading edge of FIG. 3B and 3C are cross-sections at a point B and at a point C of an air inlet of the prior art. FIG. FIG. 5 is a longitudinal section of a part of an air inlet of the nacelle of FIG. 4, FIG. 6 is a perspective view illustrating a view of an aircraft nacelle which illustrates the invention. FIG. air flows at the surface of a sector of the air inlet of the nacelle of FIG. 4, FIG. 7A is a curve which shows FIG. 7B is a cross section at a point B of the air inlet of FIG. 5, FIG. 7C is a cross-section at a point C of the air inlet of the figures, FIG. 8 is a curve which illustrates an expanded profile described by the leading edge of an air inlet according to another variant of the FIG. 9 is a curve which illustrates an expanded profile described by the leading edge of an air inlet according to another variant of the invention.
    [0007] An aircraft nacelle 40 comprises at the front an air inlet 46 configured to channel an air flow 48 towards a motor that integrates a blower 42 with an axis of rotation 44. For the rest, the axis of rotation 44 corresponds to the longitudinal direction. As illustrated in Figure 5, the air inlet 46 comprises a lip 50, a substantially circular inner conduit 52 which extends the lip 50 inside the nacelle. The lip 50 is extended outside the nacelle by a substantially circular outer wall 54. In a longitudinal plane visible in FIG. 5, the outer wall and the air inlet 46 have a longitudinal section which comprises a first rectilinear portion 56 which corresponds to the outer wall 54, a second rectilinear portion 58 which corresponds to a portion inner duct 52 and a curved intermediate portion 60 which corresponds to lip 50.
    [0008] The lip 50 comprises a vertex A which corresponds to the foremost point of the nacelle for each longitudinal plane. The apex A describes along the circumference of the nacelle a closed curve which forms a leading edge 62. This leading edge 62 makes it possible to split the flow of air in two, a first flow which enters the inlet air 46 and another flow of air that passes outside the nacelle. The inner conduit 52 comprises a cross section (perpendicular to the longitudinal direction) to the right of the circular blower 42. Thus, as illustrated in FIG. 7C, a point C of the inner duct 52 describes a circular profile 64 with a radius Rc. The blower 42 comprises a plurality of blades 66 each with an end 68 separated from a clearance J of the inner duct 52. According to one characteristic of the invention, the air inlet 46 comprises a plurality of circumferentially distributed bumps 70. of the air inlet 46. Advantageously, the bumps 70 are distributed in a plane parallel to the leading edge 62. This succession of bumps 70 forms a profile 72, shown in FIG. 7B, comprising an alternation of hollow shapes. and protruding forms. Each hollow shape has a low point Pb. Preferably, all the low points Pb are arranged on the same circle which is centered on the axis of rotation 44 and which has a radius equal to the radius Rc of the inner duct 52. protruding which corresponds to a hump 70 has a vertex S which corresponds to the highest point. Preferably, all the vertices are arranged on the same circle which is centered on the axis of rotation 44 and which has a radius Rs less than the radius Rc. Advantageously, each bump 70 has a height of the order of 1 to 5 cm.
    [0009] Thus, the distance between the rc and rs rays is of the order of 1 to 5 cm. According to one characteristic of the invention, the air inlet 46 comprises at least thirty-two (32) bumps 70 distributed regularly over the circumference of the air inlet 46, to obtain the desired effect. According to one embodiment, the profile 72 formed by the alternation of recessed shapes and projecting shapes is a sinusoid carried by a circle of radius equal to (Rc + Rs) / 2. However, the invention is not limited to this geometry for the bosses 70. According to another characteristic of the invention, for each boss 70, the vertex S is disposed on a circle of radius Rs less than that of a point P of the air intake 46 located just after the boss 70. Thus, as shown in FIG. 5, in a longitudinal plane, each boss has, in the direction of flow of the air flow 48, a profile such as the diameter of the air inlet 46 decreases continuously from a downstream leg of the hump to its vertex S and then continuously increases from the vertex S to an upstream foot P of the hump. In this way, the air inlet 46 comprises after each boss 70 a concave shape 74 at the upstream leg P configured to obtain a compression 76 (visible in Figure 6) of the air flow. This compression causes an acceleration of the air flow in contact with the inner duct 52 and therefore a reduction of the thickness of the boundary layer 78 at the surface of the inner duct 52. As shown in FIG. thickness less than the clearance J between the ends 68 of the blades 66 and the inner conduit 52. In this way, the ends 68 of the blades 66 no longer interfere with the turbulent flow of the boundary layer 78 which limits the occurrence of noise.
    [0010] Preferably, the distance between the vertex S and the upstream leg P in the longitudinal direction is less than or equal to twice the distance between the vertex S and the upstream leg P in the transverse direction which corresponds to the height of the hump. Advantageously, the bumps 70 are positioned at the theoretical location of the detachment of the air flow in the absence of bumps. According to one embodiment, the bumps 70 are positioned at the junction zone between the lip 50 and the inner pipe 52. According to one embodiment, each bump 70 has an identical profile in all the planes containing a line passing through the top S and perpendicular to the axis of rotation 44. The presence of the bumps 70 according to the invention at the air inlet 46 provides the following advantages: The bumps 70 limit the separation of the air flow entering in the air inlet 46 which tends to reduce the thickness of the boundary layer 78 at the surface of the inner pipe 52. In this way, the ends 68 of the blades 66 of the fan no longer interfere with the turbulent flows of the boundary layer 78 which on the one hand reduces the appearance of noise, and on the other hand, improves the performance of the fan. According to another advantage, the succession of hollow and projecting shapes on the circumference of the air inlet 46 makes it possible to maintain air flows oriented in the longitudinal direction and prevents them from swirling, which also contributes to improve the performance of the blower.
    [0011] According to another characteristic of the invention illustrated in FIGS. 5, 7A, 8 and 9, the leading edge 62 has a profile 80, 80 ', 80 "which alternately comprises recessed shapes 82 and shaped shapes. projection 84, a protruding shape being separated from the fan from a larger distance than the adjacent recessed shapes This configuration improves the aerodynamic characteristics of the air inlet According to a first embodiment shown in FIG. , the profile 80 of the leading edge 62 describes a sinusoid centered with respect to a plane 86 perpendicular to the axis of rotation 44 of the fan According to another embodiment shown in FIGS. 5 and 9, the profile 80 ' Leading edge 62 describes a succession of hollow shapes 82 and projecting shapes 84 distributed on either side of a plane 88 which forms an angle of 75 to 105 ° with respect to the axis of rotation 44. according to another embodiment of the invention. shown in FIG. 8, the recessed shapes 82 and the projecting shapes 84 are not symmetrical. According to this embodiment, all the recessed shapes 82 are identical, all the projecting shapes 84 are identical and the projecting shapes 84 have a radius of curvature less than the radius of curvature of the recessed shapes 82.
    权利要求:
    Claims (10)
    [0001]
    REVENDICATIONS1. An aircraft nacelle comprising an air inlet (46) configured to channel a flow of air (48) towards a motorization which includes a blower (42) with an axis of rotation (44), said inlet of air (46) comprising, in a flow direction of the airflow (48), a lip (50) with a leading edge (62) and an inner conduit (52) which extends said lip (50) towards the blower (42), characterized in that the air inlet (46) comprises a plurality of bumps (70) distributed around the circumference of the air inlet (46) and after each bump (70) a shape concave (74) configured to obtain a compression (76) of the airflow.
    [0002]
    2. aircraft nacelle according to claim 1, characterized in that the air inlet (46) comprises at least thirty-two bumps (70) regularly distributed on its circumference.
    [0003]
    3. aircraft nacelle according to claim 1 or 2, characterized in that the bumps (70) are distributed in a plane parallel to the leading edge (62).
    [0004]
    4. Aircraft nacelle according to one of the preceding claims, characterized in that the bumps (70) are disposed at a junction zone between the lip (50) and the inner conduit (52).
    [0005]
    5. Aircraft nacelle according to one of the preceding claims, characterized in that the bumps (70) form in a plane parallel to the leading edge (62) a succession of recessed shapes and protruding shapes with a profile (72) sinusoidal.
    [0006]
    6. Aircraft nacelle according to one of the preceding claims, characterized in that each bump has a height of between 1 and 5 cm.
    [0007]
    7. Aircraft nacelle according to claim 6, characterized in that each boss (70) comprises a vertex (S) and an upstream foot (P) separated by a distance in a direction parallel to the axis of rotation (44). ) less than or equal to twice the height.
    [0008]
    8. Aircraft nacelle according to one of the preceding claims characterized in that the leading edge (62) has a profile (80, 80 ', 80 ") which alternately comprises recessed shapes (82) and protruding shapes (84), a protruding shape being separated from the fan at a greater distance than the adjacent recessed shapes.
    [0009]
    9. Aircraft nacelle according to claim 8, characterized in that the profile (80 ') of the leading edge (62) describes a succession of recessed shapes (82) and projecting shapes (84) distributed from and other of a plane (88) which forms an angle of 75 to 105 ° with respect to the axis of rotation (44) of the fan.
    [0010]
    10. An aircraft nacelle according to claim 8 or 9, characterized in that the projecting shapes (84) have a radius of curvature less than the radius of curvature of the recessed shapes (82).
    类似技术:
    公开号 | 公开日 | 专利标题
    EP2955112B1|2017-01-04|Aircraft nacelle including an air intake
    CA2722077C|2016-07-26|Centripetal air bleed from a turbomachine compressor rotor
    EP1873388B1|2010-04-14|Turbomachine cowl having noise suppression triangular tabs with double crests
    CA2694996C|2015-10-06|Noise control chevron for a nozzle, and nozzle and turboshaft engine provided with such a chevron
    CA2873942C|2020-05-12|Fan blade for a turbojet of an aircraft having a cambered profile in the foot sections
    EP3477050B1|2022-03-02|Turbomachine compressor with asymmetricly contoured wall ring
    FR2565631A1|1985-12-13|EXHAUST MIXER FOR A GAS TURBINE AIRCRAFT ENGINE
    WO2008029062A2|2008-03-13|Device for improving the effectiveness of acoustic treatments in an aircraft drive duct
    FR2947868A1|2011-01-14|Nacelle for airplane, has thrust reverser device including flaps adapted to occupy folded and deployed states, and ring arranged remotely from inner wall of nacelle and outer surface supporting flaps
    CA2640236A1|2009-03-28|Cowling for turbine engine nozzle with jet noise reduction patterns
    EP3477049A1|2019-05-01|Indented ring for a turbine engine compressor
    WO2009138597A1|2009-11-19|Dual-flow turbine engine for aircraft with low noise emission
    EP2029879B1|2016-08-10|Aircraft propulsion unit and corresponding aircraft
    EP3295009B1|2019-11-20|Lobed mixer with scoops
    CA2980794A1|2016-09-29|Device with gratings for ejecting microjets in order to reduce the jet noise of a turbine engine
    WO2015036684A1|2015-03-19|Afterbody for a mixed-flow turbojet engine comprising a lobed mixer and chevrons with a non-axisymmetric inner surface
    EP2839168A1|2015-02-25|Improved casing for turbomachine blisk and turbomachine equipped with said casing
    WO2017055757A1|2017-04-06|Aerodynamic system comprising a vortex generator supplied by exhaust gases
    EP3693597A1|2020-08-12|Acoustic resonator
    FR2978731A1|2013-02-08|Aircraft propulsion assembly, has reinforcement ring connected at air inlet and/or engine with opening and outer peripheral edge concentric with opening, on level of junction zone between interior air inlet pipe and engine pipe
    FR3093756A1|2020-09-18|secondary flow rectifier has integrated nozzle
    FR3078106A1|2019-08-23|TURBOMACHINE NACELLE COMPRISING AN ACOUSTIC EVACUATION DRIVE
    FR3034142A1|2016-09-30|MICROJET EJECTION GRID DEVICE FOR REDUCING JET NOISE FROM A TURBOMACHINE
    FR3016390A1|2015-07-17|DAWN SUPPRESSING ACOUSTIC REFLECTIONS, AND TURBOMACHINE
    FR3022597A1|2015-12-25|TRIPLE FLUX DIFFUSER FOR TURBOMACHINE MODULE COMPRISING AIR PIPING DEVICES BETWEEN THE TWO WALLS OF THE DIFFUSER SEPARATION
    同族专利:
    公开号 | 公开日
    US9623979B2|2017-04-18|
    EP2955112B1|2017-01-04|
    FR3022218B1|2016-07-15|
    US20150360790A1|2015-12-17|
    EP2955112A1|2015-12-16|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    GB1171508A|1966-10-28|1969-11-19|Snecma|Improved Silencer Devices for Gas Turbine Jet Engines|
    US3664612A|1969-12-22|1972-05-23|Boeing Co|Aircraft engine variable highlight inlet|
    WO1991001034A2|1989-06-26|1991-01-24|Grumman Aerospace Corporation|An acoustic liner|
    EP1607603A2|2004-06-10|2005-12-21|United Technologies Corporation|Gas turbine engine inlet with noise reduction features|
    US20110006165A1|2009-07-10|2011-01-13|Peter Ireland|Application of conformal sub boundary layer vortex generators to a foil or aero/ hydrodynamic surface|
    EP2607657A2|2011-12-20|2013-06-26|Rolls-Royce plc|Active flow control intake for a gas turbine engine|GB2534663A|2014-11-24|2016-08-03|Snecma|Reduction of turbofan noise|
    CN112572810A|2020-11-25|2021-03-30|北京空天技术研究所|Air inlet passage boundary layer separation eliminating device and aircraft|US3611724A|1970-01-07|1971-10-12|Gen Electric|Choked inlet noise suppression device for a turbofan engine|
    CA1324999C|1986-04-30|1993-12-07|Walter M. Presz, Jr.|Bodies with reduced surface drag|
    CN104011376A|2011-10-20|2014-08-27|奥金公司|Aerodynamic Modification Of A Ring Foil For A Fluid Turbine|US10415505B2|2013-08-12|2019-09-17|United Technologies Corporation|Non-axisymmetric fan flow path|
    US10145301B2|2014-09-23|2018-12-04|Pratt & Whitney Canada Corp.|Gas turbine engine inlet|
    US10837362B2|2016-10-12|2020-11-17|General Electric Company|Inlet cowl for a turbine engine|
    US10829198B2|2017-06-21|2020-11-10|The Boeing Company|Krueger flap apparatus and methods incorporating a bullnose having a contour variation along a spanwise direction|
    US20190225318A1|2018-01-25|2019-07-25|General Electric Company|Aircraft systems and methods|
    GB201815759D0|2018-09-27|2018-11-14|Rolls Royce Plc|Nacelle intake|
    CN111942600B|2020-08-06|2022-03-08|四川航天中天动力装备有限责任公司|Boundary layer-free partition air inlet channel|
    法律状态:
    2015-06-19| PLFP| Fee payment|Year of fee payment: 2 |
    2015-12-18| PLSC| Search report ready|Effective date: 20151218 |
    2016-06-27| PLFP| Fee payment|Year of fee payment: 3 |
    2018-03-30| ST| Notification of lapse|Effective date: 20180228 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1455341A|FR3022218B1|2014-06-12|2014-06-12|AIRCRAFT NACELLE COMPRISING AN ENHANCED AIR INTAKE|FR1455341A| FR3022218B1|2014-06-12|2014-06-12|AIRCRAFT NACELLE COMPRISING AN ENHANCED AIR INTAKE|
    US14/728,593| US9623979B2|2014-06-12|2015-06-02|Aircraft nacelle comprising an improved air intake|
    EP15171344.3A| EP2955112B1|2014-06-12|2015-06-10|Aircraft nacelle including an air intake|
    [返回顶部]