Classifications

• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
  • F02K—JET-PROPULSION PLANTS
  • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
  • F02K1/54—Nozzles having means for reversing jet thrust
  • F02K1/64—Reversing fan flow
  • F02K1/70—Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing
  • F02K1/72—Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow
• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
  • B64D29/00—Power-plant nacelles, fairings, or cowlings
  • B64D29/06—Attaching of nacelles, fairings or cowlings
• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
  • B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
  • B64D33/04—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of exhaust outlets or jet pipes
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
  • F02K—JET-PROPULSION PLANTS
  • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
  • F02K1/78—Other construction of jet pipes
  • F02K1/80—Couplings or connections
  • F02K1/805—Sealing devices therefor, e.g. for movable parts of jet pipes or nozzle flaps
• • 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
  • F05D2300/00—Materials; Properties thereof
  • F05D2300/50—Intrinsic material properties or characteristics
  • F05D2300/501—Elasticity
• • 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

• the present invention relates to a turbojet engine nacelle comprising a variable nozzle section.
• a nacelle generally has a tubular structure comprising an air inlet upstream of the turbojet engine, a median section intended to surround a fan of the turbojet engine, a downstream section housing a thrust reverser means and intended to surround the combustion chamber of the turbojet engine. , and is generally terminated by an ejection nozzle whose output is located downstream of the turbojet engine.
• Modern nacelles are intended to house a turbofan engine capable of generating through the blades of the rotating fan a flow of hot air (also called primary flow) from the combustion chamber of the turbojet engine, and a flow of cold air (secondary flow) flowing outside the turbojet engine through a vein formed between a fairing of the turbojet and an inner wall of the nacelle.
• the two air flows are ejected from the turbojet engine from the rear of the nacelle.
• the role of a thrust reverser is, during the landing of an aircraft, to improve the braking capacity thereof by redirecting forward at least a portion of the thrust generated by the turbojet engine.
• the inverter obstructs the cold flow vein and directs the latter towards the front of the nacelle, thereby generating a counter-thrust which is added to the braking of the wheels of the aircraft.
• the means implemented to achieve this reorientation of the cold flow vary according to the type of inverter.
• the grid inverters as illustrated in FIGS. 1 to 3, are known in which the reorientation of the air flow is carried out by deflection grids 1 associated with a sliding cover 2 intended to reveal or cover these grids. 1, the translation of the hood 2 being effected along a longitudinal axis substantially parallel to the axis of the nacelle.
• This sliding cowl 2 is thus able to pass alternately from a closed position illustrated in FIG. 1 in which it ensures the aerodynamic continuity of the nacelle and covers the deflection grids 1 at an open position in which it opens a passage in the nacelle for the deviated flow and discovers the deflection grids 1.
• the sliding cowl belongs to the rear section and has a downstream side forming the ejection nozzle 3 for channeling the ejection of the air flows.
• This nozzle 3 comprises at least one panel 4 rotatably mounted, said panel 4 being adapted to pivot between a normal position illustrated in FIG. 1 in which it ensures the aerodynamic continuity of the nacelle, a thrust reversal position in which it obstructs the cold flow vein 5 and a position shown in Figure 2 causing a variation of the nozzle section 3. It is possible to adjust the degree of pivoting of the movable panel 4 and allow to either vary the nozzle section of ejection 3 is to cause the reversal of the cold air flow in the vein 5 in reverse jet according to the degree of displacement of the movable cover 2.
• the movable cover 2 must be advanced upstream in the direction of the structure fixed upstream 6 of the nacelle.
• a cavity 7 is created at the interface between these two elements.
• this cavity 7 generates a discontinuity of the aerodynamic lines of the surface of the nacelle, resulting in a degradation of the performance of the propulsion unit of the latter and an increase in consumption. of the aircraft.
• An object of the present invention is to overcome the aforementioned drawbacks.
• the invention proposes a nacelle comprising a fixed structure, a nozzle of variable section and a cover slidably mounted on said fixed structure in particular in a race for varying the section deladite nozzle characterized in that it comprises, in in addition, a set of aerodynamic continuity arranged between the fixed structure and the movable cowl, said assembly comprising elastic means able to be compressed between the fixed structure and said movable cowl when the cowl is in the upstream portion of its travel and to relax when the hood gets located in a downstream part of its race so as to ensure the aerodynamic continuity of the lines between said fixed structure and said movable cowl.
• the present invention offers the advantage of eliminating the aerodynamic line defect in a direct jet between the upstream fixed structure and the movable cowl while allowing the free movement between these two elements, in particular so as to move the movable cowl 30 upstream of the the nacelle to vary the section of the nozzle.
• the device may comprise one or more of the following features, taken in isolation or in combination technically possible:
• the aerodynamic continuity assembly comprises a flexible profile of elastomer type
• the aerodynamic continuity assembly comprises a rigid section able to be deformed elastically; -
• the profile is pin spring type, leaf spring or bellows;
• the aerodynamic continuity assembly comprises a non-deformable rigid profile associated with elastic return means
• the profile is formed of a single piece or of several sectors of parts to be assembled together;
• the aerodynamic continuity assembly is capable of ensuring the aerodynamic continuity of the external lines between the upstream fixed structure and the movable hood;
• the aerodynamic continuity assembly is capable of ensuring the aerodynamic continuity of the internal lines between the fixed structure and the movable hood;
• a heel can be attached to the upstream end of the movable cowl
• the aerodynamic continuity element is integral with the fixed structure
• the aerodynamic continuity element is integral with the movable cowl
• the nacelle comprises a downstream section equipped with a thrust reverser device.
• Figures 1 and 2 are schematic longitudinal sectional representations of a thrust reverser device of the prior art having movable panels respectively in normal nozzle position and closed nozzle position;
• Figure 3 is a detail view of Figure 1 on the interface between a movable cover and the upstream fixed structure of the thrust reverser device;
• FIG. 4 is a diagrammatic representation in longitudinal section of a thrust reverser device comprising a set of aerodynamic continuity according to a first embodiment of the present invention;
• Figures 5 to 8 are detail views of Figure 4 on the interface between a movable cowl and the fixed upstream structure of the thrust reverser device respectively before and after moving the movable cowl upstream;
• FIG. 9 is a diagrammatic representation in longitudinal section of a thrust reverser device comprising a set of aerodynamic continuity according to a second embodiment of the present invention
• FIGS. 10 to 13 are detailed views of FIG. 9 on the interface between a movable cowl and the upstream fixed structure of the thrust reverser device respectively before and after movement of the upstream moving cowl;
• Figures 14 and 15 are diagrammatic representations in longitudinal section of a thrust reverser device comprising a set of aerodynamic continuity according to a third embodiment of the present invention respectively before and after displacement of the movable cowl upstream;
• FIGS. 16 and 17 are diagrammatic representations in longitudinal section of a thrust reverser device comprising a set of aerodynamic continuity according to a fourth embodiment of the present invention respectively before and after movement of the upstream moving cowl.
• a nacelle generally has a structure comprising an upstream section forming an air inlet, a central section surrounding the turbojet fan and a downstream section surrounding the turbojet engine.
• this downstream section comprises an external structure 10 comprising a thrust reverser device and an internal motor fairing structure (not shown) defining with the external structure 10 a vein (not shown) intended to to the circulation of a cold flow in the case of the turbojet engine nacelle as presented here.
• the downstream section further comprises an upstream fixed structure 20 comprising a front frame 21, said upstream fixed structure 20 being extended by a movable thrust reverser cover 30 and an ejection nozzle section (not shown).
• the movable thrust reverser cover 30 is intended to be actuated in a substantially longitudinal direction of the nacelle between a closed position in which it covers grids 31 of airflow deflection and an open position in which it is spaced from the frame before 21 opening, then a passage in the nacelle by discovering the airflow deflection grids.
• the ejection nozzle section in the extension of the movable cover 30 comprises a series of movable panels rotatably mounted at a downstream end of the movable cover and distributed over the periphery of the ejection nozzle section.
• Each movable panel is adapted to pivot between several positions, namely a normal position in which it ensures the aerodynamic continuity of the nacelle, a reverse thrust position in which it obstructs the flow of cold flow and return this air to the grids deflection 31 which ensure the reorientation of the flow thus allowing the inverted jet and positions where they allow to vary the section of the nozzle.
• a plurality of recesses 50, 51 may be provided between the front frame 21 and the movable cover 30 on the outer surface. of the nacelle and the inner surface of the front frame 21.
• a set of aerodynamic continuity 40 is arranged at the interface of the front frame 21 and the movable cover 30 and housed in the recess or recesses 50, 51.
• This aerodynamic continuity assembly 40 comprises elastic means capable of being compressed between the front frame 21 and the movable hood 30 when the hood is in the upstream portion of its stroke to vary the nozzle section and able to relax when the hood is in a downstream portion of its stroke so as to ensure the aerodynamic continuity of the lines between the front frame 21 and the movable hood 30.
• this set of aerodynamic continuity 40 fits into the aerodynamic profile of the nacelle taking several positions, namely:
• the presence of the aerodynamic continuity element 40 makes it possible to overcome the aerodynamic line defect in a direct jet while allowing displacement of the movable cover 30 upstream of the nacelle to vary the section of the nozzle.
• the aerodynamic continuity element 40 can be placed at the interface between the front frame 21 and the movable cover 30 on the outer side the nacelle and / or the side of the cold flow vein thus ensuring the continuity of external aerodynamic lines and / or internal aerodynamic lines between these two elements as will be seen later with reference to the various figures.
• an aerodynamic continuity member 40 in the form of a flexible section 41 adapted to ensure the continuity of the outer aerodynamic lines of the nacelle.
• This section 41 thus has a shape and dimensions adapted to fill the space existing between the downstream end of the front frame 21 and the upstream free end of the movable cover 30 when the cover 30 is in a position in which the nozzle section does not vary.
• This profile 41 may have a downstream end having a form of overlap with the upstream end of the movable cover 30, in order to maintain the lowest aerodynamic line break possible even when moving upstream of the movable cover 30 to vary the section. of nozzle.
• the aerodynamic continuity element 40 is sufficiently flexible to deform in contact with the upstream end of the movable cover 30 when the latter moves in a translation movement upstream towards the front frame 21 to rotate the movable panels of the nozzle and then to resume its neutral position and its shape after retreating the hood 30 downstream.
• the profile 41 has a general shape in L.
• an upstream end configured to be fixed to the front frame 21 and having a shape complementary to the downstream end of the front frame 21 and, on the other hand, a downstream end having a shape complementary to the upstream end of the mobile cover 30.
• the cover 30 having at its free end upstream a protrusion 32 extending along the longitudinal axis of the nacelle, the upstream end of the profile defined by the cavity of the L is the space necessary to receive the protrusion of the cover 30 mobile.
• the aerodynamic external lines of the nacelle are thus smoothed by the presence of the profile 41 between the front frame 21 and the hood 30. During a displacement of the hood 30 mobile upstream, the cover 30 tends to come into contact with the front frame 21 deforming the passage of the profile 41 which is housed under the hood 30.
• composition of the profile 41 may be reinforced with fibers. It can also be provided to introduce rigid elements in certain areas of the profile 41 desired indeformable.
• a heel 33 of suitable shape can be attached to the upstream end of the movable cover 30, in order to improve the contact interface with the aerodynamic continuity element 40 and thus to offer a controlled contact zone.
• a flexible aerodynamic continuity element 40 is observed capable of ensuring the continuity of the internal aerodynamic lines of the nacelle on the side of the cold flow vein.
• this aerodynamic continuity element 40 may take the form of a tongue 42 intended to fill the recess 51 provided in the inner surface of the front frame 21 and, more precisely on the guide structure. of the air flow in the inversion phase.
• This tongue 42 elongate comprises two opposite ends bearing on the inner face of the inner surface of the front frame 21 and a central portion facing the upstream end of the hood 30.
• the central portion of the tongue 42 has a curved section which the curvature is defined so as to ensure the continuity of the internal aerodynamic line of the inner surface of the front frame 21.
• a heel 34 of suitable shape may be attached to the upstream end of the movable cover 30, as in FIGS. 4 to 6.
• the central portion of the tongue may be formed of a double wall to reinforce the aerodynamic continuity element 40. The central portion illustrated in Figures 7 and 8 is thus attached to an inner wall. T shape whose T bar forms the second wall of the double wall.
• the aerodynamic continuity unit 40 comprises a rigid section 43 able to be deformed elastically and intended to ensure the continuity of the aerodynamic external or internal lines of the nacelle. .
• the profile 43 may take the form of a pin-like spring.
• the profile 43 has a general shape in J whose concavity is directed towards the front frame 21. More specifically, it has a first downstream end folded and fixed to the front frame 21, said end being extended by a transition portion intended to allow the inflection of the section 43, it itself extended by a rectilinear upstream end being fixed to the outer surface of the front frame 21 without breaking the slope with the latter.
• the cover 30 When the cover 30 is in a position where the section of the nozzle does not vary, the upstream end of the cover 30 is housed against the transition portion so that the front frame 21, the aerodynamic continuity member 40 and the movable hood then have no slope break on the outer surface of the nacelle.
• the upstream end of the cover 30 abuts and bears against the transition portion of the profile 43 causing the flexion of the latter, thus releasing a passage for advancing the upstream end of the cover 30 to the front frame 21.
• the bending point may advantageously be positioned downstream of its structure as shown in FIGS. 9 to 11. It may also be inverted but requires this a larger exhaust travel.
• the section 43 may be made of metallic material or of composite material having elastic characteristics permitting the bending of part of its structure. Moreover, insofar as the upward displacement of the movable cover 30 causes the reduction of the outer diameter of the profile 43, it is necessary to provide slots which extend from the upstream end of the profile to the zone bending of the latter to allow its deformation. In addition, in order not to reduce the aerodynamic performance of the nacelle by the presence of these slots, they can be filled with flexible material like elastomer.
• the section 44 may take the form of a leaf type spring adapted to ensure continuity of the aerodynamic internal lines of the nacelle on the side of the cold flow vein.
• This profile 44 has a shape similar to that already described with reference to Figures 7 and 8.
• the section 44 is rigid and able to be elastically deformed.
• the upstream end of the movable cover 30 bears against the section 44 which bends inwardly of the front frame 21 so as not to create aerodynamic disturbances. Specifically, a free end of the section 44 not fixed on the inner face of the inner surface of the front frame 21 tilts inside the front frame to allow this bending.
• the fixed portion of connection with the frame of the section 44 can be positioned at the bottom or the top relative to the recess 51. After retreating the cover 30 downstream of the nacelle, the section 44 resumes its initial position.
• a bead 34 can be attached to the upstream end of the movable cover.
• bellows 45, 46 have shapes and dimensions adapted, as in the other embodiments, to ensure the aerodynamic continuity between the front frame 21 and the cover 30 when the cover is not in the nozzle section variation position .
• the upstream end of the cover 30 bears on the bellows 45, 46 which compresses, releasing a passage for the advance further upstream of the hood towards the front frame 21.
• each bellows 45,46 After retreat of the movable cover 30, each bellows 45,46 returns to its original shape.
• the number of corrugations of each bellows is determined so as not to exceed the elastic limit of the latter.
• the aerodynamic continuity assembly 40 intended to ensure continuity the outer and inner lines of the nacelle comprises a non-deformable rigid section 47,48 associated with elastic return means.
• the carrier structure 49 may be either integrated with the front frame 21 as illustrated for the outer surface of the nacelle is independent of the front frame 21 but fixed on the latter as illustrated for the inner surface of the front frame 21.
• the section 47,48 has a shape and dimensions adapted to fill the recesses 50,51 between the front frame 21 and the cover 30 and ensure the continuity of the aerodynamic lines between these elements. Moreover, during a displacement of the hood 30 upstream in the direction of the front frame 21 as illustrated in FIG. 15, the upstream end of the hood 30 bears on the profile 47, 48 which is compressed by means of the spring releasing a passage for the advance of the hood 30 to the front frame 21.
• the direction of movement of the profile is defined by the direction of the contact forces at the interface between the section 47,48 and the upstream end of the movable cover.
• the profile considered may consist of a single piece or an assembly of parts sectors.
• a nacelle proposing a variable nozzle in takeoff and landing by moving the movable cowl beyond its closed position upstream of the fixed structure of the nacelle while not having any aerodynamic line defect direct jet flight phase.
• the invention can be applied to a nacelle not comprising a thrust reverser device.
• the invention can also be applied to a nacelle comprising a thrust reverser device comprising, upstream, thrust reversing flaps under the deflection grilles and an associated movable cowl, downstream of the nacelle, to panels movable to ensure the sectional variation of the ejection nozzle.
• a thrust reverser device comprising, upstream, thrust reversing flaps under the deflection grilles and an associated movable cowl, downstream of the nacelle, to panels movable to ensure the sectional variation of the ejection nozzle.
• the aerodynamic continuity assembly is secured to the movable hood and compressed by the front frame during a movement of the movable cowl upstream of the nacelle.

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• Engineering & Computer Science (AREA)
• Chemical & Material Sciences (AREA)
• Combustion & Propulsion (AREA)
• Mechanical Engineering (AREA)
• Aviation & Aerospace Engineering (AREA)
• General Engineering & Computer Science (AREA)
• Structures Of Non-Positive Displacement Pumps (AREA)
• Wind Motors (AREA)
• Dairy Products (AREA)

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• 2008
  • 2008-09-24 FR FR0805243A patent/FR2936222B1/fr not_active Expired - Fee Related
• 2009
  • 2009-07-10 WO PCT/FR2009/000861 patent/WO2010034893A1/fr active Application Filing
  • 2009-07-10 CA CA2734417A patent/CA2734417A1/fr not_active Abandoned
  • 2009-07-10 US US13/120,703 patent/US8919667B2/en not_active Expired - Fee Related
  • 2009-07-10 CN CN200980132953.0A patent/CN102131705B/zh not_active Expired - Fee Related
  • 2009-07-10 RU RU2011115146/11A patent/RU2532682C2/ru not_active IP Right Cessation
  • 2009-07-10 BR BRPI0917696A patent/BRPI0917696A2/pt not_active IP Right Cessation
  • 2009-07-10 EP EP09784266A patent/EP2326555A1/fr not_active Withdrawn
• 2011
  • 2011-04-15 CO CO11047419A patent/CO6361863A2/es active IP Right Grant

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