• 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:
    To optimize the risk management of propeller blade burst, the invention proposes an aircraft part comprising a fuselage defining a pressurized compartment and a turbomachine comprising at least one propeller (20) intended to be rotated. along a helical axis of rotation (18) and comprising blades (28) variable pitch, each blade comprising a foot (28a) and an aerodynamic portion (28b). According to the invention, the propeller comprises a structure (34) for radial retention of blades in case of bursting thereof, this structure (34) being connected to the outer radial end of the aerodynamic portion (28b) of each blade by means of a mechanical connection (36) allowing pivoting along a wedge axis (30) of the blade (28), this propeller (20) being moreover arranged in a transversal notional plane passing through the pressurized compartment ( 16).
    公开号:FR3030445A1
    申请号:FR1463128
    申请日:2014-12-22
    公开日:2016-06-24
    发明作者:Julien Guillemaut；Guillaume Gallant
    申请人:Airbus Operations SAS；
    IPC主号:
    专利说明:

    [0001] BACKGROUND OF THE INVENTION The present invention relates to the field of aircraft turbomachine propellers. BACKGROUND OF THE INVENTION It relates more particularly to the management of the propeller blade burst risk, also known as the UERF risk, of the English "Uncontained Engine Rotor Failure".
    [0002] The invention applies to all types of aircraft, commercial and military, including turbomachines with propellers such as turboprops or turbine turbomachines with twin counter-rotating propellers, also called "open rotor" turbomachines. STATE OF THE PRIOR ART From the prior art, several types of propeller turbomachines are known, of which the blades of the propeller are of variable pitch. Such propellers are used in conventional turboprop engines, but also turbomachines with twin blades of contra-rotating propellers called "open rotor" propellers. A turbomachine propeller is for example known from WO20100116080.
    [0003] To meet specific standards, aircraft must provide a solution to the risk of propeller blades bursting. In operation, such a burst blade may indeed generate debris likely to damage the neighboring turbine engine, and / or damage the fuselage of the aircraft, for example at a pressurized compartment delimited by this fuselage. To avoid damage to the fuselage, particularly when defining a pressurized compartment such as the passenger cabin, the fuselage may be equipped with an anti-debris shield, which extends over a large angular sector of the portion of the fuselage located to the right of the propeller. Nevertheless, this shield is penalizing in terms of mass and costs.
    [0004] Another alternative, for the particular case of aircraft equipped with turbomachines reported at the rear fuselage, is to move these turbomachines back so that their propellers are no longer located in the right of a pressurized compartment, but behind the waterproof bottom delimiting this compartment. However, this solution requires extending the tail of the aircraft and therefore also leads to penalties in terms of weight and cost. There is therefore a need for optimizing the design of propeller turbomachinery vis-à-vis risk management UERF. SUMMARY OF THE INVENTION To at least partially meet this need, the invention firstly relates to an aircraft part comprising a fuselage defining a pressurized compartment, preferably a passenger cabin, and at least one turbomachine attached to the fuselage by means of an attachment pylon, said turbomachine comprising at least one propeller intended to be rotated about an axis of propeller rotation and having variable pitch blades, each blade comprising a foot and a aerodynamic part. According to the invention, the propeller further comprises a radial retention structure of blades in the event of their bursting, said retention structure extending around the axis of propeller rotation and being connected to the propeller. external radial end of the aerodynamic portion of each blade by means of a mechanical connection allowing pivoting along a pinning axis of the blade, and the helix is arranged in a transverse fictitious plane passing through said pressurized compartment. By judiciously integrating the blade retention structure with the propeller, the invention is remarkable in that it allows a saving in weight and cost. Indeed, any blade debris being retained radially by the dedicated structure, it is not necessary to integrate a shield on the fuselage to the right of the propeller, or unnecessarily lengthen the rear tip of the aircraft when propeller turbomachines are reported on this tip. In addition, the risk of damaging a neighboring turbine engine by blade debris is reduced to nothing, still due to the radial retention of these debris by the specific structure of the invention. The invention preferably has at least one of the following optional features, taken singly or in combination.
    [0005] According to a first preferred embodiment, said mechanical connection comprises a pin integral with the retention structure and extending along the axis of alignment of the associated blade, said pin being housed pivotally in an orifice of the outer radial end of the the aerodynamic part of the blade. According to a second preferred embodiment, said mechanical connection comprises a pin secured to the outer radial end of the aerodynamic portion of the associated blade, said pin extending along the axis of wedging of the blade and housed pivotally in an orifice of the retention structure. In this case, the retention structure then preferably comprises two axial portions fixedly attached to one another, each of the two axial portions defining half of each of the orifices. It is also possible to provide that the two axial portions are each made using several angular sectors fixedly attached to each other, and that the angular sectors of one of the two axial portions are angularly offset from the angular sectors of each other. the other axial part. Preferably, the retention structure is a structure extending continuously around the axis of rotation of the helix, that is to say in a closed manner, without discontinuity, and this retention structure is for example realized from several parts attached to each other. Preferably, the retention structure is of generally annular shape, with an axial half-section in the form of an aerodynamic profile.
    [0006] Preferably, said turbomachine comprises two counter-rotating propellers, each arranged in a transverse fictitious plane passing through said pressurized compartment. Preferably, said turbomachine is hung on a rear part of the fuselage via the attachment pylon located in front of or behind the sealed bottom delimiting a passenger compartment of the aircraft.
    [0007] Finally, the subject of the invention is an aircraft comprising such a part. Other advantages and features of the invention will become apparent in the detailed non-limiting description below. BRIEF DESCRIPTION OF THE DRAWINGS The invention will be better understood on reading the following detailed description, non-limiting examples of implementation thereof, as well as on examining the appended drawings among which: - Figure 1 shows a top view of an aircraft comprising a rear portion according to a preferred embodiment of the invention; Fig. 1a is a sectional view taken along the line A-A of Fig. 1; - Figure lb is a view similar to that of Figure la, according to an alternative embodiment; - Figure 2 shows a front view of a turbomachine propeller equipping the rear portion shown in the preceding figures; FIG. 3 represents an axial half-sectional view of the helix shown in the preceding figure, according to a first preferred embodiment of the invention; - Figure 4 shows a top view of that of the previous figure; - Figure 5 shows a view similar to that of Figure 3, after a blade burst; - Figure 6 is an axial half-sectional view showing a detailed example for the realization of the blade retention structure equipping the propeller; FIG. 7 is a view similar to that of FIG. 3, with the helix being in the form of a second preferred embodiment of the invention; - Figure 8 shows a top view of that of the previous figure; - Figure 9 shows a similar view to the previous, according to an alternative embodiment; FIG. 10 represents a view similar to that of FIG. 1, with propeller turbomachines arranged differently on the fuselage of the aircraft; and FIG. 10a is a sectional view taken along line AA of FIG. 10. DETAILED DESCRIPTION OF PARTICULAR EMBODIMENTS Referring to FIGS. 1 and 1a, there is shown a rear portion 1a of an aircraft 1 of FIG. commercial aircraft type, according to the present invention.
    [0008] Conventionally, the rear portion la comprises a fuselage 4, a vertical stabilizer 6 also called "vertical drift", and two horizontal stabilizers 7. It also comprises a sealed bottom 14 for separating, in the longitudinal direction of the aircraft, a pressurized compartment 16 forming a passenger cabin, a non-pressurized rear compartment 17. In this regard, it is noted that in the description, the terms "before" and "rear" are to be considered in a direction of advance 3 of the aircraft, following the thrust exerted by the turbomachines. The two compartments 16, 17 are of course delimited by the fuselage 4. The rear portion is also equipped with two propeller turbomachines 10, reported laterally on the fuselage 4 via conventional attachment poles 13. The two turbomachines 10 each have a counter-rotating pair of counter-rotating propellers, in the manner of so-called open-rotor, so-called "open rotor" turbomachines. However, as will become clear in the following, the propellers here have the particularity of each comprising a radial blade retention structure which forms a shield around the blades of the propeller. In operation, this shield rotates with the rest of the helix along an axis of rotation 18 of the propeller. In this configuration in which the turbomachines 10 are arranged on the rear part of the fuselage 4, the two contra-rotating propellers 20 are each arranged in a transverse fictional plane Pl, passing through the pressurized compartment 17 and the fuselage portion 4a located in front of the bottom. 14. The two helical planes P1 associated with each turbomachine 10 are arranged transversely one behind the other, while remaining in front of the sealed bottom 14. This does not cause any problem with respect to UERF risk, because the blade retention structure, which will be described later, prevents the projection of blade debris to the fuselage and to the other turbomachine 10. This allows in particular to shorten the length of the non-pressurized compartment 17, with a gain significant mass that also results from the fact that no protective shield is needed on the fuselage, on and near Pl plans.
    [0009] Furthermore, it is noted that each turbomachine 10 is designed in the so-called "puller" configuration, the particularity of which lies in the fact that the pair of counter-rotating propellers is situated in front of a gas generator 22 of the turbomachine, causing the doublet of propellers. The gas generator 22 is generally located behind the sealed bottom 14, since there is at least one transverse plane AA passing through the portion of the fuselage 4b located behind the sealed bottom 14. Conventionally, this part 4b of the fuselage located behind the sealed bottom 14 is referred to as the tail or tail of the aircraft. In the configuration of Figure la, the turbomachines 10 arranged on either side of the fuselage 4 are connected thereto by poles 13 which are laterally assembled on the same fuselage. A beam 15 connecting these two masts 13 passes through the non-pressurized compartment 17. In another configuration shown in FIG. 1b, the two turbomachines 10 are mounted in the upper part of the fuselage 4b, by means of a connecting beam 15 which crosses over the compartment 17, but which is assembled on this upper part of the fuselage. Moreover, in this configuration, the beam 15 and the two masts 13 can be made using one and the same structure, preferably straight. Referring now to Figures 2 to 4, there is shown one of the propellers 20 according to a first preferred embodiment of the invention. The other propellers of the aforementioned turbomachines are of identical or similar design.
    [0010] The propeller 20, intended to be rotated along the axis 18, comprises a hub 26 from which radially extend a plurality of blades 28 spaced angularly from each other. Conventionally, these blades are variable pitch, that is to say that each of them can be rotated according to a pinning axis of its own. This pinning axis is identified by the numeral 30 in FIGS. 3 and 4. The means for controlling the pitch of the blades are conventional, and will therefore not be further described. These means generally cooperate with a foot 28a of the blade which is housed in the hub of the propeller, and from which extends an aerodynamic portion 28b of the blade.
    [0011] As mentioned above, one of the peculiarities of the invention lies in the fact that the helix further comprises a structure 34 of radial retention of blades, in case of bursting thereof. The retention structure 34 is generally annular, extending around the axis of rotation of the helix 18. In axial half-section, the structure 34 has a shape of aerodynamic profile, as is best seen in Figure 3 In other words, it is a closed structure having two surfaces, namely an extrados surface and an intrados surface interconnected at a leading edge and at the level of a trailing edge. The annular structure 34 surrounds the blades 28, since it is connected to the outer radial end of the aerodynamic portion 28b of each blade 28. This connection is made by means of a mechanical connection 36 allowing pivoting according to the invention. In this first preferred embodiment, the mechanical connection 36 comprises a pin 38 integral with the retention structure 34, and extending radially inwards along the clamping axis 30. pin 38, projecting from the retention structure 34, is housed pivotally in an orifice 40 of the radial outer end of the aerodynamic portion 28b of the blade. As can be seen in FIG. 3, the orifice 40 opens radially outwards, facing the underside surface of the retention structure 34. Alternatively, in this first preferred embodiment, the pin 38 is the retention structure 34 may not be integral with the latter, but pivotally mounted on the same structure 34.
    [0012] For the assembly of the helix, the annular retention structure 34 is slid axially around the blades 28 mounted on the hub 26, then the pins 38 are mounted on this structure 34 from outside thereof, being introduced in the orifices 40 of the blades.
    [0013] In the event of a blade burst as shown schematically in FIG. 5, a blade debris 28 'is generated radially outwardly with respect to the rupture 42. This debris 28' is advantageously retained radially by the annular structure 34, which implies that this debris can not be ejected from the helix in this radial direction. At most, the blade debris 28 'can be extracted from the pin 38 and be discharged axially out of the helix, but preferably, it is designed so that after the rupture, the debris 28' remains in position. Place by the pin 38. As an illustrative example, it is shown in Figure 6 an embodiment of the annular retention structure 34, shaped aerodynamic profile. First of all, a core 46 made of fibers with high mechanical strength is provided. This core is the part intended to house the pin 38. A reinforcing structure 48 extends around the core 46, except at a front spar thereof. Indeed, the reinforcing structure 48 extends to the leading edge of the profile, and is covered at this point by a foil material 50 for corrosion resistance. A skin 52 surrounds the assembly while running along the profile, the remaining parts are filled by a filling material 54 at low density, for example foam. Referring now to Figures 7 and 8, there is shown a second preferred embodiment of the helix 20. This second mode differs from the first mode in that the components of the mechanical link 36 are reversed, since the pin 38 is now secured to the outer radial end of the aerodynamic portion 28b of the blade. Also, this pin which extends along the axis of wedging 30 is housed pivotally in an orifice 40 of the retention structure 34, this orifice being through or opening radially inwards. To facilitate assembly of the components of the propeller, in particular its annular retention structure 34, it is manufactured in two axial portions 34a, 34a fixedly attached to one another by conventional means, for example to using welds. Each axial portion 34a then defines half of each of the orifices 40. Thus, for assembly, the two parts 34a, 34a are slid axially towards each other and centered on the axis of rotation 18, so that each peg 38 already in place on the blades 28 is surrounded on either side by two half-orifices 40, respectively. According to an alternative embodiment shown diagrammatically in FIG. 9, each axial portion 34a is made by means of several angular sectors 34 'fixedly affixed to one another, also conventionally, for example by means of welds. In addition, it is preferably made to have a staggered type arrangement between the sectors 34 'of the front portion 34a, and the sectors 34' of the rear portion 34a. In other words, the angular sectors 34 'of the front axial portion 34a are angularly offset from the angular sectors 34' of the rear axial portion 34a, the length of the angular offset corresponding to half a sector length 34 '.
    [0014] Finally, in Figures 10 and 10a, there is shown another embodiment of the invention wherein the turbine engine 10 propellers are arranged further forward than in the previous configurations, always being arranged on either side of the fuselage. The masts 13 are here assembled at the portion 4a of the fuselage delimiting the passenger cabin 16, and no longer at the portion 4b forming the rear tip of the aircraft. The masts 13 are therefore arranged forward with respect to the sealed bottom 14. Also, to avoid damage to the fuselage at the pressurized compartment 16, the connecting beam 15 is assembled on the upper part of the fuselage 4a delimiting the cabin . In this configuration, the gas generator 22 is also located in front of the sealed bottom 14, since any transverse fictional plane AA of this gas generator 22 passes through the portion of the fuselage 4a located in front of the sealed bottom 14 In this configuration of Figures 10 and 10a, the design of the helices 20 is the same as or similar to that described above.
    [0015] Whatever the embodiment envisaged, in addition to the main advantages described above, the retention structure confers the following additional advantages. First of all, in the event of a blade burst, it is noted that the retention of the debris makes it possible to limit the level of vibrations after this bursting, since the imbalance that derives therefrom is lower. This makes it possible to improve the safety of the subsequent phase of rotation of the propeller, without having to implement heavy and expensive means dedicated to the management of this problem of vibrations, such as flexible fasteners known by the English name " soft engine mount ».
    [0016] The retention structure also limits the risks associated with the impacts of ice that is likely to grow on the blades, sliding along them. Indeed, the ice is also retained radially by the specific structure of the invention. This radial blade retention structure can also be functionalized to ensure the balancing of the blades, for example by means of sliding adjustable masses arranged on the trailing edge of this annular structure. Of course, various modifications may be made by those skilled in the art to the invention which has just been described, solely as non-limiting examples.
    权利要求:
    Claims (10)
    [0001]
    REVENDICATIONS1. Aircraft part (1a) comprising a fuselage (4) defining a pressurized compartment (16), preferably a passenger cabin, and at least one turbomachine (10) attached to the fuselage by means of an attachment mast (13), said turbomachine comprising at least one propeller (20) intended to be driven in rotation along an axis of propeller rotation (18) and comprising blades (28) with variable pitch, each blade comprising a foot (28a) and an aerodynamic portion (28b), characterized in that the propeller (20) further comprises a structure (34) for radial retention of blades in case of burst thereof, said retention structure (34) is extending about the helical axis of rotation (18) and being connected to the outer radial end of the aerodynamic portion (28b) of each blade through a mechanical link (36) for pivoting along an axis setting (30) of the blade (28), and in that said propeller (20) is arranged in s a transversal fictitious plane (P1) passing through said pressurized compartment (16).
    [0002]
    2. Aircraft part according to claim 1, characterized in that said mechanical connection (36) comprises a pin (38) integral with the retention structure (34) and extending along the pinning axis (30) of the associated blade, said pin (38) being housed pivotally in an orifice (40) of the radial outer end of the aerodynamic portion (28b) of the blade (28).
    [0003]
    3. Aircraft part according to claim 1, characterized in that said mechanical connection (36) comprises a pin (38) integral with the outer radial end of the aerodynamic portion (28b) of the associated blade (28), said pin (38) extending along the axis of wedging (30) of the blade and housed pivotally in an orifice (40) of the retention structure (34).
    [0004]
    Aircraft part according to one of the preceding claims, characterized in that the retention structure (34) is a structure extending continuously around the axis of rotation of the propeller (18), and preferably made from several parts attached to each other.
    [0005]
    5. Aircraft part according to claim 3 combined with claim 4, characterized in that the retention structure (34) comprises two axial parts (34a, 34a) fixedly attached to one another, each of the two parts. axial defining half of each of the orifices (40).
    [0006]
    6. Aircraft part according to claim 5, characterized in that the two axial portions (34a, 34a) are each produced by means of a plurality of angular sectors (34 ') fixedly attached to each other, and in that the angular sectors (34 ') of one of the two axial portions are angularly offset from the angular sectors (34') of the other axial part.
    [0007]
    7. Aircraft part according to any one of the preceding claims, characterized in that the retention structure (34) is of generally annular shape, with an axial half-section shaped aerodynamic profile.
    [0008]
    8. Aircraft part according to any one of the preceding claims, characterized in that said turbomachine (10) comprises two contra-rotating propellers (20, 20), each arranged in a transverse fictitious plane (P1) passing through said pressurized compartment (16). ).
    [0009]
    9. Aircraft part according to any one of the preceding claims, characterized in that said turbomachine (10) is hooked on a rear part of the fuselage (4) through the attachment pylon (13).
    [0010]
    10. Aircraft (1) having a portion (1a) according to any one of the preceding claims.
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    同族专利:
    公开号 | 公开日
    FR3030445B1|2017-01-27|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US2366795A|1942-10-10|1945-01-09|George R Lamorcaux|Propeller|
    EP0244515A2|1986-04-16|1987-11-11|The Boeing Company|Hoop fan jet engine|
    US5096382A|1989-05-17|1992-03-17|Gratzer Louis B|Ring-shrouded propeller|
    EP2774852A2|2013-03-05|2014-09-10|Rolls-Royce plc|Gas turbine engine installation|FR3054527A1|2016-07-29|2018-02-02|Airbus Operations|AIRCRAFT ASSEMBLY COMPRISING A PROTECTIVE SHIELD AGAINST MOTOR SHOCK, MOUNTED ON THE CASING OF A TURBOMACHINE MODULE|
    FR3054526A1|2016-07-26|2018-02-02|Safran Aircraft Engines|AIRCRAFT COMPRISING AN INTEGRATED REAR-FUSELAGE TURBOREACTOR COMPRISING A FITTING FOR EJECTING OF BLADES|
    EP3315787A1|2016-10-28|2018-05-02|Rolls-Royce Corporation|Thrust ring and rotor fan system|
    DE102019006484B3|2019-09-11|2020-08-06|Friedrich Grimm|MANIFOLDED POWER PLANT WITH AT LEAST ONE TORQUE LEVEL|
    US11060527B2|2016-10-28|2021-07-13|Rolls-Royce Corporation|Thrust ring and rotor fan system with passive leading edge slats|
    法律状态:
    2015-12-21| PLFP| Fee payment|Year of fee payment: 2 |
    2016-06-24| PLSC| Publication of the preliminary search report|Effective date: 20160624 |
    2016-12-22| PLFP| Fee payment|Year of fee payment: 3 |
    2017-12-21| PLFP| Fee payment|Year of fee payment: 4 |
    2019-12-19| PLFP| Fee payment|Year of fee payment: 6 |
    2020-12-23| PLFP| Fee payment|Year of fee payment: 7 |
    2021-12-24| PLFP| Fee payment|Year of fee payment: 8 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1463128A|FR3030445B1|2014-12-22|2014-12-22|AIRCRAFT TURBOMACHINE PROPELLER, COMPRISING A BLADE RETENTION STRUCTURE CONNECTED TO THE EXTERNAL RADIAL END OF EACH BLADE|FR1463128A| FR3030445B1|2014-12-22|2014-12-22|AIRCRAFT TURBOMACHINE PROPELLER, COMPRISING A BLADE RETENTION STRUCTURE CONNECTED TO THE EXTERNAL RADIAL END OF EACH BLADE|
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