• 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:
    In order to reduce the overall mass of an aircraft engine assembly (1), this assembly comprises: a fuselage part (102a) of the aircraft; - A turbomachine (10) comprising a non-faired propeller (14), and an annular row (32) of non-careened exit guide vanes (50b) arranged downstream of the propeller and fixed in rotation relative to each other. a longitudinal axis (5) of the turbomachine; and - an attachment mast (20). According to the invention, at least a portion of the mast leading edge (26) is integrated within said annular row (32) between two exit guide vanes (50b) thereof.
    公开号:FR3050721A1
    申请号:FR1653814
    申请日:2016-04-28
    公开日:2017-11-03
    发明作者:Olivier Pautis；Jerome Colmagro；David Ewens
    申请人:Airbus Operations SAS；
    IPC主号:
    专利说明:

    AIRCRAFT ENGINE ASSEMBLY COMPRISING A MATTRESS ATTACK EDGE INTEGRATED WITH AN ANNULAR ROW OF OUTER CARRIER OUTPUT GUIDELINES
    DESCRIPTION
    TECHNICAL AREA
    The present invention relates to the field of motor assemblies comprising a turbomachine attached to a fuselage portion, via a hooking mast. It more preferably relates to designs in which the turbomachine is attached to the rear lateral part of the fuselage.
    STATE OF THE PRIOR ART
    From the prior art, it is known to report turbomachines at the rear lateral part of the fuselage of the aircraft, these turbomachines comprising a doublet of contra-rotating propellers not careened and also being called "Open Rotor" or "CROR" "Contra Rotative Open Rotor". Such turbomachines are for example known from FR 3,024,125.
    In the so-called "Puller" configuration, in which the propellers are arranged upstream of a gas generator of the turbomachine, a minimum axial distance is provided between the trailing edge of the blades of the rearmost propeller, and the edge Attack of the attachment mast. This minimum distance, for example of the order of one meter, makes it possible in particular to limit the risks of delamination of the boundary layer on the blades of the rearmost propeller.
    However, the implementation of this minimum distance generates a consequent cantilever of the turbomachine on the pylon. The presence of this overhang requires appropriate sizing for the mast and the part of the fuselage supporting this mast, resulting in a large overall mass of the assembly.
    STATEMENT OF THE INVENTION
    To at least partially address this problem, the invention relates to an engine assembly for aircraft comprising: a fuselage part of the aircraft; - A turbomachine comprising a non-faired propeller, and an annular row of non-careened exit guide vanes arranged downstream of the propeller and fixed in rotation with respect to a longitudinal axis of the turbomachine; and a mast for attaching the turbomachine to the fuselage portion, the mast comprising a leading edge of the mast.
    According to the invention, at least a portion of the leading edge of the mast is integrated within said annular row between two exit guide vanes thereof.
    With this retreat of the turbomachine until the integration of the leading edge of the mast in the annular row of exit guide vanes, the overhang of the turbomachine is considerably reduced. The mast and the part of the fuselage concerned are then advantageously less loaded, which makes it possible to reduce their sizing and thus lead to a less overall overall mass for the engine assembly. The invention preferably has at least one of the following optional features, taken singly or in combination.
    A leading end of the at least a portion of the mast leading edge is in a dummy plane P2 passing through the leading edge of the exit guide vanes.
    The attachment mast has a main structure located downstream of the leading edge of the mast, the main structure connecting said fuselage portion to the turbomachine. This main structure has a slender shape, such as an overall shape of rectangular box, and is the main path of effort between the engine and the fuselage. It is conventionally made using side panels, lower and upper side rails, and internal ribs.
    According to a preferred embodiment of the invention, the leading edge of the mast is fixed relative to the main structure of the mast.
    According to another preferred embodiment of the invention, said at least part of the leading edge of the mast is rotatable relative to the main structure of the mast, along a pivot axis substantially parallel to a span direction. the leading edge of the mast.
    It is then preferentially provided for said at least one portion of the leading edge of the mast to be connected to a wedging device at the incidence of the exit guide vanes, said device being designed to vary in incidence the said at least part of the edge of the mast. mast attack, by pivoting it along said pivot axis.
    Preferably, the pivot axis lies in the same fictitious plane as that integrating the radial axes of the exit guide vanes, radial axes in which the outlet guide vanes are designed to be rotated in incidence by said wedging device in incidence .
    Preferably, said at least a portion of the leading edge of the mast has a cross section of generally concave shape, and preferably of general shape substantially identical to that of the cross section of the exit guide vanes.
    According to a preferred embodiment of the invention, said leading edge extends substantially rectilinearly between the turbomachine and the fuselage portion.
    According to another preferred embodiment of the invention, said mast leading edge has a forwardly curved portion forming said at least a portion of the integrated mast leading edge within said annular array of vanes. output guides, the convex portion extending from the turbomachine to the fuselage portion and being extended by a junction portion with this fuselage portion, said domed portion preferably having a leading edge of substantially identical shape to an edge attack of the exit guide vanes.
    Preferably, the joining portion of the leading edge of the mast is substantially rectilinear.
    Preferably, the turbomachine comprises a gas generator arranged downstream of the propeller. In addition, said fuselage portion is preferably a lateral rear portion of this fuselage.
    As mentioned above, the turbomachine also comprises a wedging device at the incidence of the exit guide vanes, this incidence setting device being also designed to vary in incidence the blades of the non-faired propeller, by rotating them along radial axes of these blades. Alternatively, it could be two separate incidence timing devices, controlled synchronously.
    Finally, the invention also relates to an aircraft comprising at least one engine assembly such as that described above. Other advantages and features of the invention will become apparent in the detailed non-limiting description below.
    BRIEF DESCRIPTION OF THE DRAWINGS
    This description will be made with reference to the appended drawings among which; - Figure 1 shows a schematic top view of a portion of an aircraft comprising an engine assembly according to a first preferred embodiment of the invention; - Figure 2 shows a perspective view of the propeller fitted to the motor assembly shown in the previous figure; FIG. 3 represents a perspective view of the annular row of exit guide vanes equipping the motor assembly shown in FIG. 1; - Figure 4a shows a sectional view taken along the line IV-IV of Figure 1; - Figure 4b is a view similar to that of the previous figure, with the leading edge being in the form of another embodiment of the invention; FIG. 4c is a view similar to that of FIG. 4b, with the leading edge of the mast being in the form of another preferred embodiment of the invention; and FIG. 5 shows a view similar to that of FIG. 1, with the leading edge of the mast still in the form of another preferred embodiment of the invention.
    DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS
    Referring firstly to Figure 1, there is shown a rear part of aircraft 100, comprising one or more motor units 1 according to a first preferred embodiment of the invention. More precisely, the aircraft 100 comprises two engine assemblies 1 (only one of which has been shown in full in FIG. 1), each being thus arranged in the rear part of the aircraft.
    Each engine assembly 1 comprises a lateral rear portion 102a of a fuselage 102. It also comprises a turbomachine 10 equipped with a single non-faired propeller 14 in "Puller" mode, that is to say that it comprises a gas generator 12 arranged downstream of the propeller 14. In this regard, it is noted that throughout the description, the terms "upstream" and "downstream" are to be considered with respect to a main direction of flow of gases to through the turbomachines 10, this direction being shown schematically by the arrow 16. Moreover, by convention, the direction X corresponds to the longitudinal direction of the engine assembly 1, which is also comparable to the longitudinal direction of the turbomachine 10 and this assembly 1. This direction X is parallel to a longitudinal axis 5 of the turbomachine 10. On the other hand, the direction Y corresponds to the direction transversely oriented relative to the motor assembly 1 and also assists imilable to the transverse direction of the turbomachine 10, while the direction Z corresponds to the vertical direction or height. These three directions X, Y and Z are orthogonal to each other and form a direct trihedron. The engine assembly 1 also comprises a latching mast 20 of the turbomachine on the fuselage portion 102a. The mast 20, also called EMS (English "Engine Mounting Structure"), comprises a main structure in the form of box 22 dedicated to the recovery of forces from the turbomachine. This box 22 is made in particular with a rear spar 28 from which extends downstream a trailing edge of mast 24, as well as with a front spar 30 from which is upstream an edge d In view from above, the casing 22 extends in the direction Y, of the turbomachine 10 to the fuselage portion 102a.
    Returning to the turbomachine 10 shown in FIG. 1, it is noted that this latter has a single rotating propeller 14 around the axis 5, and not two counter-rotating propellers, as is the case with turbomachines of the Open Rotor plus type. conventional. The turbine engine 10 is also called USE (of the English "Unducted Single Fan"). Nevertheless, it comprises downstream of the propeller 14 an annular row 32 of non-keeled exit guide vanes. This annular row 32 is not rotatably mounted relative to the axis 5, but fixed in rotation relative thereto.
    With reference to FIGS. 2 and 3, the turbomachine comprises a device for stalling in incidence, making it possible to vary in incidence both the blades 50a of the propeller 14 along their radial axes 54a, and the exit guide vanes 50b of the fixed annular row 32, also along their radial axes 54b.
    In FIG. 2 representing the propeller 14, the latter comprises a hub 40 centered on the axis of rotation of the helix, corresponding to the axis 5. This hub 40 comprises a main ring 42 centered on the axis 5 , pierced with several openings 44 circumferentially spaced from each other, and oriented radially relative to this axis 5. The hub 40 also comprises hollow members 46 associated with the openings 44 and extending radially outwardly from the ring main 42. Each hollow member 46 is centered on one of the radial axes 54a and for receiving a blade 50a of the helix. In addition, the helix comprises an outer cowling 52 shown only schematically in FIG. 2, this cowling 52 being centered on the axis 5 and arranged around the hollow members 46. In a conventional manner, the outer surface of this cowling 52 is intended to to be married by the air circulating around the turbomachine, before reaching the blades of the propeller. The propeller 14 thus comprises a plurality of blades 50a, provided in a number for example between eight and twelve. The foot 56 of each blade 50a projects radially inwardly of the hub 40, so that it can be mechanically connected in a conventional manner to the wedging device 38.
    The annular row 32 shown in Figure 3 may have a design identical or similar to that of the propeller 14. Therefore, in Figures 2 and 3, the elements bearing the same reference numerals correspond to the same or similar elements. One of the differences lies in the fact that the hub 40 is not rotatable relative to the axis 5. In addition, one of the particularities of the invention lies in the fact that the annular row 32 has between two blades 50b a free space 60 for the integration of at least a portion of the leading edge of the mast.
    Indeed, returning to Figure 1, it is noted that the leading edge of the mast 26 is integrated in the annular row 32 between two directly consecutive guide vanes.
    In the first embodiment shown in FIGS. 1 to 4a, it is the entire mast leading edge 26 that is integrated with the annular row 32. The leading edge of the mast 26 is here fixed relative to the box 22, and has a conventional section of aerodynamic profile, similar to that of an aircraft wing. In addition, it extends substantially rectilinear in a span direction 62, between the turbomachine 10 and the fuselage portion 102a.
    In this first preferred embodiment, the leading edge of mast 26 is considered as fully integrated with the annular row 32, since it is traversed over its entire length by a notional plane PI passing through the radial axes of the exit guide vanes. 50b.
    In addition, a front end of the leading edge of mast 26 is in a dummy plane P2 passing through the leading edge of the exit guide vanes 50b, this plane P2 being substantially orthogonal to the axis 5. Even more preferably. the leading end of the leading edge of mast 26 is in the same fictitious transverse plane as the leading end of the leading edge of the exit guide vanes 50b. The invention is thus remarkable in that it makes it possible to approach the center of gravity of the turbomachine, referenced 66 in FIG. 1, as close as possible to the casing 22 of the attachment mast 20. This makes it possible to reduce the carrier overhang the turbomachine, and thus reduce the forces that pass through the mast 20 and the fuselage portion 102a. The mass of the motor assembly 1 is advantageously reduced. Preferably, the center of gravity 66 is in the extension of the box 22, outwards in the span direction 62.
    In addition, the turbomachine has been strongly shifted downstream in comparison with known embodiments of the prior art, it results in a decrease in the fuselage section exposed to acoustic pollution caused by the turbomachine. Also, the design according to the invention implies that no harmful interactions exist between the air flow leaving the exit guide vanes 50b, and the leading edge of the mast 26. This advantageously makes it possible to increase the aerodynamic performance as well as the overall performance of the engine assembly 1.
    According to a second preferred embodiment shown in FIG. 4b, the leading edge of mast 26 is no longer of conventional shape, but its cross section is generally concave in shape, with the concavity oriented in the same direction as that of blades 50b. Moreover, it is preferentially made so that the cross section of the leading edge 26 has a general shape substantially identical to that of the cross section of the exit guide vanes 50b, for a better integration of the leading edge of mast 26 within the annular row 32.
    According to a third preferred embodiment shown in FIG. 4c, the shape of the leading edge of mast 26 is retained, but it is arranged mobile in rotation with respect to box 22, along a pivot axis 54b substantially parallel to the span direction 62. Therefore, the mast leading edge 26 can be likened to an output guide vane, and be driven in incidence in the same manner as the other blades of the annular row, through the stalling device 38 to which it is mechanically connected.
    In this case, it is preferentially provided that the pivot axis 54b of the leading edge of the mast 26 is located in the same imaginary plane PI as that integrating the radial axes 54b of the exit guide vanes 50b shown in FIG. 3.
    Finally, FIG. 5 shows a fourth preferred embodiment of the invention, in which only a part 26a of the leading edge of mast 26 is integrated with the annular row 32. It is a portion 26a curved towards the front, in the manner of a leading edge of an exit directional dawn. This curved portion 26a extends from the turbomachine 10 to the fuselage portion, over a radial length substantially identical to that of the exit guide vanes 50b. In addition, the domed portion 26a has a front edge 26a ', similar to its leading edge, whose shape is substantially identical to that of a leading edge of the exit guide vanes 50b.
    The convex portion 26a is extended by a portion 26b of junction with the fuselage portion 102a, this joining portion 26b being substantially rectilinear and parallel to the span direction 62.
    Here too, the front end of the edge 26a 'is in the transversal notional plane P2 passing through the leading edge of the exit guide vanes 50b, and more preferably, the notional transversal plane P2 passes through the front end of the edge attacking guide vanes 50b.
    Of course, various modifications may be made by those skilled in the art to the invention which has just been described, solely by way of non-limiting examples. In particular, the embodiments which have been described above are not exclusive of each other, but can instead be combined with each other.
    权利要求:
    Claims (14)
    [1" id="c-fr-0001]
    An aircraft engine assembly (1) comprising: - a fuselage portion (102a) of the aircraft; - A turbomachine (10) comprising a non-faired propeller (14), and an annular row (32) of non-careened exit guide vanes (50b) arranged downstream of the propeller and fixed in rotation relative to each other. a longitudinal axis (5) of the turbomachine; and - a hooking mast (20) of the turbomachine (10) on the fuselage portion (102a), the mast comprising a mast leading edge (26), characterized in that at least a portion of the edge mast driver (26, 26a) is integrated within said annular row (32) between two exit guide vanes (50b) thereof.
    [2" id="c-fr-0002]
    2. Engine assembly according to claim 1, characterized in that a front end of said at least a portion of the leading edge of the mast (26, 26a) is in a fictitious plane (P2) crossing the leading edge. guide vanes (50b).
    [3" id="c-fr-0003]
    3. Engine assembly according to claim 1 or claim 2, characterized in that the leading edge of the mast (26) is fixed relative to a main structure (22) of the mast.
    [4" id="c-fr-0004]
    4. Engine assembly according to claim 1 or claim 2, characterized in that said at least a portion of the mast leading edge (26, 26a) is rotatable relative to a main structure (22) of the mast, along a pivot axis (54b) substantially parallel to a span direction (62) of the mast leading edge (26).
    [5" id="c-fr-0005]
    5. Engine assembly according to claim 4, characterized in that said at least a portion of the leading edge of the mast (26, 26a) is connected to a device (38) for wedging in incidence of the exit guide vanes (50b). said device (38) being adapted to cause the at least one portion of the leading edge of the mast (26, 26a) to vary in incidence by pivoting it along said pivot axis (54b).
    [6" id="c-fr-0006]
    6. Motor assembly according to claim 4 or claim 5, characterized in that the pivot axis (54b) lies in the same fictitious plane (PI) as that integrating the radial axes (54b) of the exit guide vanes ( 50b), radial axes in which the outlet guide vanes are adapted to be rotated in incidence by said incidence wedging device (38).
    [7" id="c-fr-0007]
    7. Motor assembly according to any one of the preceding claims, characterized in that said at least a portion of the leading edge of mast (26, 26a) has a cross section of generally concave shape, and preferably generally of substantially identical to that of the cross section of the exit guide vanes (50b).
    [8" id="c-fr-0008]
    8. Motor assembly according to any one of the preceding claims, characterized in that said leading edge of the mast (26) extends substantially rectilinearly between the turbomachine (10) and the fuselage portion (102a).
    [9" id="c-fr-0009]
    9. Engine assembly according to any one of claims 1 to 7, characterized in that said leading edge of mast has a convex portion forward (26a) forming said at least a portion of the leading edge of the mast. integrated within said annular row (32) of exit guide vanes (50b), the convex portion (26a) extending from the turbomachine (10) to the fuselage portion (102a) and being extended by a joining portion (26b) with this fuselage portion, said domed portion (26a) preferably having a leading edge (26a ') of substantially identical shape to a leading edge of the exit guide vanes (50b).
    [10" id="c-fr-0010]
    10. Engine assembly according to claim 9, characterized in that the junction portion (26b) of the leading edge of the mast (26) is substantially rectilinear.
    [11" id="c-fr-0011]
    11. Engine assembly according to any one of the preceding claims, characterized in that the turbomachine (10) comprises a gas generator (12) arranged downstream of the propeller (14).
    [12" id="c-fr-0012]
    12. Engine assembly according to any one of the preceding claims, characterized in that said fuselage portion (102a) is a lateral rear portion of the fuselage.
    [13" id="c-fr-0013]
    13. Motor assembly according to any one of the preceding claims, characterized in that the turbomachine also comprises a device (38) for wedging the incidence of the guide vanes (50b), said incidence wedging device (38) being of preferably also designed to vary in incidence the blades (50a) of the non-faired propeller (14), by rotating them along radial axes (54a) of these blades.
    [14" id="c-fr-0014]
    14. Aircraft (100) comprising at least one engine assembly (1) according to any one of the preceding claims.
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    法律状态:
    2017-04-19| PLFP| Fee payment|Year of fee payment: 2 |
    2017-11-03| PLSC| Publication of the preliminary search report|Effective date: 20171103 |
    2018-04-20| PLFP| Fee payment|Year of fee payment: 3 |
    2019-04-18| PLFP| Fee payment|Year of fee payment: 4 |
    2020-04-20| PLFP| Fee payment|Year of fee payment: 5 |
    2021-04-23| PLFP| Fee payment|Year of fee payment: 6 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1653814|2016-04-28|
    FR1653814A|FR3050721B1|2016-04-28|2016-04-28|AIRCRAFT ENGINE ASSEMBLY COMPRISING A MATTRESS ATTACK EDGE INTEGRATED WITH AN ANNULAR ROW OF OUTER CARRIER OUTPUT GUIDELINES|FR1653814A| FR3050721B1|2016-04-28|2016-04-28|AIRCRAFT ENGINE ASSEMBLY COMPRISING A MATTRESS ATTACK EDGE INTEGRATED WITH AN ANNULAR ROW OF OUTER CARRIER OUTPUT GUIDELINES|
    US15/496,183| US10556699B2|2016-04-28|2017-04-25|Aircraft engine assembly comprising a pylon leading edge incorporated with an annular row of unfaired after-guide vanes|
    GB1706723.2A| GB2551882B|2016-04-28|2017-04-27|Aircraft engine assembly comprising a pylon leading edge incorporated with an annular row of unfaired after guide vanes|
    CN201710290106.3A| CN107336837A|2016-04-28|2017-04-28|Engine pack and corresponding aircraft for aircraft|
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