• 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 further benefit from the principle of ingestion of the boundary layer by engines of an assembly (1) for aircraft, the invention provides that the rear fuselage portion (10) of this assembly comprises a front portion (12) which splits into at least two separate rear parts (14) spaced from each other and each incorporating the rotary ring of the receiver (18) of one of the motors (2).
    公开号:FR3052743A1
    申请号:FR1655719
    申请日:2016-06-20
    公开日:2017-12-22
    发明作者:Jean-Michel Rogero；Camil Negulescu；Renaud Faure
    申请人:Airbus Operations SAS；
    IPC主号:
    专利说明:

    AIRCRAFT ASSEMBLY COMPRISING PROPULSION ENGINES BY INGESTION OF THE LIMIT LAYER
    DESCRIPTION
    TECHNICAL AREA
    The present invention relates to the field of aircraft comprising a rear fuselage portion equipped with propulsion engines by ingestion of the boundary layer. In a known manner, propulsion by ingestion of the boundary layer corresponds to ingestion by the motors of a flow of low kinetic energy air circulating around the rear fuselage portion. This technique reduces the kinetic energy expended for propulsion as well as the drag of the aircraft, resulting in a decrease in fuel consumption.
    STATE OF THE PRIOR ART
    It is known to report, in the rear fuselage portion, engines propelled by ingestion of the boundary layer. This is for example two semi-buried engines arranged side by side, projecting upwardly or laterally from the rear fuselage portion.
    However, in this type of configuration, the two engines can only ingest part of the air boundary layer flowing on the rear fuselage portion. For these configurations, the boundary layer is also ingested non-axisymmetrically with respect to the axis of the air inlet, thus generating a distortion of the inlet flow of the engine.
    There is therefore a need for optimization to benefit more from the principle of propulsion by ingestion of the boundary layer.
    STATEMENT OF THE INVENTION
    To at least partially meet this need, the invention relates to an assembly for aircraft comprising a rear fuselage portion and at least two propulsion engines by ingestion of the boundary layer flowing on the rear fuselage portion, each engine comprising a receiver equipped with a rotating ring from which bladed elements protrude radially outwards. According to the invention, said rear fuselage portion comprises a front portion which splits into at least two distinct rear parts spaced from each other and each incorporating the rotary ring of one of said engines. The invention thus provides for the separation of the fuselage into several rear parts each of which is associated with a motor, so that its receiver can ingest all of the boundary layer flowing on its associated rear portion. This advantageously results in overall performance gains of the aircraft. The invention also provides for the implementation of the following optional features, taken singly or in combination.
    Each rear portion comprises successively, from front to rear - a fuselage front section of convergent shape towards the rear; - the rotating ring; and - a rear section.
    The receivers of said motors are spaced from each other in a transverse direction and / or in a direction of the height of said assembly. Optionally, the two receivers of two motors are spaced apart from one another in a longitudinal direction, so that a distance between the two longitudinal parallel axes of the two distinct rear parts respectively carrying the two receivers, is less than the sum a radius of the bladed elements of one of the two receivers and a radius of the bladed elements of the other receiver.
    Each engine comprises a gas generator driving the receiver of said engine, said gas generator comprising a compressor assembly, a combustion chamber and a turbine assembly, said gas generator being preferably arranged forward with respect to the receiver. Alternatively, the gas generator could be placed at the rear of the receiver.
    Each motor preferably has an inverted design in which the turbine assembly is located in front of the compressor assembly, the exhaust gas discharge ports through the rear fuselage portion being preferably arranged in front of the gas generator .
    For two motors facing each other, the respective longitudinal axes of the two gas generators are inclined with respect to a longitudinal direction of the assembly, so that a separation distance between the two compressor assemblies is less than at a separation distance between the two turbine units.
    The respective longitudinal axes of the two gas generators are inclined with respect to the longitudinal direction of the assembly, so that the turbine disks of the turbine assembly of one of said motors are inscribed in the transverse notional turbine planes. which do not intercept the gas generator of the other of said engines, and vice versa.
    Each receiver is a non-faired propeller, whose blades are preferably variable pitch.
    Alternatively, each receiver is a fan surrounded by a nacelle connected to the associated rear fuselage portion, via front support arms and / or output guide vanes.
    Each nacelle is structural and configured to ensure transmission of forces from one or more empennages of the assembly, towards the associated rear fuselage portion, and / or the assembly comprises at least one beam passing through 'efforts between a tail and the rear fuselage portion.
    The nacelles are mechanically connected to each other.
    Each nacelle comprises thrust reverser means, preferably comprising thrust reversal grids covered by one or more movable covers. The set includes two empennages.
    Preferably, the two empennages and the nacelles are crossed by the same transversal fictional plane of the assembly.
    Finally, the invention also relates to an aircraft comprising an assembly such as that described above, the aircraft being preferably of the commercial type. 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; FIG. 1 represents a perspective view of an aircraft according to the invention; - Figure 2 shows an enlarged perspective view of an assembly according to a preferred embodiment of the invention, the assembly belonging to the aircraft shown in the previous figure; - Figure 3 is a perspective view similar to that of Figure 2, specifically showing the rear fuselage parts of the assembly; FIG. 3a is a view similar to that of FIG. 3, seen from above, showing an alternative embodiment; - Figure 4 is a top view of the assembly shown in Figure 2; Figure 5 is a sectional view taken along the line V-V of Figure 4; - Figure 5a is a sectional view taken along the line Va-Va of Figure 4; FIGS. 5b and 5c respectively show a front view of another tail configuration, and a still perspective view of another configuration of empennages; Figure 6 is a sectional view taken along the line VI-VI of Figure 5; FIG. 7 is a view from above of an assembly according to another preferred embodiment of the invention; - Figure 8 is a sectional view similar to that of Figure 6, showing thrust reversal means; - Figure 9 is a top view of the assembly shown in Figure 8; and - Figure 10 is a sectional view similar to that of Figure 6, with the assembly being in accordance with yet another preferred embodiment of the invention.
    DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS
    Referring firstly to Figure 1, there is shown an aircraft 100 of the commercial type, comprising a set 1 corresponding to its rear end, provided with engines 2. On this aircraft, the wings 4 are not equipped with engines, even if this could be the case, without departing from the scope of the invention. The motors 2 are only arranged on the assembly 1, a preferred embodiment of which will now be described with reference to FIGS. 2 to 6. In these figures, the terms "front" and "rear" are to be considered in relation to a direction of advance 8 of the aircraft, following the thrust generated by the motors 2. The assembly 1 comprises a rear fuselage portion 10, forming the rear end of the fuselage of the aircraft. This portion 10 has a front portion 12, a front end 12a has a fuselage for example oval, circular or other shape.
    Going towards its rear end 12b, the front portion 12 is progressively pinched at its center until it splits into two distinct rear fuselage parts, referenced 14. The two rear portions 14, preferably shaped identical and revolutionary, are spaced apart from each other in a transverse direction Y of the assembly. In this regard, it is stated that by convention, the direction X corresponds to the longitudinal direction of the assembly 1, which is also comparable to the longitudinal direction of each engine of this set 1. This direction X is parallel to a longitudinal axis 5 On the other hand, the direction Y corresponds to the direction transversely oriented relative to the assembly 1 and also comparable to the transverse direction of each engine, while the direction Z corresponds to the vertical direction or the height. These three directions X, Y and Z are orthogonal to each other and form a direct trihedron.
    Each rear portion 14 is intended to integrate all or part of one of the motors 2. Therefore, in the preferred embodiment which provides two motors spaced in the direction Y, there are two rear parts 14. In a different case where it would be added a third motor spaced from the first two in each of the directions Y and Z so as to have a triangular arrangement, it would then be provided three rear fuselage parts. In the case of four engines, the latter could be arranged in square or rectangle, being respectively integrated to four rear fuselage parts 14.
    In the preferred embodiment shown in FIGS. 2 to 6, two fuselage rear portions 14 spaced from each other in the Y direction, and running parallel in the X direction from the end, are provided. rear 12b of the fuselage portion 12. From this end 12b, each rear portion 14 has first a front section 14-1 which narrows, for example frustoconical or similar, converging rearwardly. This part 14 then integrates an element of its associated engine 2, as will be explained hereinafter, then ends backwards by a rear section 14-2 of circular section and of substantially constant diameter, in the form of a warhead, convergent or more complex form
    In this preferred embodiment, each fuselage rear portion 14 is centered on the longitudinal axis 5 of its associated engine 2. Each engine is here of the jet engine type by ingestion of the boundary layer flowing on the corresponding rear fuselage part. 14. Referring more specifically to Figure 4, each engine 2 is thus equipped with a gas generator 16 driving a receiver 18. The generator 16 is arranged forward with respect to the receiver 18, which allows it to be integrated wholly or partly inside the converging front section 14-1 of the rear fuselage portion 14. This avoids the presence of a large mass at the rear end of the aircraft, and facilitates the balancing of the latter while reducing the balancing drag.
    The gas generator 16 has a so-called inverted design, in which a turbine assembly 20 is arranged in front of a compressor assembly 22, with a combustion chamber 24 located between them. As has been schematized in FIG. 6, this makes it possible to arrange orifices 28 for evacuating the hot gases coming from the turbine assembly 20, in front of the generator. These orifices 28 passing through the front fuselage section 14-1 thus adopt an advanced position, offering several advantages.
    Firstly, since the hot gases are ejected very far upstream through the orifices 28, their cooling is favored by a mixture with the ambient air over a considerable length, before any impact of these gases on rear parts of the aircraft.
    In addition, this arrangement very upstream of the orifices 28 simplifies the use of energy recovery systems, with the benefits of increasing the efficiency of the engine and power generation for the cabin of the aircraft.
    The receiver is here a turbomachine blower, which comprises a rotary ring 30 also called fan hub, from which project bladed elements 32 referred to as fan blades. It is the rotary ring 30 which is integrated in the rear fuselage portion 14, being interposed between the two sections 14-1, 14-2 and providing aerodynamic continuity between them, as is best seen. in Figure 3. This allows the blower to ingest the entire boundary layer flowing around the rear fuselage portion 14, 360 °. Thus, the boundary layer is ingested axisymmetrically with respect to the axis of the air inlet, thus avoiding a distortion of the flow at the engine inlet which could have the effect of reducing the efficiency of the fan, and to increase the risk of operability problems of this fan.
    The blower 18 is surrounded by a structural nacelle 36 mechanically connected to the front section 14-1 by radial support arms 40 spaced circumferentially from each other, and mechanically connected to the rear section 14-2 by guide vanes 42 also called OGV (Outlet Guide Vanes). Each of the two nacelles 36 may also have a steerable blower nozzle, thus controllable vertically and horizontally, to generate a vector thrust.
    In Figure 3, it is shown that the two fuselage rear portions 14 have the same length, and the two rotating rings 30 are arranged in the same transverse plane. Nevertheless, according to an alternative embodiment shown in FIG. 3a, the two rotary rings 30 could be axially offset with respect to each other, so that the bladed elements 32 of one of the motors are offset axially. bladed elements 32 of the other engine. These bladed elements 32 may therefore be partially superimposed in the axial direction. In other words, this makes it possible to bring the two fuselage rear portions 14 closer to one another, the distance Ds separating their two parallel longitudinal axes 5 then being smaller than the sum of the radius RI of the bladed elements 32 of one of the motors, and the radius R2 of the bladed elements 32 of the other engine. In this regard, it is noted that the embodiment shown in Figure 3a has a longitudinal offset for two motors spaced apart from each other in the transverse direction, in a horizontal plane. Nevertheless, this embodiment could also apply to two motors spaced from each other in the vertical direction.
    Figures 4 and 5 show that there are two empennages 50 for the assembly 1, arranged on both sides of the engines 2. A solution to a tail or a number of empennages greater than two is also possible, without depart from the scope of the invention.
    In the invention, the two empennages 50 are not necessarily vertical, but can be inclined so as to deviate from a central axis 52 of the assembly 1, going upwards. In this case, the two empennages are deemed arranged in V. Nevertheless, other arrangements may be retained as a T-arrangement shown in Figure 5b, or by providing double empennages as shown in Figure 5c . In the latter mode also called "twin-tail", on each side of the rear structure of the aircraft, there are two empennages respectively 50 substantially vertical and substantially horizontal, or slightly inclined relative to the vertical and horizontal directions.
    The two empennages 50 and the two nacelles 36 are substantially aligned transversely, being traversed by the same transversal fictitious plane PI of the assembly 1.
    For the recovery of forces from each of the two empennages 50, there is provided a beam 60 associated with each motor 2, extending generally in the direction X. At its rear end 60a, the beam connects a front end of the the empennage 50 to a structural part before the nacelle 36, which can thus ensure the transmission of forces from the empennage 50, towards the sections 14-1, 14-2 via the support arms 40 and the guide vanes of exit 42.
    In addition, the front end 60b of the beam 60 is connected to the front section 14-1 fuselage, which provides another path of effort between the empennage 50 and section 14-1. In addition, it is noted that in the rear part, the two nacelles are also mechanically connected to one another by a material ligament 64.
    It is also noted, with reference to Figure 5a, that the two pods 36 may be partially fused near the rear ends 60a. In other words, they do not each extend over 360 °, but on a lower angular sector being connected to each other at two points to form a single structure, preferably pinch shape vertically in its center.
    In the embodiment shown in FIG. 7, another benefit arises from the fact that the gas generator 16 is located in the front fuselage section 14-1 of convergent rearward shape. Indeed, this allows to tilt the gas generator, by providing longitudinal axes of generators 5 'which are no longer coincident with the longitudinal axis 5 of the fan, but inclined relative thereto.
    The two inclinations, preferably symmetrical, are such that the gas generators 16 deviate from the central axis 52 while going forward, which implies that a separation distance between the two compressor assemblies 22 is less than a separation distance between the two sets of turbines 20. In other words, the two gas generators 16 are arranged at V, symmetrically with respect to a longitudinal median plane of the assembly.
    This makes it possible for the turbine disks of the turbine assembly 20 of each engine 2 to fit into transverse turbine impeller planes P2 which do not intercept the gas generator 16 of the other engine. Thanks to this specificity, the management of the propeller blade burst risk, also known as the UERF risk, of the English "Uncontained Engine Rotor Failure", is facilitated. Indeed, it is no longer necessary to provide specific shield between the two gas generators, which advantageously reduces the overall mass of the assembly.
    FIGS. 8 and 9 illustrate the fact that the nacelle comprises thrust reverser means which are here grids 70 covered in an inactive position by one or more covers 72 movable in translation in the X direction. A rotating or other movement can also be expected. Preferably, there are provided two grids 70 respectively arranged at the top and bottom of the nacelle, ie at 12 o'clock and at 6 o'clock, so that the inverted flow of thrust does not disturb the flow of the air on empennages 50 provided laterally.
    Finally, it is noted that another preferred embodiment may consist in providing engines of the turboprop type, in which the receiver is a helix 18 'as shown diagrammatically in FIG. 10, with blades 32' that are not careened and preferably variable pitch, in particular to be able to perform the function of reverse thrust. A solution with an electric motor could also be envisaged, without departing from the scope of the invention.
    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 (15)
    [1" id="c-fr-0001]
    An assembly (1) for an aircraft (100) comprising a rear fuselage portion (10) and at least two engines (2) propelled by ingestion of the boundary layer flowing on the rear fuselage portion, each engine (2). ) comprising a receiver (18, 18 ') equipped with a rotating ring (30) from which bladed elements (32, 32') project radially outwards, characterized in that said rear fuselage portion (10 ) has a front portion (12) which splits into at least two separate rear parts (14) spaced from each other and each incorporating the rotary ring (30) of one of said motors (2).
    [2" id="c-fr-0002]
    2. An assembly according to claim 1, characterized in that each rear portion (14) comprises successively, from front to rear: - a front fuselage section (14-1) of convergent shape rearwardly; the rotary ring (30); and - a rear section (14-2).
    [3" id="c-fr-0003]
    3. An assembly according to claim 1 or claim 2, characterized in that the receivers (18, 18 ') of said motors (2) are spaced from each other in a transverse direction (Y) and / or in a direction of the height (Z) of said assembly, and optionally, the two receivers (18, 18 ') of two motors are spaced from one another in a longitudinal direction (X), so that a distance (Ds) between the two longitudinal parallel axes (5) of the two distinct rear parts (14) carrying respectively the two receivers (18, 18 '), or less than the sum of a radius (RI) of the bladed elements (32, 32') of the one of the two receivers and a radius (R2) of the bladed elements (32, 32 ') of the other receiver.
    [4" id="c-fr-0004]
    4. An assembly according to any one of claims 1 to 3, characterized in that each motor (2) comprises a gas generator (16) driving the receiver (18, 18 ') of said engine, said gas generator comprising a set compressor (22), a combustion chamber (24) and a turbine assembly (20), said gas generator (16) being preferentially arranged in front of the receiver (18, 18 ').
    [5" id="c-fr-0005]
    5. An assembly according to the preceding claim, characterized in that each motor (2) has an inverted design in which the turbine assembly (20) is located in front of the compressor assembly (22), the gas evacuation ports exhaust (28) through the rear fuselage portion being preferentially arranged in front of the gas generator (16).
    [6" id="c-fr-0006]
    6. An assembly according to the preceding claim, characterized in that for two motors (2) facing one another, the respective longitudinal axes (5 ') of the two gas generators (16) are inclined with respect to a longitudinal direction (X) of the assembly, so that a separation distance between the two sets of compressors (22) is less than a separation distance between the two sets of turbines (20).
    [7" id="c-fr-0007]
    7. Assembly according to the preceding claim, characterized in that the respective longitudinal axes (5 ') of the two gas generators (16) are inclined relative to the longitudinal direction (X) of the assembly, so that the disks of turbine of the turbine assembly (20) of one of said engines (2) fit into transverse turbine planes (P2) which do not intercept the gas generator (16) of the other of said engines, and reciprocally.
    [8" id="c-fr-0008]
    8. An assembly according to any one of the preceding claims, characterized in that each receiver is a non-ducted propeller (18 '), the blades (32') are preferably variable pitch.
    [9" id="c-fr-0009]
    9. An assembly according to any one of claims 1 to 7, characterized in that each receiver is a blower (18) surrounded by a nacelle (36) connected to the rear fuselage associated portion (10) via arms front support (40) and / or exit guide vanes (42).
    [10" id="c-fr-0010]
    10. An assembly according to the preceding claim, characterized in that each nacelle (36) is structural and configured to ensure a transmission of forces from one or more empennages (50) of all, towards the rear portion associated fuselage (14), and / or in that the assembly comprises at least one beam (60) for the passage of forces between a stabilizer (50) and the rear fuselage portion (14).
    [11" id="c-fr-0011]
    11. An assembly according to claim 9 or claim 10, characterized in that the nacelles (36) are mechanically connected to each other.
    [12" id="c-fr-0012]
    12. An assembly according to any one of claims 9 to 11, characterized in that each nacelle (36) comprises thrust reverser means, preferably comprising thrust reversal grids (70) covered by one or more movable hoods (72).
    [13" id="c-fr-0013]
    13. Assembly according to any one of the preceding claims, characterized in that it comprises two empennages (50).
    [14" id="c-fr-0014]
    14. An assembly according to the preceding claim combined with any one of claims 9 to 12, characterized in that the two empennages (50) and nacelles (36) are traversed by a same transverse fictitious plane (PI) of all .
    [15" id="c-fr-0015]
    15. Aircraft (100) comprising an assembly (1) according to any one of the preceding claims.
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    法律状态:
    2017-06-21| PLFP| Fee payment|Year of fee payment: 2 |
    2017-12-22| PLSC| Search report ready|Effective date: 20171222 |
    2018-06-26| PLFP| Fee payment|Year of fee payment: 3 |
    2020-06-19| PLFP| Fee payment|Year of fee payment: 5 |
    2021-06-22| PLFP| Fee payment|Year of fee payment: 6 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1655719|2016-06-20|
    FR1655719A|FR3052743B1|2016-06-20|2016-06-20|AIRCRAFT ASSEMBLY COMPRISING PROPULSION ENGINES BY INGESTION OF THE LIMIT LAYER|FR1655719A| FR3052743B1|2016-06-20|2016-06-20|AIRCRAFT ASSEMBLY COMPRISING PROPULSION ENGINES BY INGESTION OF THE LIMIT LAYER|
    US15/626,698| US10633101B2|2016-06-20|2017-06-19|Assembly for aircraft comprising engines with boundary layer ingestion propulsion|
    CN201710469385.XA| CN107521705B|2016-06-20|2017-06-20|Assembly for an aircraft comprising an engine with boundary layer suction propulsion|
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