Assembly for aircraft comprising engines with boundary layer ingestion propulsion

In order to further benefit from the principle of boundary layer ingestion by engines of an aircraft assembly, the rear portion of the fuselage of this aircraft assembly includes a front portion which splits up into at least two distinct rear portions, spaced apart from each other, and each integrating the rotary ring of the receiver of one of the engines.

CROSS-REFERENCES TO RELATED APPLICATIONS

This application claims the benefit of the French patent application No. 1655719 filed on Jun. 20, 2016, the entire disclosures of which are incorporated herein by way of reference.

BACKGROUND OF THE INVENTION

The present invention pertains to the field of aircraft comprising a rear portion of the fuselage equipped with engines with boundary layer ingestion propulsion. As is known, boundary layer ingestion propulsion corresponds to the engines taking in an air flow with weak kinetic energy, circulating around the rear portion of fuselage. This technique reduces the kinetic energy expended for the propulsion as well as the drag of the aircraft, with the result of less fuel consumption.

It is known how to attach, in the rear portion of the fuselage, engines with boundary layer ingestion propulsion. For example, this involves two half-sunken engines placed side by side, protruding upward or to the side from the rear portion of fuselage.

However, in this type of configuration, the two engines are only able to take in one portion of the boundary layer of air circulating on the rear portion of fuselage. For these configurations, the boundary layer is also taken in without axial symmetry in relation to the air inlet axis, thus generating a distortion of the incoming flow of the engine.

Thus, there is a need for optimization in order to better profit from the principle of boundary layer ingestion propulsion.

SUMMARY OF THE INVENTION

In order to meet this need at least in part, the invention concerns an assembly for aircraft comprising a rear portion of fuselage as well as at least two engines with propulsion by boundary layer ingestion circulating on the rear portion of fuselage, each engine comprising a receiver equipped with a rotary ring from which bladed elements project radially to the outside. According to the invention, the rear portion of fuselage includes a front portion which splits up into at least two distinct rear portions spaced apart from each other and each integrating the rotary ring of one of the engines.

The invention thus calls for a separation of the fuselage into several rear portions, each of which is associated with an engine, such that its receiver can take in all of the boundary layer circulating on its associated rear portion. This results advantageously in better overall performance of the aircraft.

The invention also calls for the implementation of the following optional characteristics, taken by themselves or in combination.

Each rear portion comprises in succession, from front to rear:a front piece of fuselage with shape converging toward the rear;the rotary ring; anda rear piece.

The receivers of the engines are spaced apart from each other along a transverse direction and/or along a height direction of the assembly. Optionally, the two receivers of the two engines are spaced apart from each other along a longitudinal direction, such that a distance between the two parallel longitudinal axes of the two distinct rear portions respectively bearing the two receivers, is less than the sum of 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 the engine, the gas generator comprising a compressor assembly, a combustion chamber and a turbine assembly, the gas generator being preferably disposed in front of the receiver. Alternatively, the gas generator could be placed behind the receiver.

By another possibility, each receiver may also be driven by an electric motor.

Each engine preferably has an inverted design in which the turbine assembly is situated in front of the compressor assembly, orifices for evacuation of exhaust gases through the rear portion of fuselage being preferably disposed in front of the gas generator.

For two engines facing each other, the respective longitudinal axes of the two gas generators are inclined with respect to a longitudinal direction of the assembly, such that a distance separating the two compressor assemblies is less than a distance separating the two turbine assemblies.

The respective longitudinal axes of the two gas generators are inclined with respect to the longitudinal direction of the assembly, such that the turbine disks of the turbine assembly of one of the engines are inscribed in imaginary transverse turbine planes which do not intercept the gas generator of the other of the engines, and vice versa.

Each receiver is an uncased propeller whose blades preferably have a variable timing.

Alternatively, each receiver is a fan surrounded by a nacelle joined to the associated rear portion of fuselage, by means of front support arms and/or outlet guide vanes.

Each nacelle is structural and designed in order to ensure a transmission of the forces coming from one or more empennages of the assembly, in the direction of the associated rear portion of fuselage, and/or the assembly comprises at least one force transfer beam between an empennage and the rear portion of fuselage.

The nacelles are mechanically joined to each other.

Each nacelle comprises thrust reversal means, preferably comprising thrust reversal grids covered by one or more mobile cowls.

The assembly comprises two empennages.

Preferably, the two empennages and the nacelles have the same imaginary transverse plane of the assembly passing through them.

Finally, the invention also concerns an aircraft comprising an assembly such as the one described above, the aircraft being preferably of commercial type.

Other advantages and characteristics of the invention will appear in the following detailed and nonlimiting description.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Referring first of all toFIG. 1, there is represented an aircraft100of the commercial type, comprising an assembly1corresponding to its rear end, provided with engines2. On this aircraft, the wings4are not equipped with engines, even though this could be the case, without departing from the scope of the invention. The engines2are only disposed on the assembly1, one preferred embodiment of which shall now be described with reference toFIGS. 2 to 6. In these figures, the terms “front” and “rear” should be considered in regard to a direction of movement8of the aircraft, due to the thrust generated by the engines2.

The assembly1comprises a rear portion of fuselage10, forming the rear end of the fuselage of the aircraft. This portion10comprises a front portion12whose one front end12ahas a fuselage shape, which may be, for example, oval, circular, or the like.

Moving in the direction of its rear end12b, the front portion12is progressively pinched at its center until it splits into two distinct rear portions of fuselage, referenced14. The two rear portions14, preferably of identical solid of revolution shape, are spaced apart from each other in a transverse direction Y of the assembly. In this context, it is noted that, by convention, the X direction corresponds to the longitudinal direction of the assembly1, which is likewise equated to the longitudinal direction of each engine of this assembly1. This direction X is parallel to a longitudinal axis5of each engine2. On the other hand, the direction Y corresponds to the direction oriented transversely with respect to the assembly1and likewise equated to the transverse direction of each engine, while the direction Z corresponds to the vertical or height direction. These three directions X, Y and Z are orthogonal to each other and form a direct trihedron.

Each rear portion14is designed to integrate all or some of one of the engines2. Consequently, in the preferred embodiment which calls for two engines spaced apart along the direction Y, two rear portions14are provided. In a different case where a third engine will be added, spaced apart from the first two along each of the directions Y and Z so as to make a triangle arrangement, there will then be provided three rear portions of fuselage. In the case of four engines, these may be arranged in a square or a rectangle, being integrated respectively in four rear portions of fuselage14.

In the preferred embodiment represented inFIGS. 2 to 6, there are thus provided two rear portions of fuselage14spaced apart from each other along the direction Y, and running in parallel along the direction X, starting from the rear end12bof the portion of fuselage12. Starting from this rear end12b, each rear portion14has, first of all, a front piece14-1which narrows, for example in a truncated conical or similar shape, converging toward the rear. This portion14then integrates an element of its associated engine2, such as will be explained below, and then terminates toward the rear with a rear piece14-2of circular cross section and substantially constant diameter, of bullet shape, converging shape, or more complex shape.

In this preferred embodiment, each rear portion of fuselage14is centered on the longitudinal axis5of its associated engine2. Each engine here is of the turbojet type with propulsion by boundary layer ingestion circulating on the corresponding rear portion of fuselage14. Referring more precisely toFIG. 4, each engine2is thus equipped with a gas generator16driving a receiver18. The generator16is disposed in front of the receiver18, which allows it to be integrated in whole or in part inside the converging front piece14-1of the rear portion of the fuselage14. This avoids the presence of a substantial weight at the rear end of the aircraft, and facilitates the balancing of the latter while reducing the balancing drag.

The gas generator16has a so-called inverted design, in which a turbine assembly20is arranged in front of a compressor assembly22, with a combustion chamber24located between the two. As is shown schematically inFIG. 6, this makes it possible to arrange orifices28for evacuation of the hot gases coming from the turbine assembly20, upstream from the generator. These orifices28pass through the front fuselage piece14-1, thus assuming a forward position, which offers several advantages.

First of all, as for the hot gases being ejected far upstream by the orifices28, their cooling is promoted by mixing with the ambient air for a substantial length, before any impacting of these gases on rear portions of the aircraft.

Moreover, this far upstream, disposition of the orifices28simplifies the use of energy recuperation systems, which has the benefit of increasing the efficiency of the engine and generating power for the aircraft cabin.

The receiver here is a turbine fan, comprising a rotary ring30, also known as a fan hub, from which bladed elements32project, also known as fan blades. It is the rotary ring30which is integrated in the rear part of the fuselage14, being interposed between the two pieces14-1,14-2and ensuring an aerodynamic continuity between them, as is best seen inFIG. 3. This allows the fan to take in all of the boundary layer circulating about the rear portion of fuselage14, over 360°. Thus, the boundary layer is taken in with axial symmetry relative to the air inlet axis, thus preventing any distortion of the incoming flow of the engine which might have the consequence of reducing the efficiency of the fan, and of increasing the risk of operational problems with this fan.

The fan18is surrounded by a structural nacelle36joined mechanically to the front piece14-1by radial support arms40spaced apart from each other circumferentially, and joined mechanically to the rear piece14-2by outlet guide vanes42, or OGV. Each of the two nacelles36may likewise have an orientable fan nozzle, i.e., one which can be piloted vertically and horizontally, to generate a vectorial thrust.

InFIG. 3, it is shown that the two rear portions of fuselage14have the same length, and the two rotary rings30are arranged in a same transverse plane. Even so, in one alternative embodiment shown inFIG. 3a, the two rotary rings30could be axially staggered with respect to each other, so that the bladed elements32of one of the engines are axially offset from the bladed elements32of the other engine. These bladed elements32may thus be partially superimposed in the axial direction. In other words, this makes it possible to bring closer together the two rear portions of fuselage14, the distance Ds separating their two parallel longitudinal axes5then being able to be less than the sum of the radius R1of the bladed elements32of one of the engines, and the radius R2of the bladed elements32of the other engine.

In this context, it is noted that the embodiment shown inFIG. 3ahas a longitudinal offset for two engines spaced apart from each other in the transverse direction, in a horizontal plane. Even so, this embodiment might likewise be applied to two engines spaced apart from each other in the vertical direction.

FIGS. 4 and 5show that there are provided two empennages50for the assembly1, arranged on either side of the engines2. A solution with one empennage or a number of empennages greater than two may likewise be contemplated, without departing from the scope of the invention.

In the invention, the two empennages50are not necessarily vertical, but may be inclined so as to depart from a central axis52of the assembly1, running upward. In this case, the two empennages are said to have a V-shape. However, other dispositions can be used, such as a T shape disposition, shown inFIG. 5b, or by providing double empennages as shown inFIG. 5c. In this latter embodiment, also called a “twin-tail,” on either side of the rear structure of the aircraft, there are provided two empennages50, respectively substantially vertical and substantially horizontal, or instead slightly inclined with respect to the vertical and horizontal directions. Finally, another possibility represented inFIG. 5dcomprises attaching conventional empennages50to the nacelles, in whole or in part. For example, the central empennage is likewise attached partly to the fuselage, to the front of the nacelles.

The two empennages50and the two nacelles36are substantially aligned transversely, having the same imaginary transverse plane P1of the assembly1passing through them.

In order to once again take up forces coming from each of the two empennages50, there is provided a beam60associated with each engine2, extending generally in the direction X. In the area of its rear end60a, the beam joins a front end of the empennage50to a front structural part of the nacelle36, which can thus ensure the transmission of the forces coming from the empennage50, in the direction of the pieces14-1,14-2via the support arms40and the outlet guide vanes42.

Moreover, the front end60bof the beam60is connected to the front piece14-1of fuselage, which makes it possible to provide a different path for forces between the empennage50and this piece14-1. Moreover, it is noted that, in a rear portion, the two nacelles are likewise joined mechanically to each other by a material ligament64.

It is likewise noted, referring toFIG. 5a, that the two nacelles36may be partly merged near the rear ends60a. In other words, they do not each extend for 360°, but rather are joined to each other at two points on a lesser angular sector so as to form only a single structure, preferably with a shape pinched vertically at its center.

In the embodiment shown inFIG. 7, another benefit comes from the fact that the gas generator16is situated in the front piece of fuselage14-1with shape converging toward the rear. In fact, this makes it possible to tilt the gas generator, providing longitudinal axes of generators5′ which are no longer merged with the longitudinal axis5of the fan, but rather inclined relative to them.

The two inclinations, preferably being symmetrical, are such that the gas generators16make away from the central axis52in the forward direction, which means that a separation distance between the two compressor assemblies22is less than a separation distance between the two turbine assemblies20. In other words, the two gas generators16are arranged in a V, symmetrically with respect to a median longitudinal plane of the assembly.

This makes it possible to have the turbine disks of the turbine assembly20of each engine2inscribed in the imaginary transverse turbine planes P2not intercepting the gas generator16of the other engine. Thanks to this feature, the managing of the risk of break-up of the propeller blades, also known as the UERF or “Uncontained Engine Rotor Failure” risk, is facilitated. In fact, it is no longer necessary to provide a specific shield between the two gas generators, which advantageously allows a reduction in the overall weight of the assembly.

FIGS. 8 and 9illustrate the fact that the nacelle comprises means of thrust reversal, which are the grids70here, covered in the inactive position by one or more cowls72able to move in translation along the direction X. A movement by rotation or another movement may also be provided. Preferably, there are provided two grids70, arranged respectively at the top and bottom of the nacelle, or at the 12 o'clock and 6 o'clock positions, so that the reversed thrust flow will not perturb the air flow provided laterally on the empennages50.

Finally, it is noted that another preferred embodiment might comprise having turboprop type engines, in which the receiver is a propeller18′ as shown schematically inFIG. 10, with uncased blades32′ and preferably having variable timing, in particular to provide a thrust reversal function. A solution with electric motor could likewise be contemplated to drive the receivers, without departing from the scope of the invention.

Of course, various modifications may be provided by the person skilled in the art for the invention that has just been described, solely by way of nonlimiting examples. In particular, the embodiments which have been described above are not mutually exclusive, but rather may be combined with each other.