FR2994941A1 - Propelling assembly for aircraft, has external aerodynamic fairing part fixed on rigid substructure, and connection unit formed such that part of connection effort of turbomotor with rigid fixing structure passes through structural layer - Google Patents

Abstract

The assembly has an external aerodynamic fairing part fixed on a rigid substructure (28) to partly cover the rigid substructure. A structural layer of the aerodynamic fairing part exhibits a concave form along direction of a longitudinal axis (17) of a turbomotor when a cross-section of the structural layer is viewed along a transverse plan (P) that is orthogonal to the longitudinal axis of the turbomotor. A connection unit is formed such that a part of connection effort of the turbomotor with a rigid fixing structure (18) passes through the structural layer.

Description

The present invention relates to the field of propulsion systems for aircraft, and more particularly relates to the means for attaching turbine engines to the structure of aircraft. BACKGROUND OF THE INVENTION STATE OF THE PRIOR ART In aircraft, in particular airplanes, conventional turbojet engines as well as turbojet engines with non-ducted fan (s), also called "open rotor", are generally supported by an attachment pylon, also called pylon or EMS, of the English "Engine Mbunting Structure", comprising a rigid structure generally box-shaped bearing means for connecting the turbojet engine to this rigid structure. These connecting means are usually divided into a front engine attachment and a rear engine attachment generally centered with respect to a median plane of the rigid structure, and two rods for taking up the thrust forces arranged on either side of this plane. median. Hanging poles of this type are illustrated in FIG. 1 of the international patent application No. WO2009 / 147341 A2 with regard to the case of the attachment of a conventional turbojet engine under a main wing of an airplane and on Figures 1 and 2 of the international patent application WO2010 / 031959 A1 with respect to the case of an unducted fanless turbojet engine attached to the side of a rear part of the fuselage of an aircraft.

However, these attachment masts do not allow an optimal distribution of the connecting means of the turbine engines to the rigid structure of these masts. These connecting means are indeed concentrated at the distal end of the rigid structure, that is to say its end facing the turbine engine. The recovery of efforts by these means of connection can not be optimal. In addition, this type of configuration limits the possibilities of integrating mechanical energy dissipating elements to limit the transmission of vibrations from the turbine engine to the structure of the aircraft. This type of configuration also considerably complicates the integration of means for controlling the inclination of the turbine engine, in particular upwards or downwards. In addition, the engine fasteners are, in general, arranged to the right of bladed wheels in rotation of the turbine engine. In order to overcome any accidental tearing of one or more blades of such a bladed wheel, a phenomenon also called UERF of the English Uncontained Engine Rotor Failure, it is thus necessary to provide redundancy in these engine fasteners, by adding unsolicited engine fasteners in normal operation, known as "engine attachments pending", and / or by allowing, in the event of failure of one or more engine attachments, that the other engine attachments are capable of taking up all of the support efforts of the turbine engine. Moreover, the engine fasteners disposed at the distal end of the rigid structure require that the attachment and detachment of the turbine engine is effected by means of a translation of the turbine engine substantially in a direction included in a median plane of the engine. rigid structure.

In the case of a turbine engine attached to the side of the fuselage of an airplane, this median plane of the rigid structure is inclined relative to the vertical direction, so that a complete procedure for mounting or dismounting such a turbine engine comprises at least two successive translations, a translation in the vertical direction and a translation along a direction of the median plane of the rigid structure. It is thus desirable to be able to simplify the assembly and disassembly procedures of the turboshaft engines. DISCLOSURE OF THE INVENTION The object of the invention is in particular to provide a simple, economical and effective solution to these problems, making it possible to avoid at least in part the aforementioned drawbacks. To this end, the invention proposes a propulsion unit for an aircraft, comprising a turbine engine, an external aerodynamic fairing surrounding at least one part and preferably all of this turbine engine, and a device for attaching the turbine engine comprising a rigid attachment structure which has a proximal end intended to be connected to an aircraft structure and a distal end arranged facing the turbine engine, said rigid attachment structure carrying connecting means which connect it to the turbine engine. According to the invention: - said connecting means comprise a rigid substructure surrounding at least a portion 10 of said turbine engine and connected to said rigid attachment structure; said external aerodynamic fairing comprises at least one structural skin connected to said rigid attachment structure, and fixed on said rigid substructure so as to cover at least part of this rigid substructure; when said structural skin of the external aerodynamic fairing is seen in section along a transverse plane orthogonal to a longitudinal axis of said turbine engine, said structural skin has a concave shape in the direction of said longitudinal axis; said connecting means are shaped so that a part of the forces of connection of the turbine engine to said rigid attachment structure 25 passes through said structural skin of said outer aerodynamic fairing. The bonding of the structural skin of the outer aerodynamic fairing to the rigid attachment structure may be a direct connection, in which case said structural skin is directly attached to the rigid attachment structure, or an indirect connection via other elements, as will become clearer in the following. By "external aerodynamic fairing", it is necessary to understand a fairing bathed by the relative wind circulating outside the propulsion unit. By "connecting means" connecting the turbine engine to the rigid attachment structure, it is necessary to understand means capable of transmitting to the rigid attachment structure the forces resulting from the support of the turbine engine and the forces resulting from the thrust generated by the turbine engine. The aircraft structure concerned may be in particular the fuselage or a wing of the aircraft. In addition, throughout the present description, the "proximal" or "distal" character of an element respectively reflects the proximity or distance of this element from the end of the rigid attachment structure intended to be connected to an aircraft structure. Thus, when the propulsion unit equips an aircraft, the "proximal" or "distal" character respectively reflects the proximity or distance of the element relative to said aircraft structure. The longitudinal direction of the turbine engine 25 or the propulsion unit is defined as being the direction of the aforementioned longitudinal axis of this turbine engine. The invention allows an optimal distribution of the connecting forces of the turbine engine to the aforementioned rigid attachment structure, by passing some of these forces in the structural skin of the external aerodynamic fairing. The invention also makes it possible to lighten the structural elements of the aforementioned fastening device, in particular with regard to the rigid substructure, because part of the binding forces of the turbine engine to the rigid structure of attachment. passes through the structural skin of the external aerodynamic fairing.

In addition, the structural skin of the external aerodynamic fairing provides an extended surface on which can be arranged motor fasteners and / or mechanical energy dissipating elements, also called dampers, intended to connect the turbine engine to the rigid structure of attachment. These engine fasteners and / or mechanical energy dissipating elements can thus be distributed optimally with respect to the turbine engine. It should be noted that the structural skin of the external aerodynamic fairing can be made in one piece or be formed of an assembly of several adjacent panels. In addition, this structural skin of the external aerodynamic fairing is preferably made of a composite material comprising reinforcing fibers embedded in a cured resin. As a variant, this structural skin of the external aerodynamic fairing may be made of a metallic material. Furthermore, when the structural skin 30 of the external aerodynamic fairing is seen in section along a transverse plane, this structural skin advantageously has a circumferential extent greater than 90 degrees, preferably greater than 120 degrees, and even more preferably greater than 180 degrees, around of the longitudinal axis of the turbine engine, and this structural skin advantageously has a circumferential extent less than 230 degrees. Such a circumferential extent allows effective transmission of forces by the structural skin of the outer aerodynamic fairing while providing facilities in terms of attachment and removal of the turbine engine, as will become more clearly apparent in the following. The outer aerodynamic fairing further advantageously comprises a closure cap articulated on a circumferential end of the structural skin of this external aerodynamic fairing. Said hood may be designed to participate also in the transmission of connecting forces of the turbine engine to the rigid structure of attachment, when the hood is in a closed position. Furthermore, the structural skin of the outer aerodynamic fairing, and said connecting means connecting the turbine engine to the rigid attachment structure, are preferably integrally included in a region of space such that: said region of the space is bounded by two half-planes; each of these two half-planes is delimited by said longitudinal axis; these two half-planes extend symmetrically on either side of a longitudinal plane including said longitudinal axis; each of said half-planes forms an angle less than 120 degrees with said longitudinal plane; said longitudinal plane is intended to extend vertically and said region of the space is intended to extend above said half-planes, when said propulsion unit equips an aircraft parked on the ground. The structural skin of the outer aerodynamic fairing and the said connecting means thus allow the turbine engine to move in translation vertically from or to its operating position, respectively during the removal or for the installation of this turbine engine in its service position. as will become clearer in the following. In a preferred embodiment of the invention, the rigid attachment structure extends laterally with respect to the turbine engine so that the proximal end of this rigid attachment structure can be laterally attached to a structural structure. aircraft. The propulsion unit can thus be attached laterally to an aircraft fuselage, in particular to a rear part of such a fuselage. The method of fixing the rigid structure 30 of attachment to the aircraft structure is not the subject of the invention. This may be a conventional method of attachment, for example by fixing the proximal end of the rigid attachment structure to a box structure disposed within the fuselage, or alternatively, by direct connection, within of the fuselage, the proximal end of the rigid attachment structure at a proximal end of a rigid structure of attachment of another propulsion unit disposed on the other side of the fuselage, as shown in Figure 2 of the International Patent Application No. WO2010 / 031959 A1 cited above. In a variant, the propulsion unit may be configured for attachment under the wings of an aircraft, or for another mode of attachment, without departing from the scope of the present invention.

In the preferred embodiment of the invention, the propulsion unit comprises an upper aerodynamic skin which covers at least in part said rigid attachment structure and which is connected to a proximal end portion of the structural skin of the aerodynamic fairing. external. The upper aerodynamic skin is advantageously made in one piece with the structural skin of the external aerodynamic fairing. Alternatively, the upper aerodynamic skin can be attached to the structural skin of the external aerodynamic fairing, preferably by splinting. In both cases, the attachment device is preferably designed so that a part of the forces of connection of the turbine engine to said rigid attachment structure passes through said upper aerodynamic skin, the latter being designed to be able to transmit this part of the connection forces of the turbine engine to the rigid structure of attachment.

The upper aerodynamic skin thus contributes to the transmission of the connecting forces of the turbine engine to the rigid attachment structure, and thereby further reduces the mass of the aforementioned connecting means. The upper aerodynamic skin also makes it possible in this case to contribute to the rigidity of the rigid attachment structure, which may make it possible to lighten this rigid attachment structure. Furthermore, the rigid attachment structure is advantageously formed of an assembly of longitudinal members and ribs forming a box structure. Alternatively, this rigid attachment structure may comprise a honeycomb-like cellular structure, also called NIDA, or be formed of a solid foam. In one embodiment of the invention, the upper aerodynamic skin is integrally formed with the structural skin of the outer aerodynamic fairing, and said proximal end portion of the structural skin of the outer aerodynamic fairing has indentations therethrough. of which respective end portions of longitudinal members of the rigid attachment structure pass respectively, these notches being separated by projecting parts fixed on a lower aerodynamic skin, the latter being itself fixed on a lower side of the rigid structure hooking. Fixing the protruding parts on the lower aerodynamic skin is preferably made by splicing. In the preferred embodiment of the invention, the structural skin of the external aerodynamic fairing comprises at least one inspection opening 10 provided with a cover for access to the turbine engine. Furthermore, said connecting means advantageously comprise at least one yoke intended for the attachment of an engine attachment and fixed on the rigid substructure by removable fastening means, such as bolts. Such a clevis can easily be attached to the rigid substructure when the turbine engine is connected to the coupling device or, conversely, detached from this rigid substructure when this turbine engine is removed. These operations can thus be performed without it being necessary to assemble or disassemble the axis connecting the motor attachment concerned to the aforementioned screed. In addition, the propulsion unit advantageously comprises at least one reinforcement panel disposed between the turbine engine and said rigid substructure and fixed thereto. This reinforcing panel preferably has a proximal circumferential end connected to the distal end of the rigid attachment structure, and a distal circumferential end forming with its proximal circumferential end an angle of between 70 degrees and 110 degrees around the longitudinal axis of the turbine engine. In the preferred embodiment of the invention, for each access opening of said structural skin of the external aerodynamic fairing arranged next to said reinforcement panel, the latter comprises a corresponding inspection opening arranged opposite said inspection opening of the structural skin of the aerodynamic fairing, so as to allow access to the turbine engine. The invention also relates to an aircraft, such as an airplane, comprising at least one propulsion unit of the type described above.

The invention finally relates to a method for attaching or removing a turbine engine of a propulsion unit of an aircraft of the type described above, comprising a displacement of the turbine engine, between a service position in which the turbine engine is hung. to the attachment device of said propulsion unit, and a "deposited" position in which the turbine engine is placed on the ground. According to the invention, said displacement is a translation in the vertical direction.

The removal of the turbine engine from its operating position, and conversely, its connection to the attachment device from the ground, can thus be operated by simply moving the turbine engine in translation in the vertical direction. BRIEF DESCRIPTION OF THE DRAWINGS The invention will be better understood, and other details, advantages and characteristics thereof will appear on reading the following description given by way of non-limiting example and with reference to the appended drawings, in which: FIG. a schematic cross-sectional view of a rear portion of an aircraft equipped with two propulsion units according to a preferred embodiment of the invention, these propulsion units being reported laterally on the fuselage of the aircraft; FIG. 2a is a schematic perspective plunging view of one of the propulsion units of the aircraft of FIG. 1, this propulsion unit comprising a turbine engine, an attachment device for this turbine engine, and an external aerodynamic fairing; Figure 2b is a diagrammatic perspective view, and against the dive, of the propulsion unit of Figure 2a; Figure 3 is a diagrammatic perspective view of a rigid attachment structure and a rigid substructure which belong to the attachment device of the propulsion unit of Figure 2a; Figure 4a is a schematic cross-sectional view along the plane P of Figure 3 of a front bow belonging to the rigid substructure of Figure 3; Figure 4b is a schematic perspective view of an additional hoop belonging to the rigid substructure of Figure 3; Figure 5 is a partial schematic perspective view, and downwardly, of the attachment device of the propulsion unit of Figure 2a, coated with a structural skin external aerodynamic fairing of the propulsion unit; FIG. 6 is a partial schematic view from above of the attachment device of the propulsion unit of FIG. 2a, coated with the structural skin of said external aerodynamic fairing; FIGS. 7 and 8 are partial schematic perspective views of the propulsion assembly of FIG. 2a seen from the front of the aircraft; FIGS. 9 and 10 are partial diagrammatic views in perspective and on a larger scale of the propulsion unit of FIG. 2a, respectively illustrating an upper engine attachment before taking up vertical forces and a lateral engine fastening for taking up transverse forces. ; Figure 11 is a partial diagrammatic view from the rear of the aircraft of the propulsion assembly of Figure 2a; - Figure 12 is a partial schematic view on a larger scale of the detail XII of Figure 11, illustrating a rear lateral engine attachment of recovery of vertical forces; FIG. 13 is a partial diagrammatic view from the front of the aircraft of the propulsion unit of FIG. 2a, illustrating a method of removing the turbine engine from this propulsion unit; - Figure 14 is a partial schematic perspective view of a propulsion unit according to another embodiment of the invention, seen from the front of the aircraft, illustrating more particularly an upper engine attachment front of the propulsion unit; Figure 15 is a partial schematic cross-sectional view of a propulsion assembly according to another embodiment of the invention; FIG. 15a is an enlarged view in longitudinal section along the plane A-A of FIG. 15 of a lateral engine attachment of the propulsion unit of this FIG. 15; - Figure 16 is a partial schematic perspective view of a rear upper engine attachment of the propulsion unit of Figure 15; FIG. 17 is a partial diagrammatic cross-sectional view of a propulsion assembly according to another embodiment of the invention, illustrating the connection zone of the rigid substructure with the rigid structure of attachment of the device 25; attachment of this propulsion unit; - Figure 18 is a view similar to Figure 17, illustrating an alternative embodiment of the propulsion system; FIG. 19 is a view similar to FIG. 17 of the propulsion unit of this FIG. 17 viewed along a plane of section offset longitudinally with respect to the plane of FIG. 17. In all of these figures, identical references may designate identical or similar elements. DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS FIG. 1 very schematically illustrates a rear portion of an aircraft 8, such as an aircraft, seen in cross-section, comprising two propulsion assemblies 10 reported laterally on the fuselage 11 of the aircraft. These propulsion units 10 are thus arranged behind the main wings of the aircraft (not visible in Figure 1). Each propulsion assembly 10 generally comprises a turbine engine 12 (FIG. 2b), a device 14 for attaching the turbine engine to an aircraft structure (FIGS. 2a and 2b), and an external aerodynamic fairing 16 for guiding the flow. of outside air, also called relative wind, around the turbine engine 12. In the illustrated example, the turbine engine 12 is a turbofan engine with unsheathed contra-rotating propellers located in a rear part of the turbojet engine, these propellers not being represented on the engines. FIGS. This type of turbojet is commonly called "open rotor pusher". Throughout the following description, by convention, X is the longitudinal direction of the turbine engine 12 and more generally of the propulsion unit 10 and the aircraft 8, this longitudinal direction X being parallel to a longitudinal axis 17 of the turbine engine 12 On the other hand, Z is the direction of the height, which corresponds to the vertical direction when the aircraft 8 is parked on the ground or is in a zero roll position in flight, and Y is referred to as the transverse direction of the aircraft. propulsion assembly 10. The three directions X, Y and Z are orthogonal to each other. Furthermore, the terms "forward" and "backward" are to be considered with respect to a direction of advancement of the aircraft 8, represented schematically by the arrow F. In addition, the terms "distal" and "proximal" translate respectively the fact that an element is distant or close to the fuselage of the aircraft. The hooking device 14 comprises a rigid attachment structure 18, as well as a leading edge fairing 20 and a trailing edge fairing 22 which are respectively attached to the front and rear ends of the rigid structure. 18. The rigid attachment structure 18 has a proximal end 24 intended to be connected laterally to the fuselage 11 of the aircraft 8, 25 and an opposite distal end 26 which extends opposite the turbine engine 12. hooking 14 further comprises a rigid substructure 28 connected to the distal end 26 of the rigid structure 30 of attachment 18.

The rigid attachment structure 18 comprises (FIG. 4) a front spar 52, a median spar 54, and a rear spar 56, as well as three internal ribs 58 and a closing rib 60. The rigid attachment structure 18 thus takes the shape of a box. In a manner known per se, each of the aforementioned longitudinal members has, for example, the form of a beam with I-section, and thus comprises a core 61a extending substantially in a plane orthogonal to the longitudinal axis 17 of the turbine engine , a lower sole 61b and an upper sole 61c respectively connected to the lower and upper edges of the core 61a and extending substantially in planes parallel to the longitudinal axis 17 of the turbine engine and orthogonal to the core 61a. The rigid substructure 28 comprises two substantially identical arches, namely a front bow 62 and a rear bow 64. These two arches have respective proximal ends 66a connected to respective distal ends 68 of the front and rear longitudinal members 56 and 56 respectively. that respective distal ends 66b opposite. These two arches 62 and 64 are curved elements which each have a concavity oriented towards the longitudinal axis 17 of the turbine engine, as can be seen in FIG. 4a. The front bow 62 continues beyond a longitudinal plane V (FIGS. 3, 4a) including the longitudinal axis 17 and intended to extend vertically when the aircraft 8 is parked on the ground. The circumferential extent of the front bow 62 is quantified by means of an angle O1 formed by two half-lines D1 and D2 which each extend from the longitudinal axis 17. The first half-line D1 passes through The distal end 66b of the front bow 62, while the second half-line D2 is in the longitudinal plane V and intercepts the front bow 62. The angle 01 of the angular sector delimited by the two half-lines D1 and D2 is substantially equal to 100 degrees. Preferably, this angle θ is between 45 degrees and 120 degrees. According to one variant, the circumferential extent of the front bow 62 is quantified by considering that a radially inner edge of the bow, seen in section along the transverse plane P of FIG. 3, extends in the direction of the distal end 66b of the arch from an upper edge 70 of the closing rib 60 of the rigid fastening structure 18. The radially inner edge thus defined of the front bow 62 has a circumferential extent equal to About 200 degrees around the longitudinal axis 17 of the turbine engine. The properties described above with regard to the front bow 62 are of course valid for the rear bow 64. The rigid substructure 28 further comprises a longitudinal beam 72 (FIG. 3) centered with respect to the longitudinal plane. V. This longitudinal beam 72 is embedded in a median portion 30 of each of the front bow 62 and rear 64.

The rigid substructure 28 also comprises a medial half-arch 74 having a proximal end 76 connected to a distal end 78 of the median spar 54 of the rigid attachment structure 18, as well as a distal end 80 connected to a lateral side. of the longitudinal beam 72. This medial half-arch 74 is shaped so that its distal end 80 is offset rearwardly relative to its proximal end 76.

Each of the front bow 62 and rear 64 has a core 81a extending in an extension of the core 61a of the spar 52, 56 corresponding, and an insole 81b and an outer sole 81c respectively connected to inner and outer edges of the soul 81a of the arch. The outer sole 81c of each of the arches 62, 64 extends in an extension of the upper sole 61c of the spar 52, 56 corresponding to the distal end 66b of the arch, while the insole 81b of the arch extends in an extension of the upper edge 70 of the closing rib 60 of the rigid fastening structure 18, to the longitudinal beam 72. The core 61a of each of the arches is thus in a transverse plane while the insoles 81b and outer 81c of each arch are each substantially shaped cylinder portion of revolution about the longitudinal axis 17. In a similar manner, the middle half-hoop 74 has a soul 81a s' extending in an extension of the core 61a of the median spar 54, and an insole 81b and an outer sole 81c respectively connected to inner and outer edges of the web 81a of the arch. The outer sole 81c of the half-medial arch 74 extends in an extension of the upper sole 61c of the median spar 54 to the longitudinal beam 72, and the insole 81b of the middle half-arch extends in an extension of the upper edge 70 of the closing rib 60 of the rigid fastening structure 18, also to the longitudinal beam 72.

The rigid substructure 28 further comprises an additional hoop 82 attached to a rear end 84 of the longitudinal beam 72 and forming a rear end of the rigid substructure 28. As shown in FIG. 3, the additional hoop 82 is shifted rearwardly with respect to the rigid attachment structure 18, and extends about 60 degrees around the longitudinal axis 17 of the turbine engine being centered with respect to the longitudinal plane V.

As shown in FIG. 3, the rigid substructure 28 further comprises two reinforcing arches 86 which respectively connect two opposite lateral ends 88 of the additional arch 82 to the longitudinal beam 72.

The front bow 62 and the front spar 52 are preferably made of metal and joined to each other by electron beam welding of their respective proximal end 66a and distal end 68 at a welding zone 90 (Figure 4a).

Alternatively, the front bow 62 and the front spar 52 can be secured to one another by other means, for example by splinting. These embodiments of the front bow 62 and the front spar 52 also apply to the rear bow 64 and the rear spar 56, as well as to the medial half-bow 74 and the median spar 54. As it appears in Figure 4a, the front spar 52 has openings 92 for lightening thereof. The front bow 62 also has openings 94, distributed in a proximal half of the bow. In addition, the front bow 62 has three yokes 96 projecting radially inwards with respect to the longitudinal axis 17 of the turbine engine, and intended for the attachment of energy dissipating elements, like this will appear more clearly in the following. One of these screeds 96a is located at one vertex of the front bow 62 while the other two screeds 96b and 96c are located respectively at the proximal and distal ends of the arch so as to be diametrically opposite on either side of the longitudinal axis 17 of the turbine engine. The additional bow 82 is illustrated in greater detail in FIG. 4b. It comprises two yokes 98a and 98b respectively forming the opposite lateral ends of the additional bow 82 and each intended for the attachment of a corresponding engine attachment 30. The additional arch 82 further comprises a rear central engine attachment portion 100 which is centered with respect to the longitudinal plane V, and which essentially comprises a bore 102 with a vertical axis intended to receive a force transmission pin, such as this will become clearer in the following. The additional arch 82 is preferably made of a metallic material and is advantageously formed of three segments assembled end-to-end, for example by an electron beam welding technique at the welding zones 104 roughly shown in FIG. 4b . Alternatively, the attachment of these segments can be made by bolting. In the illustrated example, the additional arch 82 has substantially an I-section and is mainly formed of a core 106a extending substantially in a plane orthogonal to the longitudinal axis 17 of the turbine engine, and an insole 106b and an outer sole 106c respectively connected to inner and outer edges of the core 106a of the additional arch 82 and taking the form of cylinder portions of revolution of respective axes coinciding with the longitudinal axis 17 of the turbine engine. In this example, the outsole 106c extends continuously from one end to the other of the additional arch 82, while the insole 106b comprises two disjoint portions 108 and 110 extending respectively from the two ends. opposing the additional arch towards a middle portion thereof.

In addition, as shown in Figures 3 and 4b, the additional arch 82 has a connecting portion 112 formed projecting forwardly in the axial direction from the portion 100 of rear central engine attachment integrated with this additional roll. This connecting portion 112 has a plane end face 114 extending orthogonally to the longitudinal axis 17 of the turbine engine, and intended to be fixed on a corresponding end face of the longitudinal beam 72 so that the connecting portion 112 above extends said beam. Furthermore, the external aerodynamic fairing 16 comprises a structural skin 30a (FIGS. 2a, 2b, and 5), which has a substantially semi-cylindrical aerodynamic shape of revolution, and which is fixed on the rigid substructure 28. It is note that the structural skin 30a has a concavity oriented towards the longitudinal axis 17 of the turbine engine 12 when the propulsion unit 10 is seen in section along a transverse plane, that is to say orthogonal to the longitudinal direction X, such as the transverse plane P visible in FIG. 3. In the example shown, the structural skin 30a has a circumferential extent equal to about 190 degrees about the longitudinal axis 17, when the propulsion unit 10 is seen in section according to a transverse plane such as the plane P of Figure 3. More generally, this circumferential extent is preferably between 90 degrees and 230 degrees.

Preferably, the structural skin 30a is arranged mainly above the longitudinal axis 17 of the turbine engine, when the aircraft 8 is parked on the ground. More precisely, this structural skin 30a is preferably integrally comprised in a region R of space, visible in FIG. 4a, delimited by two half-planes P1, P2 each delimited by the longitudinal axis 17 and extending symmetrically from on both sides of the longitudinal plane V. Under these conditions, said region R of the space is intended to extend above the half-planes P1, P2. In addition, each of the half planes P1, P2 forms an angle 13 advantageously less than 120 degrees with the longitudinal plane V. In the example illustrated, the angle 13 is thus equal to about 105 degrees. The attachment device 14 further comprises an upper aerodynamic skin 30b having a substantially planar aerodynamic shape and extending laterally in an extension of the aforementioned structural skin 30a so as to cover an upper face of the rigid attachment structure 18 (Figures 2a, 2b, and 5). The structural skin 30a and the upper aerodynamic skin 30b are made of a thermosetting or thermoplastic composite material, but may alternatively be made of metal. In the case of a thermosetting material, these skins can in particular be produced by means of an automatic dispensing machine 30, according to a method known to those skilled in the art used for producing single or double curvature fuselage panels. . In addition to the structural skin 30a, the outer aerodynamic fairing 16 also includes a lower cover 32, an air inlet 34, and complementary panels 36 (Figures 2a and 2b). The lower cover 32 has a distal end 42 articulated on a distal end 44 of the rigid substructure 28, as well as a proximal end 46 provided with means for connection to the rigid attachment structure 18. The lower cover 32 is and movable between an open position, illustrated in Figures 2a and 2b, wherein the proximal end 46 of the lower cover 15 is detached from the rigid attachment structure 18, and a closed position, in which the end proximal 46 of the lower cover is connected to this rigid attachment structure 18. The air inlet 34 is connected to a front end of the rigid substructure 28. This air inlet 34 has an outer skin 48 generally annular centered with respect to the longitudinal axis 17 of the turbine engine and extending the structural skin 30a and the lower cover 32 when the latter is in the closed position. The air inlet 34 also comprises a generally annular inner skin 50 delimiting externally an air intake duct of the turbine engine 12. Each of the aforementioned complementary panels 36 30 is connected to one or more elements of the attachment device 14, among the structural skin 30a, the leading edge fairing 20 and trailing edge fairing 22, the lower cover 32 and the air inlet 34, so as to fill spaces between these elements. Furthermore, the fixation of the structural skin 30a on the rigid substructure 28 is achieved by bolting or riveting this structural skin 30a on the respective outer soles 81c, 106c of the arches 62, 64, the half-arch 74 and the additional arch 82 of the rigid substructure 28. In a similar way, the fixation of the upper aerodynamic skin 30b on the rigid attachment structure 18 is achieved by bolting or riveting this upper aerodynamic skin 30b on the upper flanges 61c respectively of the spars 15 52, 54, 56 of the rigid attachment structure 18. The elements which form the rigid substructure 28, in particular the front arches 62 and rear 64 and the longitudinal beam 72 and the half middle hoop 74, are dimensioned so that a portion of the forces exerted by the turbine engine 12 is transmitted from this turbine engine to the rigid attachment structure 18 via the structural skin This dimensioning thus aims to limit the rigidity of the elements forming the rigid substructure 28, for example by limiting the cross-section of these elements, and / or by limiting the Young's modulus by an adequate choice. of the material or materials forming these elements. In addition, the aforementioned elements which form the rigid substructure 28 as well as the rigid attachment structure 18 are dimensioned so that a part of the forces induced by the turbine engine 12 is transmitted to the rigid attachment structure 18 by the upper aerodynamic skin 30b. As illustrated in FIG. 5, the underside of the rigid attachment structure 18 is streamlined by a lower aerodynamic skin 116 of the attachment device 14, fixed on the respective lower flanges 61b of the longitudinal members of the rigid attachment structure 18. This lower skin 116 is provided with inspection openings 118 allowing access to the inside of the propulsion unit 10 and provided with removable covers, not shown in the figures. The closing rib 60 of the rigid fastening structure 18 also includes such inspection openings 118. In addition, the attachment device 14 comprises an inner skin 120 fixed on the inner soles 81b of the respective front and rear hoops 62 and 64, respectively. and the medial half-hoop 74. This inner skin 120 forms a reinforcement panel participating in the rigidification of the rigid substructure 28. The connection between the rigid substructure 28 and the turbine engine 12 is performed by a plurality of first clevises. intended for the attachment of mechanical energy dissipating elements, hereinafter referred to as dampers, as well as a plurality of second clevises intended for fastening engine fasteners, and the portion 100 of integrated rear central engine fastener at the additional roll 82.

The plurality of first clevises comprises the three clevises 96a, 96b, 96c already described of the front bow 62, as well as two clevises carried by the rear bow 64, one of which 122a is located at a vertex of the while the other 122b is located at the distal end 66b of the arch, as well as a yoke 124 carried by the longitudinal beam 72 and located between the front bow 62 and the middle half hoop 74 These first clevises each have an axis parallel to the longitudinal axis 17 of the turbine engine.

It should be noted that the first clevises 96c and 122b, which respectively belong to the front and rear hoops 62, form, according to the terminology of the invention, distal hooking parts arranged on the same side as the distal end 66b. of each arch with respect to the longitudinal plane V. As shown in FIG. 4a, the aforementioned distal hooking portions 96c and 122b each extend outside an angular sector of angle 02 equal to about 80 degrees. extending from the second half-line D2 described above towards the distal end 66b of the corresponding arch. The angle 02 mentioned above is preferably between 25 degrees and 100 degrees. In addition, the first clevis 96b belonging to the front bow 62 forms, according to the terminology of the invention, a proximal hooking portion, arranged on the same side as the proximal end 66a of the bow and that the rigid structure 18, relative to the longitudinal plane V.

In FIG. 4a, the aforementioned proximal attachment portion 96b extends outside an angular sector of angle θ equating to about 80 degrees extending from the second half-line D2 towards the Rigid hooking structure 18. The aforementioned angle 03 is preferably between 25 degrees and 100 degrees. The plurality of second yokes comprises the two yokes 98a and 98b forming the ends of the additional bow 82, as well as a yoke 126 carried by the longitudinal beam 72 and located substantially at the front end of this beam, and a clevis 128 carried jointly by the proximal end of the medial half-hoop 74 and the distal end 78 of the median spar 54. These second yokes each have an axis parallel to the longitudinal axis 17 of the turbine engine, with the exception of the clevis 126 disposed at the front of the longitudinal beam 72, which has an axis parallel to the transverse direction Y. The first and second said yokes 20 each extend radially inwardly relative to the longitudinal axis 17 turbine engine. As can be seen in FIG. 6, the structural skin 30a has inspection openings 130 which are, for example, four in number. These inspection openings are provided with removable covers, not shown in the figures. One of these inspection openings is formed opposite a viewing opening 132 formed in the inner skin 120 of the attachment device 14, and also visible in FIG. 5.

With reference to FIGS. 7 to 12, the manner in which the turbine engine 12 is connected to the attachment device 14 will now be described. In these figures, the motor fasteners and the energy dissipating elements enabling the turbine engine 12 to be connected to the rigid substructure 28 and to the rigid attachment structure 18 are illustrated by functional symbols that are not related to the geometry of these elements. elements, it being understood that these elements may be of a known type, so that their structure will not be described in detail in the following. Figures 7 to 9 show first an upper engine attachment 134 connecting the yoke 126 carried by the front end of the longitudinal beam 72 to a similar yoke 136 carried by a front casing 138 of the turbine engine 12. The attachment of this upper engine attachment before 134 on the yokes 126 and 136 is such that this engine attachment transmits only to the rigid substructure 28 the 20 forces oriented in the vertical direction Z. For this purpose, the connection of the upper engine attachment before 134 on each of the aforementioned yokes 126, 136 is effected by means of a ball joint. The recovery of the forces oriented in the vertical direction Z is also ensured by two rear lateral engine fasteners 140a and 140b visible in FIG. 11, respectively connecting the clevises 98a and 98b forming the lateral ends of the additional bow 82 to screeds. 142a and 142b of respective axes 30 parallel to the longitudinal axis 17 of the turbine engine 12 and formed projecting radially outwardly on a rear casing 144 of the turbine engine. These rear lateral engine fasteners 140a and 140b are also connected to the abovementioned clevises by means of ball joints making it possible to avoid the recovery of the forces oriented in the longitudinal X and transverse Y directions. FIGS. 8 and 10 make it possible to distinguish a front lateral engine attachment 146 connecting the yoke 128 carried by the proximal end of the middle half-hoop 74 to a yoke 148 of axis parallel to the longitudinal axis 17 of the turbine engine and formed projecting radially outwardly on a casing 150 of the turbine engine 12 disposed between the casings before 138 and back 144 above. The attachment of this front lateral engine attachment 146 on the yokes 128 and 148 is such that this engine attachment transmits only to the rigid substructure 28 and the rigid attachment structure 18 the forces oriented in the transverse direction Y. A this effect, this front lateral engine attachment 146 is connected to the yokes 128 and 148 by ball joints. The engine fasteners 134, 140a, 140b and 146 described above take for example the form of connecting rods or shackles.

In addition, these engine fasteners are preferably of a so-called flexible type, that is to say comprising elastic means allowing relative movement of the ends of the engine attachment. The recovery of the forces oriented in the transverse direction Y is also ensured by a rear central engine attachment 152 (FIG. 11) comprising the portion 100 of motor attachment integrated with the additional bow 82, as well as a fitting 154 formed on a part the top of the rear casing 144 and provided with a shear pin 156 of vertical axis passing through the longitudinal axis 17 of the turbine engine. This shear pin 156, commonly referred to as a spigot, projects upwardly and has an upper portion fitly inserted into the bore 102 (FIG. 4b) of the motor attachment portion 100 integral with the hoop. 82. The tight adjustment of the shear pin 156 in the bore 102 allows the rear central engine attachment 152 to recover not only the aforementioned transverse forces but also the thrust forces of the turbine engine 12, that is to say the forces oriented in the longitudinal direction X. The shear pin 156 is mounted with a clearance along the axis of the bore 102 so as to avoid the resumption of the forces oriented in the vertical direction Z. In addition, the bore 102 is formed in a ball joint (not shown in the figures) itself integrated with the portion 100 of motor attachment integrated with the additional arch 82. Such a connection by ball joint prevents the transmission of forces or moments related to a rotation of the shear pin 156. Furthermore, the recovery of the moment in the vertical direction Z is provided jointly by the front lateral engine attachment 146 and the rear central engine attachment 152.

The recovery of the moment in the transverse direction Y is provided jointly by the upper engine attachment before 134 and by the rear lateral engine fasteners 140a and 140b. Finally, the recovery of the moment in the longitudinal direction X is provided by the rear lateral engine fasteners 140a and 140b. As can be seen in FIGS. 7 and 8, six mechanical energy dissipating elements 158a, 158b, 158c, or dampers, intended for damping the vibrations of the turbine engine 12, respectively connect the yokes 96a, 96b, 96c, 124, 122a, 122b carried by the rigid substructure 28 to yokes 160a, 160b, 160c disposed on the housings 138 and 150 of the turbine engine 12. The two dampers 158c which are respectively connected to the yoke 96c arranged at the distal end 66b of the front bow 62, and the yoke 122b arranged at the distal end 66b of the rear bow 64, each constitute a distal connecting element which connects the turbine engine to the corresponding bow, in the terminology specific to the 'invention. In a similar way, the damper 158b connected to the yoke 96b arranged at the proximal end of the front bow 62 constitutes a proximal connecting element which connects the turbine engine to the front bow 62. As illustrated 12, the yoke 98a which forms one of the ends of the additional bow 82, and which allows the attachment of the rear lateral engine attachment 140a, is fixed on the additional bow 82 by removable fixing means such as bolts, of which only the respective axes 162 are shown. It is the same with regard to the yoke 98b forming the other end of the additional hoop 82. For this purpose, these two yokes 98a, 98b each have a plate 164 applied to an end plate 166 of the additional hoop 82, these plates, also visible in Figure 4b, with respective holes aligned for the passage of the aforementioned bolts.

The other clevises 126 and 128 carried by the rigid substructure 28 and intended for fastening engine fasteners are also attached to the corresponding element of this rigid substructure by removable fastening means.

The fixing method by removable fastening means of the yokes for attachment of the engine fasteners allows attachment / unhitching of the turbine engine 12 to / of the attachment device 14, not by mounting / dismounting the axes of each engine attachment, but by mounting / disassembling the corresponding yoke of each engine attachment on / of the rigid substructure 28. Similarly, the yokes 96a-96c, 122a, 122b, and 124, which are intended for the attachment of the dampers 158 , are preferably also attached to the corresponding element of the rigid substructure 28 by removable fixing means. It should be noted that the set of engine fasteners 134, 140a, 140b, 146, 152 and dampers 158 and the rigid substructure 28 together form connecting means of the turbine engine 12 to the rigid attachment structure 18 , in the terminology of the invention. As is diagrammatically illustrated in FIG. 13, the attachment of the turbine engine 12 to the attachment device 14, as well as the removal of this turbine engine 12, can be effected by a displacement of the turbine engine 12 in translation only in the vertical direction, symbolized by the arrow 168. It should be noted that the turbine engine 12 is shown without its engine fasteners 134, 140a and 140b in Figure 13, for the sake of clarity. This method of attachment and removal of the turbine engine, which therefore has the advantage of not requiring several successive translations in different directions, is facilitated by the fact that the connecting means of the turbine engine 12 to the rigid structure of attachment 18 are entirely included in the region R of the space defined above (FIG. 4a). In particular, it should be noted that no motor attachment is connected to a lower part of the turbine engine 12. It is the same with regard to the dampers 158. More specifically, none of the engine attachments is fully located in a half-space defined in a plane passing through the longitudinal axis 17 and parallel to the transverse direction Y. Indeed, all the engine fasteners are included in a half-space defined above the aforementioned plane, to the except for the front lateral engine attachment 146 which is divided by the aforementioned plane 30 in two parts substantially symmetrical to each other.

Furthermore, Figure 6 illustrates a zone 170, commonly called "UERF area", defined as the area likely to be impacted by a blade of the turbine engine 12 in case of tearing of such a blade.

It appears in this FIG. 6 that the additional arch 82 is located outside the UERF zone 170. As a result, the rear lateral engine fasteners 140a, 140b as well as the rear central engine attachment 152 are protected from dawn impact.

In addition, the upper engine attachment before 134 being designed for a recovery of vertical forces only, this engine attachment has a clearance in the transverse direction Y at the yokes 126 and 136 (Figure 9) which this engine attachment is connected. However, in the event of accidental destruction of the front lateral engine attachment 146 intended to take up the transverse forces and participating in the recovery of the moment in the vertical direction Z, the upper engine attachment before 134 naturally tends, in its movements according to the transverse direction Y, abutting in the yokes 126 and 136, and thus replaces the deficient engine attachment to ensure the recovery of the moment in the vertical direction Z together with the rear lateral engine fasteners 140a and 140b located outside the UERF zone 170. In a similar way, the front lateral engine attachment 146 intended for the recovery of forces in the transverse direction Y has a displacement in the vertical direction Z at the yokes 128 and 148 (FIG. 10) to which this fastener motor is connected. However, in the event of accidental destruction of the upper engine attachment before 134 intended to take up the vertical forces and participating in the recovery of the moment in the transverse direction Y, the front engine attachment 146 naturally tends, in its vertical displacements, to abut in the yokes 128 and 148, and thus replaces the deficient engine attachment to ensure the recovery of the moment in the transverse direction Y together with the rear lateral engine fasteners 140a and 140b located outside the UERF area 170. The configuration of the propulsion unit 10 thus makes it possible to satisfy the regulatory requirements with regard to the risk of dawn loss, while avoiding the use of additional engine fasteners, called "waiting", which do not participate in the recovery. of effort in normal operation of the turbine engine.

Alternatively, the longitudinal members 52, 54, 56 and / or the ribs 58, 60 of the rigid fastening structure 18 may not be made of metal but of a composite material. In another variant, the rigid attachment structure 18 may not comprise longitudinal members and ribs, but may be a honeycomb structure, also called honeycomb, or be formed of a solid foam. In addition, the structural skin 30a may have a circumferential extent smaller than the circumferential extent of the rigid substructure 28. The latter is then only partially covered by the structural skin 30a. Furthermore, although the lower cover 32 described above only fulfills an aerodynamic function and does not participate in the transmission of the forces induced by the turbine engine, it is possible alternatively to size this lower cover 32 and the constituent elements the attachment device 14 so that the lower cover 32 plays a structural role, that is to say participates in the transmission of the forces induced by the turbine engine, in particular to the transmission of the torsion moment in the longitudinal direction X. On the other hand, the motor attachment portion 100 including the bore 102 may not be integrated in the additional arch 82 but take the form of a fitting attached to this additional arch and integrating the bore 102. In addition, the upper front engine attachment 134 may be replaced or supplemented by a linear actuator to enable the inclination of the longitudinal axis of the turbine engine 12 to be altered with respect to in the horizontal direction. Such a modification of the inclination is facilitated by the fact that the rear central engine attachment 152 incorporates a ball joint, as explained above. FIG. 14 illustrates a propulsion assembly 10 according to another embodiment of the invention, in which the attachment device 14 does not include a front lateral engine attachment such as the engine attachment 146 described above, but comprises two upper engine attachments 172a and 172b respectively connecting two yokes 174a, 174b secured to the front casing 138 of the turbine engine 12 to a spreader 176 itself hooked to a fitting 178 carried by the longitudinal beam 72. The two yokes 174a, 174b above are arranged on either side of the longitudinal plane V, so that the two upper front engine fasteners 172a and 172b are arranged substantially in an inverted "V" configuration. The spreader 176 has two yokes 180a, 180b also arranged on either side of the longitudinal plane V, respectively in alignment with the two yokes 174a, 174b integral with the front casing 138.

Each of the upper front engine fasteners 172a and 172b is connected to the corresponding yoke 174a, 174b of the front casing 138 by a ball joint connection, and to the corresponding yoke 180a, 180b of the yoke 176 by a pivot connection. In addition, the spreader is itself connected to the fitting 178 carried by the longitudinal beam 72 by a ball joint having a play in the longitudinal direction X. The two upper front engine fasteners 172a and 172b together allow a recovery efforts directed in the vertical direction Z as well as forces directed in the transverse direction Y. FIGS. 15 to 16 illustrate a propulsion unit 10 according to another embodiment of the invention, in which the attachment device 14 comprises an engine attachment. lateral 182 dedicated to the recovery of efforts oriented along the longitudinal direction X, and not the recovery of forces in the transverse direction Y as in the embodiments described above. This lateral engine attachment 182 comprises a fitting 184 fixed to the rigid attachment structure 18, for example on the distal end 78 of the median spar 54. This fitting 184 comprises a bore 186 in which a shear pin 188 is inserted, or spigot, integral with the casing 150 of the turbine engine. The bore 186 has an elongated shape in the vertical direction Z (FIG. 15a) so that the shear pin 188 is locked in the bore 186 in the longitudinal direction X but free to move in the vertical direction Z. In this embodiment, it is not necessary that the rear central engine attachment 152 participates in the recovery of the longitudinal forces, so that it can be dedicated only to the recovery of forces in the transverse direction Y. In this case as illustrated in FIG. 16, the bore 102 of this rear central engine attachment 152 may be elongate in the longitudinal direction so that the corresponding shear pin 156 is locked in the transverse direction Y but not in the longitudinal direction X In addition, in this embodiment, the forward engine attachment is preferably of the same type as that of FIG. orts oriented in the vertical direction Z and the recovery of forces oriented in the transverse direction Y. Alternatively, the recovery of the longitudinal forces can be provided by two thrust rods of known type, each having a rear end connected to the rear casing 144 of the turbine engine and a front end connected to the distal end 26 of the attachment structure 18. In this case, the recovery of the transverse forces can be provided by a lateral engine attachment of the same type as the attachment motor 146 described above with reference to Figures 7 to 12. In addition, the upper front engine attachment can then be dedicated to the recovery of the vertical forces as in the first embodiment described above, and can therefore be the same type as the upper front engine attachment 134 described above (Figure 9). FIG. 17 illustrates a propulsion assembly 10 according to another embodiment of the invention, in which the structural skin 30a and the upper aerodynamic skin 30b have generally the same conformation as in the embodiments described above, but are not however not made in one piece. These two skins 30a, 30b are here spliced to one another. The structural skin 30a has a proximal end portion 190 on which is applied a first portion 192 of a splint 194, for example by means of bolts or rivets of which only the pins 196 are shown in Figure 17. The splint 194 comprises a second portion 198 applied to a distal end 200 of the upper aerodynamic skin 30b, also by means of bolts or rivets of axes 202. The splint 194 preferably extends over the entire length of the rigid attachment structure 18.

In the example of Figure 17, the longitudinal members 52, 54, 56 of the rigid attachment structure 18 are respectively fixed to the front bow 62, the middle half-bar 74, and the rear bow 64, or directly by means of bolts or rivets (not shown) axes 204, or via a splint. Alternatively, as shown in Figure 18, the longitudinal members 52, 54, 56 may be respectively connected to the front bow 62, the middle half-hoop 74, and the rear bow 64, by respective yokes 206. In any case, as illustrated in FIG. 19, the structural skin 30 a can be extended downwards between each pair of successive longitudinal members of the rigid fastening structure 18, so as to allow the attachment of this structural skin 30 a by This splicing is for example carried out by means of two fishplates 208 (only one being visible in FIG. 19), the first of which is arranged between the front spar 52 and the median spar 54, while the second is disposed between the median spar 54 and the rear spar 56. In the illustrated example, these fishplates 208 are applied to the radially inner face of the structural skin 30a, but these fishplates 208 can alternatively be applied on the radially outer face thereof. The invention can of course be transposed to other types of turboshaft engines, in particular to a turbojet engine comprising a pair of counter-rotating, non-ducted propellers mounted at the front. In this case, the arrangement of the attachment device 14 may be inverted in the longitudinal direction, the additional arch 82 then being arranged at the front. 10

Claims (8)

REVENDICATIONS1. A propulsion unit (10) for an aircraft, comprising a turbine engine (12), an external aerodynamic fairing (16) surrounding at least a portion of said turbine engine, and a device (14) for attaching the turbine engine comprising a rigid attachment structure (18) having a proximal end (24) intended to be connected to an aircraft structure and a distal end (26) arranged facing the turbine engine, said rigid attachment structure (18) carrying connecting means (28, 134, 140a, 140b, 146, 152, 172a, 172b, 182, 158a, 158b, 158c) connecting said turbine engine (12) to said rigid attachment structure (18), characterized in that: - said connecting means comprise a rigid substructure (28) surrounding at least a portion of said turbine engine (12) and connected to said rigid attachment structure (18); said external aerodynamic fairing (16) comprises at least one structural skin (30a) connected to said rigid attachment structure (18), and fixed on said rigid substructure (28) so as to cover at least in part this sub-structure rigid structure (28); when said structural skin (30a) of the external aerodynamic fairing (16) is seen in section along a transverse plane (P) orthogonal to a longitudinal axis (17) of said turbine engine, said structural beam (30a) has a concave shape in the direction of said longitudinal axis (17); - Said connecting means are shaped so that a portion of the turbine engine connecting forces (12) to said rigid attachment structure (18) passes through said structural skin (30a) of said external aerodynamic fairing (16). The propulsion assembly of claim 1, wherein said structural skin (30a) has a circumferential extent of between 90 degrees and 230 degrees about said longitudinal axis (17). 3. propulsion unit according to claim 2, wherein: - said structural skin (30a), and said connecting means connecting said turbine engine (12) to said rigid attachment structure (18), are integrally included in a region ( R) of the space delimited by two half-planes (P1, P2); - said half-planes (P1, P2) are each delimited by said longitudinal axis (17) and extend symmetrically on either side of a longitudinal plane (V) including said longitudinal axis (17); each of said half-planes (P1, P2) forms an angle 13 less than 120 degrees with said longitudinal plane (V); said longitudinal plane (V) is intended to extend vertically, and said region (R) of the space is intended to extend above said half-planes (P1, P2), when said propulsion unit (10) ) equip an aircraft parked on the ground. 4. propulsion unit according to any one of claims 1 to 3, wherein said rigid attachment structure (18) extends substantially laterally relative to the turbine engine (12) so that the proximal end (24) of this Rigid attachment structure (18) can be attached laterally to an aircraft structure (11). A propulsion assembly according to claim 4 comprising an upper aerodynamic skin (30b) which at least partially covers said rigid attachment structure (18) and which is connected to a proximal end portion of said structural skin (30a). said external aerodynamic fairing (16). 6. propulsion unit according to any one of claims 1 to 5, wherein said structural skin (30a) of said external aerodynamic fairing (16) comprises at least one inspection opening (130) provided with a cover for access to the turbine engine (12). 7. Propulsion unit according to any one of claims 1 to 6, comprising at least one reinforcing panel (120) disposed between the turbine engine (12) and said rigid substructure (28) and fixed thereto. 8. Aircraft (8), characterized in that it comprises at least one propulsion unit (10) according to any one of the preceding claims.