Abstract:
The present invention relates, on the one hand, to a nacelle ( 1 ) for a double flow turboreactor ( 2 ) comprising a front air intake section ( 5 ), a median section ( 6 ) intended for surrounding a fan ( 3 ) of the turboreactor and a rear section ( 7 ), said rear section having an internal structure ( 7   b ) intended for serving as a housing to a rear portion of the turboreactor, characterized in that the internal structure possesses attachment means suitable for enabling the nacelle to be fastened to a pylon ( 12 ) intended to be connected to a fixed structure ( 13 ) of an aeroplane on at least one portion of said internal structure, and on the other hand, to a propellant assembly and to an aircraft provided with such a nacelle.

Description:
TECHNICAL FIELD OF THE INVENTION 
     The present invention relates to a nacelle for a turbofan. 
     BRIEF DISCUSSION OF RELATED ART 
     An aircraft is propelled by several turbojets each housed in a nacelle also accommodating an assembly of associated actuation devices linked to its operation, such as a thrust reverser device, and performing various functions when the turbojet is in operation or stopped. 
     A nacelle usually has a tubular structure comprising an air intake upstream of the turbojet, a mid-section designed to surround a fan of the turbojet, a downstream section accommodating thrust reverser means and designed to surround the combustion chamber of the turbojet, and is usually terminated by an exhaust nozzle whose outlet is situated downstream of the turbojet. 
     Modern nacelles are designed to accommodate a turbofan capable of generating, by means of the air foils of the fan in rotation, a flow of hot air (also called the main flow) originating from the combustion chamber of the turbojet, and a flow of cold air (the bypass flow) which travels on the outside of the turbojet through an annular passageway, also called a stream, formed between a fairing of the turbojet (or an internal structure of the downstream structure of the nacelle and surrounding the turbojet) and an internal wall of the nacelle. The two air flows are exhausted from the turbojet through the rear of the nacelle. 
     Each propulsion system of the aircraft is therefore formed by a nacelle and a turbojet, and is suspended on a fixed structure of the aircraft, for example beneath a wing or on the fuselage, by means of a pylon attached to the turbojet in its front and rear portions by suspension elements. 
     In such a configuration, it is the turbojet that supports the nacelle. 
     Such an architecture sustains many combined external forces during the aircraft&#39;s mission. 
     Amongst other things these are forces resulting from gravity, external and internal aerodynamic forces, gusts of wind, and thermal effects. 
     These stresses applied to the propulsion system are transmitted to the turbojet and cause deformations of casings which directly impact the performance of the various stages of the turbojet. More particularly, in the case of a propulsion system called a wasp-waisted propulsion system, that is to say having a long downstream portion that is relatively thin relative to the intermediate and air intake structures, these stresses result in a particularly harmful deformation called a “banana-shaping” deformation, the downstream portion curving considerably. 
     Such a “banana-shaping” is reflected by a deformation of the external structure of the nacelle formed by the various successive casings while the drive shaft, the blades of the fan and the internal blades of the turbojet remain rectilinear. The result of this is that the heads of the blades of the shaft move closer to the internal periphery of the casings. The general performance of the turbojet is thereby reduced relative to a configuration in which the casings sustain little or no deformations, because it is then necessary to take account of this deformation in the design of the nacelle so as always to arrange a sufficient clearance between the heads of the blades and the periphery of the casings. This results in a portion of the supply air that is not compressed by the blades because it escapes through this considerable clearance. 
     BRIEF SUMMARY OF THE INVENTION 
     The invention alleviates the aforementioned disadvantages, and for this reason includes in a nacelle for a turbofan comprising a front air intake section, a mid-section designed to surround a fan of the turbojet and a rear section, said rear section having an internal structure designed to serve as a casing to a rear portion of the turbojet, characterized in that the internal structure has coupling means suitable for allowing an attachment of the nacelle to a pylon designed to be connected to a fixed structure of an aircraft on at least one portion of said internal structure. 
     Therefore, by allowing the pylon to be directly attached to a structure of the nacelle instead of attaching it directly to the turbojet, it is the nacelle which supports the turbojet. In this manner, the turbojet does not have to sustain and transmit the deformations of the nacelle and vice versa. As explained above, it is then possible to optimize the clearance existing between the fan blades and the blades inside the turbojet and their respective casings in order to improve the performance of the propulsion system. 
     Preferably, the internal structure is fitted with means for rigid connection to the turbojet, for example by bolting. 
     Advantageously, the internal structure is connected to the mid-section by means of a casing surrounding the fan. 
     Preferably, the internal structure is connected to the mid-section of the downstream portion of the casing surrounding the fan on at least a portion of its periphery by means of a groove of the latter. 
     Again preferably, the internal structure is connected to the mid-section of the downstream portion of the casing over the whole of its periphery. Evidently, this attachment may be made only on a portion of the periphery of the groove. 
     Advantageously, the peripheral groove of the casing has a V-shaped internal profile. 
     Again advantageously, the internal structure is fitted with a means for recentering the turbojet. 
     Preferably, the internal structure is designed so that the pylon can extend over the whole length of the internal structure. 
     Advantageously, the pylon is incorporated into the internal structure. 
     Preferably, the internal structure comprises at least one external wall forming an aerodynamic surface mounted on a framework. Advantageously, the external wall is partially or totally made from at least one acoustic panel. In this manner, the external wall fulfills no structural role, this function being performed by the framework, and it can therefore be lightened to the maximum without it being necessary to provide high-density structural zones in this wall. In the case of an acoustic panel, it is therefore possible to dedicate the whole surface of the acoustic panel to the acoustic function without it being necessary to provide structural zones which prevent any acoustic function. 
     According to a first variant embodiment, the framework only partially surrounds the turbojet, preferably over at least 180°. 
     According to a second variant embodiment, the framework totally surrounds the turbojet. 
     Advantageously, the framework of the internal structure is made from radial frames. Again advantageously, the radial frames are made from force-absorbing link rods. 
     Preferably, at least a portion of the radial frames are made in a single piece. 
     In an alternative or complementary manner, at least a portion of the radial frames are made from several elements linked together, for example by bolting. 
     Advantageously, the framework of the internal structure is made from radial frames distributed over the length of the internal structure. 
     Preferably, the framework comprises at least one front radial frame and one rear radial frame connected by an intermediate structure forming a mesh. 
     Advantageously, the intermediate structure is made in the form of a caisson. 
     Advantageously, the intermediate structure is made from link bars connecting at least two radial frames together. 
     Again advantageously, at least a portion of the link bars are incorporated into at least one radial frame. 
     Preferably, the link bars are hollow. 
     Advantageously, the link bars are placed relative to one another so as to form triangles, preferably isosceles triangles. 
     In a yet more advantageous manner, the framework comprises at least one longitudinal reinforcement on either side of a longitudinal axis of the internal structure. 
     Preferably, the internal structure comprises at least one link rod for absorbing thrust attached, on the one hand, to at least one point of an upstream portion of the internal structure, for example at a horizontal mid-plane, and, on the other hand, at at least one point of a downstream portion of the internal structure in the vicinity of the pylon or optionally incorporated into the latter. 
     The presence of such thrust-absorbing link rods mounted obliquely makes it easier to transmit longitudinal forces to the pylon. 
     Advantageously, the link rod for absorbing thrust is oriented substantially in the structural alignment of the pylon. 
     Again advantageously, the link rod for absorbing thrust has a fork attached to the internal structure at at least two points of the upstream portion of the structure, on either side of the horizontal mid-plane, the fork of the link rod having a junction point situated, for example, at a radial frame of the framework. 
     Preferably, at least one portion of the framework elements, namely in particular the radial frames, force-absorbing link rods, intermediate structure and longitudinal reinforcements, are fitted with a heat protection. 
     According to a first variant embodiment, the framework is made in one piece. 
     According to a second variant embodiment, the framework is made in two half-pieces designed to be assembled substantially vertically. 
     The present invention also relates to an aircraft, characterized in that it comprises at least one propulsion system comprising a nacelle according to the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The implementation of the invention will be better understood with the aid of the detailed description that is explained below with reference to the appended drawing in which: 
         FIG. 1  is a schematic representation in perspective of a nacelle according to the invention attached to a pylon by means of an internal structure surrounding the turbojet. 
         FIG. 2  is a view in longitudinal section of the nacelle of  FIG. 1 . 
         FIG. 3  is a partial schematic view showing the arrangement of the internal structure relative to a casing of the fan. 
         FIG. 4  is a schematic representation of the structure of  FIG. 3  with a complete internal structure attached to the pylon. 
         FIG. 5  is a representation in solid lines of  FIG. 4  with the internal structure accommodating the turbojet. 
         FIG. 6  is a view in cross section of the nacelle of  FIG. 1 . 
         FIG. 7  is a schematic representation of a first variant embodiment of the internal structure. 
         FIG. 8  is a schematic representation of a second variant embodiment of the internal structure. 
         FIG. 9  is a simplified illustration of a recentering means fitted to the internal structure. 
         FIG. 10  is a view in cross section of a nacelle according to the invention with an internal structure fitted with means for recentering the turbojet. 
         FIGS. 11 and 12  are representations respectively in perspective and from the side of a third embodiment comprising a short internal structure. 
         FIGS. 13 and 14  are representations of the structure represented in  FIGS. 11 and 12  in a turbojet-supporting situation. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIGS. 1 and 2  represent a nacelle  1  for a turbofan  2 . 
     The nacelle  1  forms a tubular housing for a turbofan  2  and is used to channel the air flows that it generates by means of the air foils of a fan  3 , namely a hot air flow passing through a combustion chamber  4  of the turbojet  2  and a cold air flow traveling outside the turbojet  2 . 
     The nacelle  1  has a structure comprising a front section forming an air intake  5 , a mid-section  6  surrounding the fan  3  of the turbojet  2 , and a rear section  7  surrounding the turbojet  2  and comprising a thrust reverser system. 
     The air intake  5  has an internal surface  5   a  designed to channel the intake air and a fairing external surface  5   b.    
     The mid-section  6  comprises, on the one hand, an internal casing  6   a  surrounding the fan  3  of the turbojet  2 , and, on the other hand, a fairing external structure  6   b  of the casing extending the external surface  5   b  of the air intake section  5 . The casing  6   a  is attached to the air intake section  5  which it supports and extends its internal surface  5   a . In addition, the casing  6   a  is connected to an upstream casing  6   c  of the turbojet  2  by means of radial struts  8  placed crosswise. Quite evidently there could be more than four radial struts, particularly on a turbojet of the CFM type. 
     The rear section  7  comprises an external structure  7   a  comprising a thrust reverser system forming an exhaust nozzle and a fairing internal structure  7   b  of the turbojet  2  defining with the external structure  7   a  a stream  9  designed for the circulation of the cold flow. 
     The internal structure  7   b  is made of a structural framework  10  covered with acoustic panels  11  producing an internal aerodynamic surface of the stream  9 . Accordingly, the acoustic panels  11  are not structural and may be lightened to the maximum, the whole surface of said acoustic panels  11  being able to be dedicated to the acoustic function without needing to provide structural zones preventing any acoustic element. 
     The structural framework  10  is designed to be attached directly to a mast  12  itself designed to be attached to a fixed portion of an aircraft such as a wing  13 . 
     The structural framework  10  is made from two half-portions  14 , one of which is represented in  FIG. 3  in perspective with the casing  6   a  of the fan  3 , designed to be attached together. 
     Each half-portion  14  has a series of radial frames  15  that are distributed over the whole length of the half-portion  14  and whose number and section are defined according to the forces to be made to pass through. 
     More precisely, each half-portion has an upstream radial frame  15   a  associated with a top strut  16   a  and a bottom strut  16   b  which, together with the upstream radial frame  15   a , are designed to serve as a connection interface between the structural framework  10  and the mid-section  6  by means of the upstream casing  6   c  and the vertical struts  8 . 
     The radial frames  15  are connected together by at least one longitudinal reinforcement  17  and by a top longitudinal reinforcement  18  and a joining bottom longitudinal reinforcement  19 . Furthermore, the half-portion  14  has a downstream top strut  20   a  and a downstream bottom strut  20   b  which supplement the half-portion  14  in order to allow a structural connection by a top reinforcement  21   a  and a bottom reinforcement  21   b  respectively connecting the struts  16   a  and  20   a  and  16   b  and  20   b  together. Other top and bottom struts may be added, for example in continuity with the radial frames  15 . 
     The transmission of the forces is improved by adding to each half-portion  14  a force-absorbing link rod  22 , as can be seen in  FIG. 4 , attached, on the one hand, upstream of the half-portion  14  at a mid-plane of the structural framework  10 , that is to say substantially at the longitudinal reinforcement  17  and the upstream radial frame  15   a , and, on the other hand, downstream of the half-portion  14  at a point designed to come close to the mast  12 , that is to say substantially on the top longitudinal reinforcement  18  and close to a downstream radial frame  15   b . Advantageously, the force-absorbing link rod  22  is therefore oriented in a direction that is substantially identical to the direction of the mast  12 . Alternatively, the downstream coupling point of the force-absorbing link rod  22  may be incorporated into the mast  12 . 
     Each half-portion  14  is connected to the other half-portion via its bottom portion, by means of their upstream bottom struts  16   b  and downstream bottom struts  20   b , and by means of the bottom longitudinal reinforcements  19  and bottom reinforcements  21   b.    
     In the top portion, each half-portion  14  is connected to the mast  12  by means of their upstream top struts  16   a  and downstream top struts  20   a , and by means of the top longitudinal reinforcements  18  and top reinforcements  21   a.    
     Alternatively, the mast may be incorporated into the structural framework  10 . 
       FIG. 5  represents the inside of the nacelle  1 , once the structural framework  10  has been covered by the acoustic panels  11 . 
       FIG. 6  shows a front view in section of the internal structure  7   b  thus assembled. 
       FIG. 7  shows a variant embodiment of the structural framework  10 . A structural framework  110  according to  FIG. 7  is made from two half-portions  114  that differ only from a half-portion  14  by the fact that each half-portion  114  comprises a force-absorbing link rod  122  having an upstream fork. Such a force-absorbing link rod  122  is therefore attached to the half-portion  114  at three points, namely two downstream points  114   a ,  114   b  situated at the upstream radial frame  15   a  on either side of the mid-plane of the structural framework  110 , that is to say on either side of the longitudinal reinforcement  17 , and at a point  114   c  situated downstream at the same location as for the force-absorbing link rod  22 . Preferably, the fork of the force-absorbing link rod  122  joins at a point  114   d  substantially situated at a radial frame  15  and is attached thereto. 
       FIG. 8  shows a structural framework  210  made in a single piece that is open at the top portion only, the portion by which it is designed to be attached to the mast  12 . 
     The structural framework  10 ,  110 ,  210  is supplemented by means for recentering between the turbojet  2  and the internal structure  7   b  situated downstream of the latter. The operating principle of the recentering means is shown in  FIG. 9 . 
     The recentering means provides a permanent contact between the turbojet  2  and the internal structure  7   b  so as to take account of a differential movement between these two structures due to the thermal expansion of the turbojet  2  in operation causing a longitudinal and axial movement of the latter. 
     To do this, the turbojet has, downstream of its structure, radial extensions  30  distributed over the whole of its circumference and each terminated by a ramp  31  in sliding contact with a complementary ramp  32  of an internal radial extension  33  of the structural framework  10 ,  110 ,  210 . The ramps  31 ,  32  are designed so that their orientation corresponds substantially to the estimated movement differential between the two structures. 
     The recentering system may be made in various ways, notably by elastic contact, by distinct or one-piece elements, on only one sector of the periphery of the turbojet  2  or on the whole of its periphery. 
       FIG. 10  shows a front view in section showing a distribution of the recentering means. 
     It will also be noted that the invention allows easier maintenance of the turbojet  2 , the access to the latter being able to be made simply by removing the acoustic panels  11  without the need to dismantle the whole internal structure  7   b.    
     It will also be noted that the internal structure  7   b  may optionally comprise a bottom structure allowing the coupling of a rear external structure  41  in the bottom portion. In this case, the result is a distance between the point of attachment of said rear external structure  41  and the downstream circumferential recentering zone of the turbojet  2 . This distance provides a force component which tends to separate the bottom structure from the internal structure  7   b  by which the rear external structure  41  is attached which no longer allows the recentering means to fulfill their function in this zone. Accordingly, it will be possible to ensure the integrity of the maintenance of the recentering by a system of locks  40  at the junction between the two half-portions  14 ,  114  at the downstream radial frames  15   b.    
       FIGS. 11 to 14  show a particular variant embodiment of the invention comprising a short internal structure  310  also maintained at a casing of the fan. 
     The support system alone is shown in  FIGS. 11 and 12 . 
     The latter comprises attachment means of the pylon  12  type to which the internal structure  310  is connected. 
     The internal structure  310  is shown in the form of a peripheral structural framework made from a front peripheral radial frame  315   a  and a rear peripheral radial frame  315   b.    
     The front radial frame  315   a  and the rear radial frame  315   b  are connected together by an intermediate structure  316  forming a mesh made from link bars  316   a ,  316   b  together forming substantially isosceles triangles. 
     The support assembly is supplemented by suspension elements  320  mounted on the structure of the pylon  12  type and designed to be connected close to one end of the turbojet  2 . 
       FIGS. 13 and 14  show the support assembly previously described in a situation of support of a turbojet  2 , the framework  310  being connected to the casing  6   c  by means of a bolting system via the front radial frame  315   a  installed in a V-shaped peripheral groove of the casing  6   c.    
     Evidently, as mentioned above, the internal structure may, as a variant, be made in the form of one or more sectors that are not entirely peripheral. 
     Although the invention has been described with particular exemplary embodiments, it is evident that it is in no way limited thereto and that it comprises all the technical equivalents of the means described and their combinations if the latter enter into the context of the invention.