Abstract:
The disclosure relates to an aircraft propulsion assembly comprising a bypass turbojet engine equipped with a nacelle, the bypass turbojet engine including a structure defining a first part of a secondary flow path for channelling secondary flow, and the nacelle having a structure defining a second part of the secondary flow path. The structure of the nacelle defining the second part of the secondary flow path is arranged such that the first part and the second part of the secondary flow path are angularly offset around a longitudinal axis of the engine when the engine is shut down/stopped.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is a continuation of International Application No. PCT/FR2015/051806, filed on Jul. 1, 2015, which claims the benefit of FR 14/56408 filed on Jul. 3, 2014. The disclosures of the above applications are incorporated herein by reference. 
     
    
     FIELD 
       [0002]    The present disclosure relates to to an aircraft propulsion unit, the propulsion unit being constituted by a turbojet engine and a nacelle. 
       BACKGROUND 
       [0003]    The statements in this section merely provide background information related to the present disclosure and may not constitute prior art. 
         [0004]    An aircraft is propelled by several turbojet engines each housed within a nacelle, each nacelle further accommodating an assembly of additional actuating devices linked to its operation and ensuring various functions when the turbojet engine is in operation or stopped. 
         [0005]    The modern nacelles are intended to accommodate a bypass turbojet engine capable of generating, by means of the fan blades in rotation, a flow of hot gases (also called primary flow) and a flow of cold air (also called secondary flow) witch circulates outside the turbojet engine through an annular passage, also called flow path, formed between two concentric walls of the nacelle. The primary and secondary flows are ejected from the turbojet engine by the rear of the nacelle. 
         [0006]    A turbojet engine nacelle generally has a tubular structure including, from upstream to the downstream (relative to the direction of the cold and hot flows):
       a front section, or air inlet, located in front of the turbojet engine;   a median section, intended to surround a fan module of the turbojet engine;   a rear section, intended to surround a high-pressure module, which in particular includes the combustion chamber of the turbojet engine, and generally embarking thrust reversal means;   an ejection nozzle, whose outlet is located downstream of the turbojet engine.       
 
         [0011]    The rear section generally has a fixed external structure, called “Outer Fixed Structure” (OFS), which defines, with a concentric internal structure, called “Inner Fixed Structure” (IFS), a downstream portion of the secondary flow path serving to channel the flow of cold air. The rear section is positioned downstream of a fan module of the turbojet engine which comprises in particular: a fan casing (inside which the fan is contained) and an intermediate casing. The intermediate casing includes a hub and an outer annular casing, as well as radial link arms therebetween. 
         [0012]    Each propulsion unit of the aircraft is thus formed by a nacelle and a turbojet engine, and is suspended from a fixed structure of the aircraft, for example under a wing or on the fuselage, by means of a pylon or a mast fastened to the turbojet engine or to the nacelle. 
         [0013]    It is thus observed that an aircraft propulsion unit integrates functional subassemblies likely to enter in relative movements, and between which it is suitable to manage the sealing. 
         [0014]    In particular, it is important that the rear section of the nacelle, which delimits the secondary flow path, can be correctly aligned with the intermediate casing, with which it cooperates to channel the flow of cold air without leakage and without aerodynamic losses. Such a leakage would be particularly harmful, because a nacelle is designed and dimensioned to withstand the pressure exerted by the cold flow, in the case where it is correctly channeled. In contrast, the nacelle is not designed to withstand the forces generated by the pressure exerted by an air leakage of the secondary flow path towards the turbojet engine. Such a leakage can thus lead to a detachment of the inner structure of the nacelle. In view of these constraints, it is therefore essential to provide for a sealing barrier between the upstream portion of the rear section and the turbojet engine, in order to prevent any leakage of the secondary flow path towards the turbojet engine. 
         [0015]    However, the sealing between the two covers and the turbojet engine presents a particular problem. First of all, the elements constituting the rear section of the nacelle are, in operation, animated by axial and radial movements relative to the turbojet engine. Given the large dimension of the concerned parts, these relative movements can, in operation, result in important displacements. 
         [0016]    On the other hand, in operation, during the flight phases, the engine also undergoes deformations. In particular, the torsional forces generated by the rotation at very high speed of the fan blades lead the engine to be deformed about its longitudinal axis. This torsional movement, known under the name of “fan twist,” leads to an angular offset between the front part (the fan module, including in particular the intermediate casing) and the rear part (including in particular the combustion chamber) of the engine. 
         [0017]    This angular offset is consequently also induced between the intermediate casing and the inner fixed structure. A gasket interposed between the inner fixed structure and the turbojet engine must therefore create a sealing barrier whatever the relative position of the inner fixed structure with respect to the turbojet engine, and for that, it must have a high crushing amplitude. 
         [0018]    However, even by providing for such a gasket, the angular deformation of the engine in operation has severe disadvantages, among which is the reduction of the aerodynamic qualities of the secondary flow path. Indeed, the alignment of the inner fixed structure of the intermediate casing, which is correct when the engine is stopped, may become defective in flight. Indeed, the angular offset (about the longitudinal axis of the engine) between the inner fixed structure and the engine results in a deviation between some engine walls located in the flow of cold air when the engine is in operation, and which should normally be aligned with corresponding walls of the inner fixed structure. These walls are for example constituted by the outer surfaces of the link arms of the intermediate casing (and in particular those located in the positions called “6h00” and “12h00” positions. These alignment deviations generate a discontinuity of the aerodynamic lines of the secondary flow path, which greatly reduces the aerodynamic qualities of the secondary flow path. 
       SUMMARY 
       [0019]    The present disclosure relates to an aircraft propulsion unit, including a bypass turbojet engine equipped with a nacelle, the turbojet engine including a structure defining a first portion of the secondary flow path intended to channel the secondary flow, the nacelle including a structure defining a second portion of the secondary flow path, the assembly being characterized in that the structure of the nacelle defining the second portion of the secondary flow path is arranged so that the first portion and the second portion of the secondary flow path are offset angularly about the longitudinal axis of the engine when the engine is stopped. 
         [0020]    In accordance with the present disclosure, when the engine is stopped, there is an angular offset about the longitudinal axis of the engine between the structure of the nacelle defining the second portion of the secondary flow path and the structure of the turbojet engine defining the first portion of this flow path. Thus, by providing for an initial angular offset between the portion of the secondary flow path delimited by the engine and the portion of the secondary flow path delimited by the nacelle, the torsional deformation undergone by the engine during the flight phases will be taken into account. In other words, the alignment between the two portions of the secondary flow path is voluntarily deteriorated when the engine is stopped, in order to be improved when the engine is in cruising speed. 
         [0021]    In one form, the first portion of the secondary flow path is delimited in particular by an intermediate casing of the engine. 
         [0022]    In another form, the second portion of the secondary flow path is delimited by the rear section of the nacelle. 
         [0023]    In still another, the second portion of the secondary flow path is delimited by an inner fixed structure and an outer fixed structure. 
         [0024]    In yet another, a gasket is interposed between the inner fixed structure and the engine. 
         [0025]    In one form, the angular offset value is comprised between 2° and 10°. 
         [0026]    In one embodiment, the angular offset value is comprised between 2° and 5°. 
         [0027]    In another form, the angular offset value is comprised between 5° and 10°. 
         [0028]    The present disclosure also concerns an aircraft including one or several propulsion unit(s) as defined above. 
         [0029]    Further areas of applicability will become apparent from the description provided herein. It should be understood that the description and specific examples are intended for purposes of illustration only and are not intended to limit the scope of the present disclosure. 
     
    
     
       DRAWINGS 
         [0030]    In order that the disclosure may be well understood, there will now be described various forms thereof, given by way of example, reference being made to the accompanying drawings, in which: 
           [0031]      FIG. 1 , shows one form of a propulsion unit in accordance with the teachings of the present disclosure; 
           [0032]      FIGS. 2 a  and 2 b    show a schematic view of a section of an intermediate casing of a turbojet engine, when stopped and in operation, respectively; 
           [0033]      FIGS. 3 a  and 3 b    show a schematic view of a section of a propulsion unit, when stopped and in operation, respectively; and 
           [0034]      FIGS. 4 a  and 4 b    show a schematic view of a section of a propulsion unit in accordance with the present disclosure, when stopped and in operation, respectively. 
       
    
    
       [0035]    The drawings described herein are for illustration purposes only and are not intended to limit the scope of the present disclosure in any way. 
       DETAILED DESCRIPTION 
       [0036]    The following description is merely exemplary in nature and is not intended to limit the present disclosure, application, or uses. It should be understood that throughout the drawings, corresponding reference numerals indicate like or corresponding parts and features. 
         [0037]      FIG. 1  shows an exploded view of a propulsion unit  1 , including a bypass turbojet engine  2  and a nacelle  3 . 
         [0038]    The turbojet engine  2  includes a fan module, including a fan casing and an intermediate casing. The fan casing has a general cylindrical shape with a circular section, and surrounds the fan of the turbojet engine, whose rotation serves in particular to generate the secondary flow. The intermediate casing  21  is disposed downstream of the fan casing and in particular includes an outer annular casing defining an upstream portion of the flow path of cold flow, or secondary flow path, of the engine. The annular casing is linked to a hub of the intermediate casing by radial link arms. The link arms are generally four in number, located at the positions called “12h00,” “3h00,” “6h00” and “9h00” positions. 
         [0039]    The nacelle  3  includes an air inlet  31 , a median section, including in the example two fan cowls  32 , and a rear section, including in the example two substantially hemispherical half-portions  33 . Finally, the nacelle includes an ejection nozzle  34 . 
         [0040]      FIGS. 2 a  and 2 b    show a schematic section of a fan module of a conventional bypass turbojet engine, the section being located downstream of the fan blades, at the intermediate casing.  FIGS. 2 a  and 2 b    show the same section, respectively, when the engine is stopped and when it is in operation, in cruising speed.  FIGS. 2 a  and 2 b    thus show the intermediate casing  41 , and the walls  42 ,  43  of the engine located in the flow of cold air when the engine is in operation. These walls  42 ,  43  are disposed in an upstream portion of the secondary flow path  44 , intended to channel the secondary flow (or cold flow). These walls are constituted, for example, by the outer surfaces of some of the link arms between the hub of the engine and the intermediate casing, in particular the link arms located in the positions called “at 6h00” and “at 12h00.” In the example, these walls are located on either side of a vertical plane containing the longitudinal axis of the engine (the longitudinal axis of the engine being normal to the plane of the figures) 
         [0041]    The comparison of  FIGS. 2 a    (stopped engine) and  2   b  (engine in cruising speed) shows the consequences of the “fan twist” effect mentioned above. In  FIG. 2 b   , it can be seen that the walls  42 ,  43  have an angular offset relative to their position in  FIG. 2 a   . This angular offset is due to the deformation of the engine under the effect of the torsional forces induced by the rotation of the fan (whose direction of rotation in materialized in  FIG. 2 b    by the arrow F). This angular offset depends in particular on the speed of rotation of the fan and can reach values comprised between 2° and 10°, typically between 2° and 5° for small dimension engines, and between 5° and 10° for large dimension engines. 
         [0042]      FIGS. 3 a  and 3 b    show sections identical to those of  FIGS. 2 a  and 2 b   , but on which the portions of the inner fixed structure  45  (or IFS  45 ) of the nacelle located facing the walls  42 ,  43  have been shown. In the example, the inner fixed structure  45  includes two panels located on either side of a vertical plane containing the longitudinal axis of the engine (the longitudinal axis of the engine being normal to the plane of the Figures). A sealing gasket  46  interposed between the walls  42 ,  43  and the panels of the inner fixed structure  45 , is also shown. This sealing gasket  46  ensures the sealing between the upstream portion (delimited by the intermediate casing) and the downstream portion (delimited by the rear section of the nacelle) of the secondary flow path. The comparison of  FIGS. 3 a  and 3 b    shows the consequences of the “fan twist” effect on a conventional turbojet engine equipped with its nacelle. Indeed, it is seen in  FIG. 3 a    that the alignment between the portions of the inner fixed structure  45  facing the walls  42 ,  43  and the walls  42 ,  43  is correct when the engine is stopped. In these conditions, the gasket  46  has a substantially uniform crushing, and the aerodynamic qualities of the secondary flow path  44  are maximal. When the engine is in operation, the “fan twist” effect involves an angular deformation essentially on the engine: the rear section of the nacelle, and therefore the inner fixed structure  45 , is little or not subjected to the forces induced by the rotation of the fan blades. 
         [0043]    When the engine is in operation, it is therefore produced an angular offset between the walls  42 ,  43  and the inner fixed structure  45 , which results in a very uneven crushing of the gasket  46 . As observed in  FIG. 3 b   , the gasket  46  undergoes a very high compression in the areas A where the distance between the walls  42 ,  43  and the inner fixed structure decreases due to the angular deformation of the engine. Conversely, the gasket  46  undergoes a zero or very low compression in the areas B where the distance between the walls  42 ,  43  and the inner fixed structure increases. 
         [0044]    These compression forces on the gasket, generating very unevenly distributed crushing values, represent a major disadvantage because it is necessary to provide for a gasket capable of undergoing very large deformation amplitudes. Further, the gasket must be able to provide a satisfactory sealing over the entire deformation range. Thus, the gasket must have satisfactory sealing performances for deformation values comprised between 10% and 60%, while the optimum deformation of a gasket normally corresponds to a value of about 35%. The need to provide for a gasket that takes account of all these constraints impacts both the cost and the weight of the assembly. 
         [0045]    Otherwise, besides the disadvantages observed above, the angular offset between the walls  42 ,  43  of the engine and the inner fixed structure  45  also results in a great reduction in the aerodynamic qualities of the secondary flow path. Indeed, this angular offset results, at the interface between the walls  42 ,  43  and the inner fixed structure  45 , in a recess in the surface of the secondary flow path, and, consequently, hampers the aerodynamic continuity of the secondary flow path. It follows a great reduction of the aerodynamic qualities of the secondary flow path. 
         [0046]      FIGS. 4 a  and 4 b    show sections analogous to those of  FIGS. 3 a  and 3 b   , but carried out on a propulsion unit in accordance with the present disclosure as shown in  FIG. 1 .  FIGS. 4 a  and 4 b    thus show an intermediate casing  51 , delimiting a downstream portion of a secondary flow path  54 , intended to channel the secondary flow.  FIGS. 4 a  and 4 b    also show the portions of the inner fixed structure  55  (or IFS  55 ) of the rear section of the nacelle, as well as the gasket  56  interposed between the engine and the inner fixed structure  55 . In the example, the inner fixed structure includes two panels located on either side of a vertical plane containing the longitudinal axis of the engine. 
         [0047]    In accordance with the present disclosure, the inner fixed is shaped so that its alignment with the engine (and in particular with the walls  52 ,  53  of the engine) is optimal in the operating condition, and more particularly when the engine is in cruising speed. More precisely, when the engine is stopped, there is an angular offset about the longitudinal axis of the engine (axis corresponding in particular to the rotation axis of the fan and which is normal to the plane of  FIGS. 2 a , 2 b , 3 a , 3 b , 4 a , 4 b   ) between the panels of the inner fixed structure  55  and the walls  52 ,  53  of the engine, this offset being such that the torsional deformation of the engine in cruising speed, as observed in  FIG. 4 b   , will cancel this angular offset. Thus, the urging of the gasket  56  and the aerodynamic qualities of the secondary flow path  54  will be optimal in cruising speed. It is thus seen in  FIG. 4 b    that the deformation of the gasket  56  is very homogeneous and that there is no longer any recess between the walls  52 ,  53  and the panels of the inner fixed structure  55 . In contrast, the deformation of the gasket  56  is very uneven when the engine is stopped, which can be seen in  FIG. 4 a   . Thus, the urging of the gasket  56  and the aerodynamic qualities are (voluntarily) deteriorated when the engine is stopped. But this voluntary deterioration, on the one hand, has no particular disadvantage and, on the other hand, allows to improve the aerodynamic qualities of the secondary flow path  54  and the urging of the gasket  56  in flight conditions. This in particular allows a gain on the consumption of the engine and, further, allows to dimension the gasket in a less constraining manner. 
         [0048]    The initial angular offset (when the engine is stopped) will be, for example, comprised between 2° and 10°, and will be in particular comprised between 2° and 5° for a small size engine, and comprised between 5° and 10° for large size engines. 
         [0049]    Although the present disclosure has been described in relation with particular forms, it is in no way limited thereto and that it comprises all the technical equivalents of the means described as well as their combinations if the latter are within the scope of the present disclosure. 
         [0050]    The description of the disclosure is merely exemplary in nature and, thus, variations that do not depart from the substance of the disclosure are intended to be within the scope of the disclosure. Such variations are not to be regarded as a departure from the spirit and scope of the disclosure.