• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to an aircraft propulsion assembly comprising a first engine comprising a first nozzle (17) and a second engine comprising a second nozzle (18), and a nacelle (1). The first nozzle (17) and the second nozzle (18) each have an outlet section, defined by an outer wall of the nozzle, a portion of which is substantially straight, and the straight portion of the outlet section of the first nozzle ( 17) and the right portion of the outlet section of the second nozzle (18) are in contact with one another or form a common wall of said nozzles (17, 18) at a center plane (PM) of symmetry of the propulsion unit. The great proximity allowed between the engines makes it possible to limit the space between the fairing of each engine formed by the nacelle, and to limit the aerodynamic drag of the propulsion unit.
    公开号:FR3079211A1
    申请号:FR1852527
    申请日:2018-03-23
    公开日:2019-09-27
    发明作者:Jean-Michel ROGERO；Camil Negulescu
    申请人:Airbus Operations SAS；
    IPC主号:
    专利说明:

    The present invention relates to the field of aircraft propulsion systems. It relates more particularly to the architecture of propulsion units comprising two adjacent engines, that is to say positioned side by side close to one another.
    Propulsion systems of aircraft comprising two adjacent engines have been used in particular on certain commercial aircraft, such as the model “VC-10” of the aircraft manufacturer Vickers of the British Aircraft Corporation and the model “IL-62” of the aircraft manufacturer llyushin. The propulsion units with adjacent engines known in the state of the art are nevertheless generally designed as the juxtaposition of two propulsion units. The integration between the engines, particularly with regard to the nacelle that hulls them, is weak. In other words, the nacelle of these assemblies is designed almost like two nacelles of propulsion assemblies with a single isolated engine, joined and faired together at the level of the inter-nacelle zone.
    According to such a design of the propulsion units with adjacent engines, a voluminous fairing must be provided between the engines. This fairing must have, for aerodynamic reasons, a rear appendage called "beaver tail". It follows from this design according to the state of the art of propulsion units with adjacent engines that their nacelle has a large wetted surface, which increases the friction drag generated by the propulsion unit. The wetted surface corresponds to the surface in contact with the outside air flow.
    In addition, the imposing fairing between the engines and the interactions between the fairing of each of the engines generates overspeeds in the air flow flowing between the engines, which causes variations in air density and the generation of '' a compressibility trail, or even shock waves in the inter-motor area.
    Thus, the aerodynamic losses, linked to the drag of friction or compressibility, are significant at the level of the space separating the nozzles of the two engines of the propulsion unit.
    The invention thus tends to propose an architecture of propulsion unit with adjacent engines optimized so as to limit aerodynamic losses.
    Thus, the invention relates to an aircraft propulsion unit comprising a first engine and a second engine adjacent and a nacelle in which said first and second engines are installed, the first engine comprising a first gas ejection nozzle and the second engine comprising a second gas ejection nozzle. The first nozzle and the second nozzle each have an outlet section, defined by an external wall of the nozzle, a portion of which is substantially straight. The propulsion unit is configured so that the straight portion of the outlet section of the first nozzle and the straight portion of the outlet section of the second nozzle are opposite one another on either side of a median plane (PM) of symmetry of the propulsion unit, in contact with one another, or form a common wall of said nozzles (17,18) at the level of said median plane (PM) of symmetry of the propulsion system.
    The use of gas ejection nozzles having a non-circular outlet section, and in particular having a substantially straight portion allows their approximation in order to limit or eliminate the space between the nozzles which requires in the prior art l use of a large fairing to avoid detachment of the air passing through this space. In particular, the walls forming the first nozzle and the second nozzle are brought closer to the median plane of symmetry of the propulsion unit so as to become very close to each other, and common or in contact at the outlet of the nozzles. This goes hand in hand with a close proximity between the engines of the propulsion unit. The close proximity between the motors makes it possible to limit the space between the fairing of each motor formed by the nacelle.
    The outlet section of each nozzle may comprise a portion substantially in a semicircle, opposite to said median plane (PM) of symmetry of the propulsion unit.
    Each nozzle may comprise a cone extending longitudinally inside said nozzle and through its outlet section, said cone being shaped so that at the outlet section of the nozzle, the distance between said cone and the outer wall of the nozzle is substantially constant over the entire periphery of said cone.
    The respective outlet sections of the first and second nozzles can be asymmetrical with respect to the plane orthogonal to the median plane of symmetry of the propulsion unit passing through a main axis of the first engine and through a main axis of the second engine.
    The nozzle outlet sections can be oriented convergently towards the median plane.
    In an aircraft propulsion unit comprising a wall common to the nozzles, said common wall can be stopped upstream of the respective external walls of said nozzles, so as to form an increased outlet section adapted to maximize the thrust during the operation of a single said first motor and second motor.
    The nacelle of the propulsion unit may include a common air intake lip for the first engine and the second engine, the air flow being distributed between the first engine and the second engine by a central lip which extends, at less in part, set back from said common air intake lip. The common air intake lip may include, below the middle lip, a lower lobe and an upper lobe extending forward of the nacelle. The middle lip may have a curvature in the median plane (PM) so that it joins said lower lobe (15) and upper lobe (14) by substantially forming an arc of a circle.
    An exterior surface of the propulsion unit nacelle can form a single aerodynamic surface common to both engines.
    The aerodynamic surface can be extended beyond the outlet section of the nozzles, at the median plane, by a local fairing.
    The first motor may comprise a first fan, and the second motor may comprise a second fan, the first fan and the second fan extending in the same plane, and in which the circle in which the first rotating fan is inscribed is distant at most thirty centimeters, and preferably at most twenty centimeters, from the circle in which the second rotating fan fits.
    The invention also relates to an aircraft comprising an aircraft propulsion unit as previously described. Said assembly can in particular be installed either at the level of a double point which comprises an aircraft fuselage, or under wing, or in the lateral position at the level of a rear point of an aircraft fuselage.
    Other features and advantages of the invention will appear in the description below.
    In the appended drawings, given by way of nonlimiting examples:
    - Figure 1 shows in a schematic three-dimensional view an aircraft propulsion unit with adjacent engines as known in the prior art;
    - Figure 2 shows in a front view in principle, an aircraft propulsion unit according to one embodiment of the invention;
    - Figure 3 shows, in a schematic view in three dimensions similar to that of Figure 1, an aircraft propulsion assembly according to an embodiment of the invention;
    - Figure 4 shows a partial schematic sectional view of the front part of the aircraft propulsion unit of Figure 3;
    - Figure 5 shows in another schematic sectional view of the propulsion unit of Figure 3;
    - Figure 6 shows in a partial schematic view in three dimensions the rear part of a propulsion unit according to an embodiment of the invention;
    - Figure 7 shows in a principle view the outlet section of the nozzles of a propulsion unit according to an embodiment of the invention;
    - Figures 8a and 8b show, in a sectional view the rear parts of two alternative embodiments of a propulsion unit according to the invention;
    - Figures 9a and 9b show, in a sectional view, the rear parts of two alternative embodiments of a propulsion unit according to the invention;
    - Figure 10 shows in a schematic three-dimensional view the rear part of an aircraft having a configuration particularly suitable for receiving a propulsion unit according to the invention.
    In a propulsion unit with adjacent engines as known in the state of the art, a first engine and a second engine are juxtaposed to each other in a so-called transverse direction (y). The longitudinal direction (x) is defined according to the general direction of extension of the engines, which also corresponds to the general direction of the gas flow in the engines and in the nacelle which surrounds them.
    Finally, a third direction, perpendicular to the longitudinal direction (x) and to the transverse direction (y), is called vertical direction (z) insofar as when the propulsion unit is installed on an aircraft having an angle of incidence zero, the longitudinal (x) and transverse (y) directions define a substantially horizontal plane.
    Similarly, in this document, the concepts of "upstream" and "downstream", or "front" and "rear" are understood according to the flow of the gas flow in the propulsion unit .
    The first motor and the second motor are installed in a nacelle 1. The nacelle 1 is designed in the state of the art as two isolated engine nacelles, juxtaposed and linked together. Thus, the first engine comprises a first aerodynamic fairing 2, provided with a first air inlet 3 defined by a circular lip (or having an edge according to a closed curve). The second engine has a second aerodynamic fairing 4, provided with a second air inlet 5 defined by a lip identical to that of the first fairing 2.
    The first fairing 2 and the second fairing 4 are treated, as regards their design, in particular in aerodynamic matter, as nacelle fairings for propulsion units with a single engine.
    However, the first aerodynamic fairing 2 is linked to the second aerodynamic fairing 4 so as to form a single nacelle which receives the two engines. The nacelle thus formed constitutes a single fairing (comprising the first and the second fairing) which has a large junction surface between the first fairing 2 and the second fairing 4.
    This results in a bulky nacelle, and leads to significant aerodynamic interactions between the first fairing 2 and the second fairing 4.
    The first engine has a first gas ejection nozzle, and the second engine has a second gas ejection nozzle. Between the exit of the first nozzle and the exit of the second nozzle, it may be necessary to add an aerodynamic appendage, called a beaver tail 6, which guides the flow leaving the nozzles as well as the air which bypasses the nacelle in the faired area between the motors
    This also increases the volume and therefore the wet surface of the nacelle. The large wetted surface of the nacelle generates a significant drag of friction, and the imperfect treatment of the interactions between the first fairing 2 and the second fairing 4, and more generally air flows around the nacelle which can present overspeeds, causes a compressibility trail. The aerodynamic losses linked to friction drag and compressibility generate excess fuel consumption which should be limited as much as possible.
    FIG. 2 represents an aircraft propulsion unit in accordance with an embodiment of the invention. Figure 2 is a front view in principle, that is to say illustrating the air inlet in the nacelle in the foreground, and intended to illustrate the approximation between the first motor and the second motor can be put in work in the invention. The approximation that can be made between the motors is generally dictated by the dimension of their larger element, in a vertical transverse plane, namely the fan. The first motor 7 (here represented by its fan cone) comprises a first fan 8 in rotation along a first main axis A1 of the first motor 7, the second motor 9 (here represented by its fan cone) comprises a second fan 10 in rotation along a second main axis A2 of the second motor 9. Each fan has a set of blades whose rotational movement is inscribed in a circle. In the invention, the distance d between the circle in which the first rotating fan is inscribed and the circle in which the second rotating fan is inscribed can be reduced to less than thirty centimeters, for example of the order of twenty centimeters.
    This great proximity is made possible by certain characteristics developed in the invention and which will be detailed below, and, conversely, allows their implementation. It allows improved integration of the first fairing 2 formed around the first motor 7 and the second fairing 4 formed around the second motor 9.
    FIG. 3 represents an aircraft propulsion unit in accordance with an embodiment of the invention.
    In the exemplary embodiment shown in FIG. 3, the first motor 7 and the second motor 9 have a common air inlet, formed by a common air inlet lip 11.
    Compared to a propulsion unit with adjacent engines as known in the state of the art, the approximation between the first engine 7 and the second engine 9 operated in this embodiment of the invention reduces the space between the first fairing 2 of the first motor 7 and the fairing 4 of the second motor 9. However, there is still a risk of overspeeds in the middle zone of the nacelle, or upper inter-fairing zone 12 and lower inter-fairing zone 13.
    The concepts of "upper" and "lower" are understood in the so-called vertical direction z. The upper part of the propulsion unit is thus that which is located upwards when the propulsion unit is installed on an aircraft.
    The inter-fairing area must be properly careened, to avoid the generation of overspeeds. In particular, the aerodynamic fairing of this zone must not be lowered (that is to say approached the plane of the main axes A1, A2 of the engines) in such a way that the volume between the first fairing 2 and the second fairing 4 becomes a constrained volume which could result in the generation of overspeeds and shock waves causing aerodynamic losses.
    However, too high a fairing (far from the plane of the main axes of the engines) is also likely to generate overspeeds in the air flow bypassing the nacelle in this area, due to the large thickness of the nacelle. In order to be able to provide an aerodynamic shape, upper and lower, raised between the first and second fairings 2,4 without risk of generating overspeeds in the air flow, the air intake of the air intake lip joint 11 comprises an upper lobe 14 which extends, towards the front of the nacelle, the fairing of the upper inter-fairing zone 12. The common air inlet lip 11 comprises a lower lobe 15 which extends, towards the before the nacelle, the fairing of the lower inter-fairing zone 13. These upper and lower lobes 14, 15 smooth the flow of air outside the nacelle 1, at the level of the inter-fairing zones 11, 12.
    As a result, the nacelle forms a single aerodynamic surface common to the two engines. This aerodynamic surface does not have a sudden break in its shape: the shapes of the nacelle are rounded, and have radii of curvature as wide as possible; the nacelle is for example devoid of concave edges in the inter-fairing zone.
    In order to separate the air flow entering the common air inlet formed by the common air inlet lip 11, a central lip 16 is formed in a vertical median plane of the propulsion unit. Thus, the upper and lower lobes 14,15 are formed plumb with the middle lip 16.
    The middle lip 16 extends back from the common lip 11, for example as illustrated in FIGS. 4 and 5.
    FIG. 4 is a view of the front part of the propulsion unit of FIG. 3, in section along a horizontal cutting plane P1 (parallel to the longitudinal direction x and to the transverse direction y) passing through the main axes A1, A2 motors. Figure 5 is a view of the propulsion unit of Figure 3, in section along a vertical section plane P2, parallel to the vertical direction z and the longitudinal direction x, passing through the first main axis A1 of the first engine 7 .
    Due to the remote position of the middle lip 16 and its shape, the air flow to the first motor 7 and the second motor 9 is separated downstream from the common air inlet. The formation of the middle lip 11 is made possible by the close proximity between the first motor 7 and the second motor 9. In a vertical plane, the middle lip 16 has a shape curved towards the inside of the nacelle 1. In particular, the middle lip joins the upper lobe 14 and the lower lobe 15 substantially in an arc.
    Local overspeed zones on the middle lip 16 may be located near the area of passage of the blades of the first fan 8 and of the second fan 10, that is to say in the part of the middle lip 16 la further back in the nacelle 1. The aerodynamic profile of the middle lip can be adapted to avoid these overspeed phenomena.
    It is notable that this risk of overspeed in the air flow absorbed in the nacelle does not exist in the context of a propellant assembly with ingestion of boundary layer, as is the case for example in the example shown in Figure 10, which will be detailed below. Indeed, the boundary layer ingested by the propellant assembly has a low flow speed compared to the rest of the flow.
    FIG. 6 schematically shows the rear part of a propulsion unit according to an embodiment of the invention. In particular, FIG. 6 illustrates the first nozzle 17 of the first engine 7 and the second nozzle 18 of the second engine 9. The invention tends to reduce or eliminate the space between the nozzles of the engines of the propulsion unit. Indeed, this space, which is located in the rear part of the upper inter-fairing zone 12 and in the rear part of the lower inter-fairing zone 13, requires in the prior art the implementation of a bulky fairing to allow air passing through the inter-fairing space to bypass the nacelle without detachment. This fairing has a large wetted surface and thus causes aerodynamic losses by friction and compressibility.
    To limit the size of this fairing, the outer lines of the first nozzle 17 and the second nozzle 18 adjacent are approached and brought into contact as much as possible.
    The nozzles 17, 18 are approached from the vertical median plane PM of the propulsion unit, so that their outer lines become, in the proximity zone, common, adjacent or at least very close. As a result, the first nozzle 17 and the second nozzle 18 have an external wall, which defines the respective outlet section of the nozzle, a portion of which is substantially straight. This cross section is parallel to and near the midplane PM. Thus, the straight portion of the outlet section of the first nozzle is arranged opposite the straight portion of the outlet section of the second nozzle, on either side of the median plane PM of symmetry of the propulsion unit. The nozzles 17, 18 thus have common external lines, adjacent or at least very close, at least at the level of this cross section. The nozzles 17, 18 are thus in contact with one another at the right portion level of their wall, or these straight portions form a common wall at said nozzles.
    A separation of the nozzles 17, 18, for example by a wall common to the two nozzles, is necessary in order to avoid performance reductions due to too large an increase in effective nozzle section for an engine in operation, in the case of an operation stop of the other engine of the propulsion unit with adjacent engines. However, in one embodiment, this separation (for example this common wall) can be stopped slightly, that is to say a few centimeters, upstream of the outlet of the nozzles, at their respective outer wall, so forming an increased and optimized outlet section to maximize the thrust of the engine remaining in operation in the event that the other engine is not in operation.
    The length (or height, this dimension being oriented in the vertical direction z) of the straight portion of the wall of the two nozzles 17, 18 is determined by the approximation made between said nozzles, but also by the angle of shrinkage of the surface external inter-nacelle. The narrowing angle is defined by the extreme orientation of the nacelle with respect to the main axes A1, A2 of the motors respectively for the first fairing 2 and the second fairing
    4. A large angle of constriction reflects a pronounced curvature at the rear of the nacelle, which can cause detachments of the air flow around the nacelle and aerodynamic disturbances generating aerodynamic losses.
    The length of the straight part of the wall of the nozzles is thus determined in order to obtain an acceptable narrowing angle in the inter-nacelle outer zone (upper and lower).
    FIG. 7 illustrates the outlet section of the nozzles 17, 18 of a propulsion unit according to an embodiment of the invention, compared with the outlet section of the nozzle according to the state of the art. The first nozzle 17 includes a first outlet cone 19. The second nozzle 18 includes a second outlet cone 20. Each cone 19, 20 extends longitudinally in the nozzle which it equips, and extends through its outlet section. In FIG. 7, the outlet section of the nozzles 17, 18 and the outline of the cones 19, 20 according to the invention are shown in solid lines, while the outlet section of the nozzles conforms to the prior art and the contour of conventional cones are shown in dotted lines.
    The outlet surface of the nozzles of the prior art being optimized to improve the thrust and the efficiency of the engines, there is a tendency to maintain the value of this area unchanged. On the one hand, the lateral external lines (wall of the nozzles furthest from the median plane PM) of the nozzles can be moved (compared to the state of the art) towards said median plane PM. This approximation is carried out in a balanced manner around the circumference of the nozzle so as not to increase locally, significantly, the angle of constriction of the aerodynamic external lines of the nacelles. On the other hand, the geometry of the cone 19, 20 of each nozzle can be modified (compared to the prior art) to be adapted to the section of the nozzles proposed in the invention. For example, in particular in the plane of the outlet section of the nozzles, the shape of each cone 19, 20 can be adapted so that the distance between said cone and the outer wall of the nozzle is substantially constant. The distance between the cone and the outer wall of the nozzle corresponds to the shortest distance separating the cone at a point considered, and said inner wall of the nozzle.
    In other words, the cone 19, 20 is modified compared to the prior art in a similar manner to the nozzle in which it extends.
    It will be noted that the straight portion of the outer wall of the nozzles is arranged vertically between a lower point h1 and an upper point h2 of the median plane PM. In the example shown here, h1 and h2 are positioned equidistant from the plane P1 passing through the main axes A1, A2 of the motors 7,
    9. Nevertheless, the points h1 and h2 can be arranged at different distances from said plane P1 passing through the main axes A1, A2, so that the nozzles are asymmetrical with respect to the plane P1. This asymmetry can be introduced to increase the resistance of the nacelle to the phenomenon of detachment of the air flow when said nacelle has a positive impact. This is the case in a majority of flight phases of an aircraft equipped with the propulsion unit.
    Finally, in order to avoid a risk of local detachment of the air flow bypassing the nacelle in the inter-fairing zone, a small local fairing 21 can be provided between the nozzles. This small local fairing 21 can in particular extend towards the rear of the inter-fairing zone, beyond the outlet of the nozzles. This small local fairing 21 has a very limited volume, without common measure with a beaver tail 6 implemented in the state of the art.
    The modification of the section of the nozzles 17, 18, and in particular the offset of the walls of the nozzles furthest from the median plane PM towards said median plane PM, requires a modification of the geometry of the rear of the nacelle to keep the surface unchanged nozzle outlet.
    Figure 8a shows, in a sectional view along the first section plane P1, the rear part of a propulsion unit according to an embodiment of the invention. In FIG. 8a, the rear profile of the nacelle and the cones known in the state of the art are shown in dotted lines, that is to say as they would appear if the motors were simply brought together without modifying the nozzle configuration. In the embodiment of FIG. 8a, the walls respectively forming the outlet of the first nozzle 17 and the second nozzle 18 are brought together at the level of the median plane PM, without being shared. The cones 19, 20, in addition to their modification of shape, are offset towards the median plane PM. In order to maintain the value of the area of the outlet section of the nozzles, the necking angle a1 is increased compared to the reference necking angle αθ of the state of the art.
    FIG. 8b corresponds to an alternative embodiment in which a single wall forms the outlet of the first and of the second nozzle, at the level of the median plane PM. This further shifts the cone of each nozzle towards said midplane PM. In this configuration, the neck angle a1 is increased as much as possible.
    In the case where the reference constriction angle αθ is already at a maximum admissible value, the modification of the configuration of the section of the nozzles may consist in preserving (compared to the state of the art) the position of the external wall. of the semicircular nozzle which is distant from the median plane PM, and to consequently increase the volume of the cone installed in the nozzle.
    Figures 9a and 9b show, in a view similar to that of Figures 8a and 8b, two alternatives that can be used to keep the outlet surface of the nozzles unchanged compared to the prior art, in the case where the angle of reference constraint αθ is already at a maximum admissible value.
    As in FIGS. 8a and 8b, the rear profile of the nacelle and the cones known in the state of the art, that is to say as they would appear if one were shown by comparison in dotted lines simply brought the motors closer without changing the configuration of the nozzles.
    To respond to this problem, it is proposed in FIG. 9a to increase the length of the nacelle towards the rear, by a length I, uniformly in the longitudinal direction x, while maintaining the reference narrowing angle αθ.
    As an alternative, it is proposed in FIG. 9b to extend the nacelle only on the lateral external parts of the nacelles. This results in nozzles with oblique nozzle outlet sections, oriented convergently towards the median plane PM of vertical symmetry of the two engines. The increase in the area of the nacelle related to its elongation is however limited compared to the alternative of FIG. 9a.
    In these two alternatives, presented respectively in FIGS. 9a and 9b, the surface of the nozzles is increased by the increase in their length, but this increase in surface is less than the surface of the fairing necessary in the prior art for fill the space between circular nozzles. This results in a significant gain in terms of aerodynamic friction losses.
    The invention thus provides an aircraft propulsion assembly comprising a nacelle, an air intake and very highly integrated nozzles to reduce aerodynamic drag and thus the fuel consumption of the aircraft equipped with the propulsion assembly.
    In particular, the invention proposes for propulsion units with adjacent engines a fully integrated aerodynamic nacelle which can have:
    - an aerodynamic surface common to the two engines, with an optimized inter-fairing zone, for example by the presence of an upper lobe and a lower lobe,
    - a common air intake for both engines with a common rear middle lip, and
    - non-circular nozzles, close to or in contact laterally in the median plane of the propulsion unit.
    The above characteristics can be applied independently of each other, in particular depending on the proximity of the two motors and their architecture. For a propulsion unit having these three characteristics, the gain in fuel consumption linked to the reduction in aerodynamic losses is of the order of two to four percent compared to a propulsion unit which would be constituted according to the state of the art.
    A propulsion unit according to the invention is very particularly suitable for use on an aircraft with a converging double rear tip with adjacent engines, the rear part of which is shown in FIG. 10. In such an aircraft, the fuselage 22 divides at its rear end at a first point 23 and a second point 24. Each point is associated with a motor: the first point 23 is associated with the first motor 7 of the propulsion unit, the second point 24 is associated with the second motor 9. The flow of air entering each engine respectively then comprises the boundary layer formed on the surface of the first tip 23 and the second tip 24. This can improve the efficiency of the engines.
    A propulsion unit according to the invention can nevertheless be used successfully under the wing of an aircraft or in a lateral position at the rear tip of an aircraft fuselage. Other locations of the propulsion unit are possible, without departing from the scope of the invention. The description of an embodiment of the invention carried out above relates to a horizontal arrangement of the adjacent motors, that is to say with adjacent motors having their main axes in the plane comprising the longitudinal directions (x ) and transverse (y). Nevertheless, a relative vertical arrangement of the adjacent motors, or a disposition of the adjacent motors in any other plane is not excluded from the invention.
    Likewise, the above description has been made on the basis of motors having their main axes oriented in the longitudinal direction (x). For practical reasons, the engines may however be set with a few degrees of angle on the yaw axis and on the pitch axis of the aircraft which they equip, and these angles may be different for the two adjacent engines. Likewise, the adjacent motors may have a slightly different position in the longitudinal direction (x), or for another configuration, be slightly offset from each other in the general direction of their main shafts. These configurations are not excluded from the invention.
    Obviously, a propulsion unit according to the invention may include more than two adjacent engines.
    权利要求:
    Claims (14)
    [1" id="c-fr-0001]
    1. Aircraft propulsion assembly comprising an adjacent first engine (7) and a second engine (9) and a nacelle (1) in which said first and second engines (7, 9) are installed, the first engine (7) comprising a first gas ejection nozzle (17) and the second motor (9) comprising a second gas ejection nozzle (18), characterized in that the first nozzle (17) and the second nozzle (18) each have an outlet section, defined by an external wall of the nozzle, a portion of which is substantially straight, the propulsion unit being configured so that the straight portion of the outlet section of the first nozzle (17) and the straight portion of the outlet section of the second nozzle (18) are opposite one another on either side of a median plane (PM) of symmetry of the propulsion unit, in contact with one of the 'other, or form a common wall of said nozzles (17,18) at the level of said plane dian (PM) of symmetry of the propulsion unit.
    [2" id="c-fr-0002]
    2. aircraft propulsion unit according to claim 1, wherein the outlet section of each nozzle (17,18) comprises a portion substantially in a semicircle, opposite said median plane (PM) of symmetry of the propulsion system.
    [3" id="c-fr-0003]
    3. aircraft propulsion unit according to claim 1 or claim 2, wherein each nozzle (17,18) comprises a cone (19,20) extending longitudinally inside said nozzle (17,18) and through its outlet section, said cone (19,20) being shaped so that at the outlet section of the nozzle, the distance between said cone (19,20) and the external wall of the nozzle (17 , 18) is substantially constant over the entire periphery of said cone (19,20).
    [4" id="c-fr-0004]
    4. aircraft propulsion unit according to one of the preceding claims, in which the respective outlet sections of the first and of the second nozzle (17, 18) are asymmetrical with respect to the plane (P1) orthogonal to the median plane (PM ) of symmetry of the propulsion unit passing through a main axis of the first engine and through a main axis of the second engine.
    [5" id="c-fr-0005]
    5. Aircraft propulsion unit according to one of the preceding claims, in which the nozzle outlet sections (17,18) are oriented convergently towards the median plane (PM).
    [6" id="c-fr-0006]
    6. aircraft propulsion unit according to one of the preceding claims, comprising a wall common to the nozzles (17,18) in which said common wall is stopped upstream of the respective outer walls of said nozzles, so as to form an outlet section increased adapted to maximize the thrust during the operation of only one of said first motor and second motor.
    [7" id="c-fr-0007]
    7. aircraft propulsion unit according to one of the preceding claims, in which the nacelle (1) comprises a common air intake lip (11) for the first engine and the second engine, the air flow being distributed between the first motor and the second motor by a central lip (16) which extends, at least in part, back from said common air inlet lip (11).
    [8" id="c-fr-0008]
    8. aircraft propulsion unit according to claim 7, in which the common air intake lip (11) comprises, below the middle lip, a lower lobe (15) and an upper lobe (14) extending towards the front of the nacelle (1).
    [9" id="c-fr-0009]
    9. aircraft propulsion unit according to claim 8, wherein the middle lip (16) has a curvature in the median plane (PM) so that it joins said lower lobe (15) and upper lobe (14) forming substantially an arc.
    [10" id="c-fr-0010]
    10. An aircraft propulsion unit according to one of the preceding claims, in which an exterior surface of the nacelle (1) forms a single aerodynamic surface common to the two engines (7, 8).
    [11" id="c-fr-0011]
    11. aircraft propulsion unit according to claim 10, in which the aerodynamic surface is extended beyond the outlet section of the nozzles (17,18), at the level of the median plane PM, by a local fairing (21).
    [12" id="c-fr-0012]
    12. Aircraft propulsion unit according to one of the preceding claims, in which the first engine (7) has a first fan (8), and the second engine (9) has a second fan (10), the first fan ( 8) and the second blower (10) extending in the same plane, and in which the circle in which the first blower (8) in rotation is inscribed is distant at most thirty centimeters, and preferably at least plus twenty centimeters, from the circle in which the second rotating fan (10) is inscribed.
    [13" id="c-fr-0013]
    13. Aircraft comprising an aircraft propulsion unit according to one of claims 1 to 12.
    [14" id="c-fr-0014]
    14. The aircraft as claimed in claim 13, in which said assembly is installed either at the level of a double point which comprises a fuselage of the aircraft, or under wing, or in the lateral position at the level of a rear point of a fuselage aircraft.
    类似技术:
    公开号 | 公开日 | 专利标题
    CA2743009C|2017-03-14|Air intake for an aeroplane engine with unducted propellers
    CA2608373C|2013-02-19|Method of reducing noise emissions at the rear of a turboshaft engine and turboshaft engine thus improved
    FR3037560A1|2016-12-23|AIRCRAFT WING INCLUDING A PILOTABLE WING FIT IN INCIDENCE
    FR2938502A1|2010-05-21|TURBOMACHINE COMPRISING A NON-CARNEY PROPELLER EQUIPPED WITH AIR GUIDING MEANS
    FR3079211A1|2019-09-27|PROPELLANT AIRCRAFT ASSEMBLY COMPRISING TWO ADJACENT ENGINES, WHOSE OUTLETS HOLES HAVE A RIGHT PORTION CLOSE TO A MEDIAN PLAN OF THE PROPULSIVE ASSEMBLY
    FR3050721A1|2017-11-03|AIRCRAFT ENGINE ASSEMBLY COMPRISING A MATTRESS ATTACK EDGE INTEGRATED WITH AN ANNULAR ROW OF OUTER CARRIER OUTPUT GUIDELINES
    CA2872207A1|2015-06-05|Air ejection device including an aerodynamic profile equipped with a flexible strip for sealing the slit
    CA2850243C|2018-09-25|Blade for a fan of a turbomachine, notably of the unducted fan type, corresponding fan and corresponding turbomachine
    FR3044295A1|2017-06-02|DEVICE FORMING A LEFT EDGE OF AERODYNAMIC PROFILE AND COMPRISING A BLOWING SYSTEM
    EP2279341B1|2013-10-23|Device for reducing noise generated by an aircraft jet engine with curved ducts
    FR2899201A1|2007-10-05|Aircraft`s e.g. twin-engine aircraft, wing arrangement, has zone including projection delimiting channel, where air flow is accelerated through channel and is evacuated in direction privileged under wing, between surface of wing and nacelle
    EP2191124B1|2015-10-07|Gas exhaust cone for an aircraft turbojet, corresponding turbojet and engine plant
    EP3380399B1|2021-04-28|Aircraft powered by a turbomachine provided with an acoustic baffle
    FR2921977A1|2009-04-10|DOUBLE FLOW TURBOMOTEUR FOR AIRCRAFT
    EP3540205A1|2019-09-18|Aircraft drive unit in which the nacelle is linked by a pivot to the drive shaft of its fan
    FR3052191B1|2019-07-12|INVERSION OF PUSH IN A TURBOMACHINE WITH VARIABLE CALIBRATION BLOWER
    FR2680831A1|1993-03-05|PROCESS FOR ESTABLISHING THE ENTRY PROFILE OF THE NACELLE OF AN AIRPLANE GAS TURBINE ENGINE AND INTAKE OF THE NACELLE AS WELL AS NACELLE OBTAINED USING THIS PROCESS.
    WO2016132073A1|2016-08-25|Aircraft propulsion unit comprising an unducted-fan turbine engine and an attachment pylon
    WO2021074535A1|2021-04-22|Assembly for turbomachine
    FR3021706A1|2015-12-04|AIRCRAFT TURBOPROPULSOR COMPRISING TWO COAXIAL PROPELLERS.
    EP3778383B1|2021-10-06|Forward section of nacelle of an aircraft propulsion assembly comprising a thermal transition zone
    FR2993861A1|2014-01-31|Multi-flow turbojet engine assembly for aircraft, has nacelle comprising upstream air inlet sleeve with air intake in region of leading edge, where shield of pylon is extended up to air intake of air inlet sleeve
    FR3105553A1|2021-06-25|Acoustic treatment system with at least two degrees of freedom comprising a quarter-wave coating allowing the passage of acoustic waves in a cavity-mode resonator
    WO2019243117A1|2019-12-26|Rear propulsion system for an aircraft
    WO2012120227A1|2012-09-13|Pipe having an upstream core having a sharp curvature
    同族专利:
    公开号 | 公开日
    FR3079211B1|2020-11-20|
    US20190291885A1|2019-09-26|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US3060685A|1959-09-17|1962-10-30|Hamburger Flugzeugbau Gmbh|Multiple engine jet-propulsion drive and thrust reverser for aircraft|
    GB1030521A|1964-02-20|1966-05-25|Rolls Royce|Mounting arrangement for gas turbine engines in aircraft|
    DE2016805A1|1969-04-09|1970-10-15|British Aircraft Corp., Ltd., London|plane|
    US5114097A|1991-04-29|1992-05-19|Williams International Corporation|Aircraft|
    US20020096598A1|2001-01-19|2002-07-25|Nelson Chester P.|Integrated and/or modular high-speed aircraft|
    US20080245925A1|2007-01-09|2008-10-09|Rolls-Royce Plc|Aircraft configuration|
    US8402740B2|2008-02-28|2013-03-26|Rolls-Royce Deutschland Ltd & Co Kg|Aircraft propulsion unit in multi-fan design|
    DE102008024463A1|2008-05-21|2009-12-03|Bauhaus Luftfahrt E.V.|Aircraft propulsion system, has turboshaft engine positioned inside fuselage, and fans directly arranged at fuselage surface and in segmental arch of fuselage outer contour on upper side of aircraft|
    US20170081036A1|2015-09-21|2017-03-23|General Electric Company|Non-axis symmetric aft engine|
    US10589869B2|2018-07-25|2020-03-17|General Electric Company|Nacelle inlet lip fuse structure|
    CN111824431B|2020-07-10|2021-10-26|南京航空航天大学|High-speed air inlet precursor based on integrally-controllable ridge type pressure distribution|
    法律状态:
    2019-03-22| PLFP| Fee payment|Year of fee payment: 2 |
    2019-09-27| PLSC| Search report ready|Effective date: 20190927 |
    2020-03-19| PLFP| Fee payment|Year of fee payment: 3 |
    2021-03-23| PLFP| Fee payment|Year of fee payment: 4 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1852527A|FR3079211B1|2018-03-23|2018-03-23|AIRCRAFT PROPULSION ASSEMBLY COMPRISING TWO ADJACENT ENGINES, OF WHICH THE OUTLET PIPES PRESENT A RIGHT PORTION NEAR A MEDIAN PLANE OF THE PROPULSION ASSEMBLY|
    FR1852527|2018-03-23|FR1852527A| FR3079211B1|2018-03-23|2018-03-23|AIRCRAFT PROPULSION ASSEMBLY COMPRISING TWO ADJACENT ENGINES, OF WHICH THE OUTLET PIPES PRESENT A RIGHT PORTION NEAR A MEDIAN PLANE OF THE PROPULSION ASSEMBLY|
    US16/355,999| US20190291885A1|2018-03-23|2019-03-18|Aircraft propulsion assembly comprising two adjacent engines, of which the outlet nozzles have a straight portion in the vicinity of a median plane of the propulsion assembly|
    [返回顶部]