Afterbody section of a turbojet engine having an increased a9 / a8 nozzle area ratio

Abstract

A rear portion of a turbojet engine with converging-diverging nozzles includes a device for controlling the position of the diverging vane relative to the position of the converging vane, the device including a first link (120) having a first end (120A) hinged to the converging vane (30) and an opposite second end (120B), a second link (122) having a first end (122A) hinged to the diverging vane (32) and an opposite second end (122B), and a third link (124) having a first end (124A) hinged to a synchronization ring (82) and an opposite second end (124B). The second end (120B) of the first link is hinged to the second end (122B) of the second link and / or to the second end (124B) of the third link, and the second end (124B) of the third link (124) is hinged to the second end (122B) of the second link (122).

Description

Technical Field

[0001] This invention relates to the field of nozzles with variable geometry intended for use in turbojet engines to propel aircraft. Background Technology

[0002] Turbojet engines designed for supersonic flight typically include an afterburner passage with an outlet defined by an axisymmetric nozzle of variable geometry, meaning the nozzle geometry can be adapted to different speeds at which such an aircraft may fly.

[0003] To this end, such a nozzle includes at least one set of movable internal vanes called convergent vanes, distributed around the longitudinal axis of the turbojet engine, and each convergent vane has an upstream end hinged to an internal structure of the housing, and each convergent vane includes a panel designed to guide exhaust gas flow within the nozzle. Such a nozzle further includes a system for controlling the convergent vanes, which enables the vanes to pivot synchronously around their hinge axis to the housing.

[0004] Typically, nozzles designed for supersonic flight further include another set of movable internal vanes called diffusers, which are distributed around a longitudinal axis. Each diffuser includes a panel designed to guide the exhaust gas flow within the nozzle, and each diffuser has a corresponding upstream end hinged to the downstream end of a convergent vane. Thus, such a nozzle is called a convergent-diffuse nozzle.

[0005] In this configuration, the control system is further configured to control the position of the diffuser based on the position of the convergent wing. Therefore, such a system allows for continuous variation in the tilt of the convergent wing relative to the longitudinal axis of the turbojet engine, and ensures that the tilt of the diffuser relative to this axis corresponds to the corresponding tilt of the convergent wing according to a defined law. Consequently, in particular, such a nozzle allows for variation in the position and shape of the nozzle neck.

[0006] It should be noted that the modifier "diffusion" does not preclude the airfoil under consideration from being oriented parallel to the longitudinal axis, or even convergent in some operational phases.

[0007] The important parameter used in operating such a nozzle is the ratio A9 / A8, where A8 refers to the cross-section of the nozzle at the neck where it is formed at the junction between the converging vane and the diffuser vane, and A9 refers to the cross-section of the nozzle at the downstream end of the diffuser vane.

[0008] For a given nozzle, the range of variation of this ratio A9 / A8 depends on the construction of the device that controls the position of the diffuser vanes based on the position of the converging vanes.

[0009] In this context, there is a need for a nozzle with variable geometry and an efficient and compact control system for its movable internal vanes, which can be integrated into a limited space while maximizing the range of possible values ​​for the ratio A9 / A8. Summary of the Invention

[0010] In particular, the purpose of this invention is to address this need in a simple, economical and efficient manner.

[0011] Therefore, the present invention provides a rear portion of a turbojet engine, the rear portion of which includes:

[0012] -Upstream stator structure;

[0013] - A convergent-diffracting nozzle with variable geometry, comprising a set of convergent vanes distributed around the longitudinal axis of the rear portion of a turbojet engine, each convergent vane comprising a panel designed to guide exhaust gas flow within the nozzle, and each convergent vane having an upstream end hinged to an upstream stator structure; the convergent-diffracting nozzle further comprises a set of diffuser vanes distributed around the longitudinal axis, each diffuser vane comprising a panel designed to guide exhaust gas flow within the nozzle, and each diffuser vane having an upstream end hinged to the downstream end of a corresponding convergent vane;

[0014] - Synchronization ring, which is arranged around a set of convergent fins or upstream stator structure;

[0015] - Drive unit, the drive unit is configured to cause the synchronization ring to translate relative to the upstream stator structure according to the longitudinal axis;

[0016] - An actuation device, through which a synchronizing ring acts on at least some convergent vanes to control the geometric shape change of the nozzle, wherein the at least some convergent vanes are referred to as controlled convergent vanes; and

[0017] - An apparatus for controlling the position of a diffuser based on the position of a converging vane, the apparatus comprising a first link, a second link, and a third link, the first link having a first end hinged to a corresponding converging vane and an opposite second end; the second link having a first end hinged to a corresponding diffuser vane and an opposite second end; and the third link having a first end hinged to a synchronizing ring and an opposite second end.

[0018] Wherein, the second end of the first link is hinged to at least one of the second ends of the second link and the second end of the third link, and

[0019] The second end of the third link is hinged to the second end of the second link.

[0020] Therefore, the device for controlling the winglets can have the most optimally limited size and mass, while allowing for a wider range of values ​​for the ratio A9 / A8.

[0021] In a preferred embodiment of the invention, the second end of the first link is hinged together with the second end of the second link and the second end of the third link according to a common hinge axis.

[0022] In other preferred embodiments of the invention, the second end of the third link is hinged to the second end of the second link according to the fifth hinge axis, and the second end of the first link is hinged to the second end of the second link according to the sixth hinge axis located between the fifth hinge axis and the first end of the second link.

[0023] Preferably, the ratio A9 / A8 is equal to the ratio of the section A9 of the nozzle at the downstream end of the diffuser vane to the section A8 of the nozzle at the neck of the joint formed between the converging vane and the diffuser vane, and varies in the range greater than 0.35.

[0024] The present invention also relates to a turbojet engine for an aircraft, which includes a rear portion of the type described above. Attached Figure Description

[0025] The invention will be better understood by reading the following description, which is given by way of non-limiting example and with reference to the accompanying drawings, and other details, advantages and features of the invention will become apparent:

[0026] - Figure 1 It is a schematic half-view of the axial section of a turbojet engine including nozzles with variable geometry;

[0027] Figure 2 is a schematic half-view of the axial section of the rear portion of a known type of turbojet engine;

[0028] - Figure 3 According to an embodiment of the present invention, in the low-convergence configuration of the convergent wing, Figure 1 A schematic half-view of the axial section of the rear portion of a turbojet engine;

[0029] - Figure 4 In the highly convergent structure of the convergent winglets Figure 1 A schematic half-view of the axial section of the rear portion of a turbojet engine;

[0030] - Figure 5 It constitutes Figure 1 A schematic perspective view of the rear section of a turbojet engine, including certain elements designed to control the convergent blades;

[0031] - Figure 6 and Figure 7 yes Figure 5 Enlarged schematic perspective view of some of the visible components;

[0032] - Figure 8 This is a partial schematic perspective view of the rear section of a variant of a turbojet engine, showing elements intended for use as diffuser blades for servo-controlled nozzles.

[0033] In all these diagrams, the same label can refer to the same or similar elements. Detailed Implementation

[0034] Figure 1 A turbojet engine 10 (e.g., a twin-rotor turbofan engine) is shown, which is designed for propulsion of an aircraft capable of supersonic flight, and is therefore specifically intended to be mounted in the fuselage of such an aircraft. Of course, the invention is applicable to other types of turbojet engines.

[0035] Throughout this description, the axial direction X is the direction of the longitudinal axis 11 of the turbojet engine. Unless otherwise stated, the radial direction R is the direction orthogonal to and passing through the longitudinal axis 11 at all points, and the circumferential direction C is the direction orthogonal to both the radial direction R and the longitudinal axis 11 at all points. Unless otherwise stated, the terms "inner" and "outer" refer to the relative proximity and relative distance of an element from the longitudinal axis 11, respectively. Finally, the modifiers "upstream" and "downstream" are defined relative to the overall direction D of gas flow within the turbojet engine 10.

[0036] For illustration, such a turbojet engine 10 includes, from upstream to downstream, an air inlet 12, a low-pressure compressor 14, a high-pressure compressor 16, a combustion chamber 18, a high-pressure turbine 20, a low-pressure turbine 22, a secondary combustion passage 26, and a nozzle 28 with variable geometry (e.g., a convergent-diffusion type nozzle). All these components of the turbojet engine are centered relative to the longitudinal axis 11 of the turbojet engine.

[0037] As is known, a high-pressure compressor 16, a combustion chamber 18, a high-pressure turbine 20, and a low-pressure turbine 22 define a main flow path PF. The main flow path PF is surrounded by a secondary flow path SF of the turbine engine, which extends downstream from the outlet of the low-pressure compressor. Therefore, in operation, air F1 entering through air inlet 12 and compressed by the low-pressure compressor 14 is then divided into a main flow F2 flowing in the main flow path and a secondary flow F3 flowing in the secondary flow path. The main flow F2 is then further compressed in the high-pressure compressor 16, mixed with fuel, and ignited in the combustion chamber 18, then undergoes expansion in the high-pressure turbine 20, and then further expansion in the low-pressure turbine 22.

[0038] Then, the exhaust gas F4, which is a mixture of the combustion gases generated from the main stream and the secondary stream F3, flows through the secondary combustion channel 26 and then escapes from the turbojet engine 10 through the nozzle 28.

[0039] When using an afterburner to control engine speed, such as when propelling an aircraft at supersonic speeds, fuel is mixed with exhaust gas F4 in afterburner passage 26, and the resulting mixture is ignited in the afterburner passage to generate additional thrust.

[0040] Figure 2 shows the rear portion of a turbojet engine in a configuration known from the prior art at a larger scale, with particular emphasis on the movable internal fins of the nozzle.

[0041] The movable internal wing consists of a set of converging winglets 30 located upstream and distributed around the longitudinal axis 11, and a set of diffusing winglets 32 located downstream and also distributed around the longitudinal axis 11.

[0042] Each of these movable internal flaps includes panels 31, 33, which help to define the exhaust gas recirculation passage 34 from the outside, which is defined in an extension of the secondary combustion passage 26. Therefore, the movable internal flaps 30, 32 enable the exhaust gas flow F4 to be directed at the exit of the turbojet engine 10 during operation.

[0043] The convergent flap 30 is hinged at its upstream end 36 to the stator structure 38 of the rear portion of the turbojet engine, in which case it is hinged to the internal U-shaped clip 40 of the beam 42 belonging to the stator structure, so that the convergent flap 30 can rotate about a first hinge axis A1 attached to the stator structure 38.

[0044] The diffuser 32 is hinged at its upstream end 44 to the downstream end 46 of the convergent 30, allowing the diffuser 32 to rotate about a second hinge axis A2, which is attached to the convergent 30 and is generally parallel to the first axis A1. The diffuser 32 is also hinged at its downstream end 48 to the first end 50A of the connecting rod 50, and the opposite second end 50B of the connecting rod 50 is hinged to the stator structure 38, in this case, to the outer U-shaped clamp 54 of the beam 42.

[0045] The control system for the movable internal vanes includes a drive mechanism configured to act on at least some of the convergent vanes, which will be referred to below as controlled convergent vanes. Where other convergent vanes are acted upon by the drive mechanism only through the controlled convergent vanes, these other convergent vanes are referred to in a known manner as follower convergent vanes.

[0046] The drive unit typically consists of cylinders 56, each cylinder 56 having a stator portion fastened to a stator structure 38 and a movable portion fixed to a corresponding roller bracket 58. The stator portion is, for example, the cylinder body 56A, and the movable portion is, for example, the cylinder rod 56B. Rollers 60 are mounted on the roller brackets 58 to support rolling on cams 62. The cams 62 are formed by a structure 64 fixed to a panel 31 of a corresponding convergent airfoil 30. The roller brackets 58 are further fixed to retaining fingers 66, which cooperate with the structure 64 to radially retain the convergent airfoil 30, particularly preventing the airfoil from lowering under gravity when the turbojet engine is stopped. Thus, a set of movable internal airfoils 30 and 32 together with the stator structure 38 form an isostatic system.

[0047] Therefore, the translational movement of the movable portion of each cylinder 56 enables the converging vane 30 to rotate about the first hinge axis A1, accompanied by the rotational movement of the diffuser vane 32 about the second hinge axis A2, during which the stroke of the diffuser vane is determined by the connecting rod 50. Such movement of the movable inner vanes 30, 32 results in modification of the nozzle profile, particularly modifying the cross-section of the nozzle neck at the junction between the converging and diffuser vanes.

[0048] The nozzle further includes a movable outer vane 70, the upstream end 72 of which is hinged to the stator structure 38, for example, to the outer U-shaped clamp 54 of the beam 42, and the downstream end 74 of which is attached to the downstream end 48 of the diffuser vane 32, for example, by means of a roller connection 76 and a sliding connection 78.

[0049] The variable geometry of nozzle 28 allows it to adapt to different flight phases. Thus, at subsonic speeds, the converging inner winglets remain in a low-convergence configuration, while at supersonic speeds, the converging inner winglets employ a more convergent configuration.

[0050] An important parameter for operating such a nozzle is the ratio A9 / A8, where A8 refers to the cross-section of the nozzle at the neck 79A formed at the junction between the converging vane 30 and the diffuser vane 32, and A9 refers to the cross-section of the nozzle at the downstream end 79B of the diffuser vane 32. This ratio A9 / A8 typically has a value between 1.1 and 1.8.

[0051] For a given nozzle, the range of values ​​for the ratio A9 / A8 variation depends on the construction of the device implemented as a position servo control of the diffuser 32 based on the position of the convergent fin 30. In the example described above, this range is therefore determined by the characteristics of the link 50 and by connecting the link 50 to the diffuser 32 and the stator structure 38, and the width of this range is typically 0.15.

[0052] Now, refer to Figures 3 to 7 A more detailed description of embodiments according to the present invention Figure 1 The rear section of the turbojet engine is designed to achieve a wider range of values ​​for the A9 / A8 ratio.

[0053] For the sake of illustration, the preferred mode for controlling the convergent vane will be described first below.

[0054] According to this preferred embodiment, each of the controlled convergent vanes 30 includes a lever 80 fixed to a panel 31 of the vane. Such a lever 80 extends from the panel 31 in a direction away from the longitudinal axis 11, or, in the example shown, from a reinforcing structure 81 disposed on and fixed to the outer surface of the panel 31 in a direction away from the longitudinal axis 11.

[0055] Similar to the description above, the rear portion of the turbojet engine includes a drive unit, which comprises a movable portion that can move axially relative to the upstream stator structure 38 upon command. For illustration, the drive unit here again consists of cylinders 56, and all the rods 56B of the cylinders form the movable portion.

[0056] In order for the movable part of the drive device to act on at least one controllable convergent blade 30 lever 80, the lever 80 is axially arranged between an upstream support wall 90 and a downstream support wall 92, which are rigidly fixed to the movable part of the drive device, such that the lever 80 is free to move relative to the upstream support wall 90 and the downstream support wall 92 at least in the radial direction R relative to the longitudinal axis 11.

[0057] In this manner, as the movable part of the drive unit (composed of rod 56B) moves downstream, the upstream support wall 90 pushes the lever 80 downstream, thus causing the converging vane 30 to pivot according to the corresponding first hinge axis A1, thereby bringing the downstream end 46 of the vane closer to the longitudinal axis 11.

[0058] Conversely, during the upstream movement of the movable portion of the drive unit, at least when the turbojet engine is not operating, the downstream support wall 92 pushes the lever 80 upstream, thus causing the convergent wing 30 to pivot according to the corresponding first hinge axis A1, resulting in the downstream end 46 of the wing moving away from the longitudinal axis 11. If the turbojet engine is operating, the gas thrust on the convergent wing 30 may be sufficient to pivot it, even before the downstream support wall 92 contacts the lever 80. Therefore, it is advantageous to provide that the upstream support wall 90 has increased stiffness compared to the downstream support wall 92. For this purpose, the upstream support wall 90 may be thicker than the downstream support wall 92, or the upstream support wall 90 may have stiffeners, while the downstream support wall 92 may not.

[0059] The lever 80 is provided with an axisymmetric cylindrical support roller 96, which is mounted on the lever 80 to rotate freely according to an axis 94 parallel to the corresponding first hinge axis A1. The support roller 96 is arranged between the upstream support wall 90 and the downstream support wall 92, such that the contact between either the upstream support wall 90 or the downstream support wall 92 on the lever 80 is a cylindrical / planar contact.

[0060] Therefore, during the pivoting maneuver of the converging vane under the thrust applied to the support roller 96 by one of the upstream support wall 90 and the downstream support wall 92, the radial movement of the support roller 96 relative to the support wall under consideration is performed by the rolling of the support roller 96 on the support wall.

[0061] The axial distance between the upstream support wall 90 and the downstream support wall 92 is greater than the diameter of the support roller 96, so that an axial clearance permanently exists between the support roller 96 and the support wall opposite to the support wall on which the thrust is applied.

[0062] Advantageously, the support roller 96 is arranged at the free end of the lever 80 to maximize the lever arm applied to the converging vane 30 by the upstream support wall 90 and the downstream support wall 92.

[0063] Therefore, for example, the support roller 96 is mounted on a shaft supported by two transverse arms 80A, 80B, which form the end fork of the lever 80. Figure 5 ).

[0064] Furthermore, one of the support walls (in this case, the downstream support wall 92) is connected to the movable part of the drive unit via another support wall (in this case, the upstream support wall 90).

[0065] For this purpose, the outer connecting wall 98 connects the corresponding radial outer ends of the upstream support wall 90 and the downstream support wall 92 together.

[0066] The foregoing description relating to the manipulation of the lever of a controlled convergent vane is preferably equally effective for other controlled convergent vanes.

[0067] Therefore, in the illustrated embodiment, the upstream support wall 90 arranged opposite to each lever 80 and the upstream support wall 90 arranged opposite to the two levers 80 closest to the lever under consideration are circumferentially spaced apart, and the downstream support wall 92 arranged opposite to each lever 80 and the downstream support wall 92 arranged opposite to the two levers 80 closest to the lever under consideration are circumferentially spaced apart. Figure 5 Therefore, the upstream support wall 90 and the downstream support wall 92 form a ring-shaped support device 100 spaced apart from each other, and each support device 100 includes a corresponding pair of support walls, which includes an upstream support wall 90 and a downstream support wall 92.

[0068] The rear portion of the turbojet engine further includes a synchro ring 82 arranged around a set of convergent blades 30, or alternatively, arranged slightly upstream around the upstream stator structure 38. Each of the support devices 100 is connected via the synchro ring 82 to a movable part of the drive unit, i.e., to all the rods 56B connected to the cylinder 56.

[0069] Specifically, the movable portion of the drive unit is connected to the synchronizing ring 82 to enable the synchronizing ring 82 to translate along the longitudinal axis 11. For this purpose, the rod 56B of the cylinder 56 is hinged to a first U-shaped clamp 84 of the synchronizing ring 82. Such a first U-shaped clamp 84 is formed to protrude from the body 86 of the synchronizing ring 82, for example, which has an annular shape. For example, the first U-shaped clamp 84 extends upstream from the body 86.

[0070] It should be noted that the body 86 of the synchronizing ring may have a more complex shape, including, for example, alternations of radially inward and radially outward projecting portions, and / or alternations of upstream and downstream projecting portions. In all cases, the body 86 of the synchronizing ring extends around the longitudinal axis 11 of the turbojet engine, and thus has a generally annular shape.

[0071] Each of the support devices 100 is connected to the synchronization ring 82 via, for example, three arms 102, which are circumferentially spaced apart from each other, and each arm 102 connects the synchronization ring 82 to the downstream support wall 92.

[0072] An example of a support device forming an actuation device, wherein a synchronization ring 82 acts on a controlled convergent vane 30 via the actuation device to control the geometric changes of the nozzle.

[0073] In the example shown, in each support device 100 (one support device in...) Figure 6 As can be seen in the image, a first connecting sidewall 104 connects the corresponding first circumferential ends of the upstream support wall 90 and the downstream support wall 92 together, and a second connecting sidewall 106 connects the corresponding second circumferential ends of the upstream support wall 90 and the downstream support wall 92 together, with the second circumferential ends opposite to the first circumferential ends. Therefore, the first connecting sidewall 104 and the second connecting sidewall 106 enable the upstream support wall 90 to connect to the downstream support wall 92 and thus to the synchronization ring 82, and to the movable part of the drive unit via the synchronization ring 82.

[0074] In operation, the lever 56B of each actuator 56 unfolds, or more generally, the movable portion of the drive mechanism unfolds downstream, causing the synchronizing ring 82 to move downstream, which in turn drives each upstream support wall 90 and downstream support wall 92 downstream. Thus, each upstream support wall 90 contacts the support roller 96 of the corresponding lever 80. Each upstream support wall 90 then pushes the support roller 96 downstream, and thus the lever 80 downstream, causing the corresponding vane to pivot along the longitudinal axis 11, which increases the convergence of the converging vane 30. During the vane pivoting, the support roller 96 rolls on the upstream support wall 90, allowing this rolling due to the gap between the roller 96 and the other support walls (in this case, the downstream support wall 92).

[0075] Conversely, the rod 56B of each actuator 56 retracts, or more generally, the movable portion of the drive mechanism retracts upstream, causing the synchronizing ring 82 to move upstream, which in turn drives each upstream support wall 90 and downstream support wall 92 upstream. Thus, if the turbojet engine stops operating, each downstream support wall 92 contacts the support roller 96 of the corresponding lever 80. Each downstream support wall 92 then pushes the support roller 96 upstream, and thus the lever 80 upstream, causing the corresponding blade to pivot in the opposite direction to the longitudinal axis 11, which reduces the convergence of the convergent blade 30. During blade pivoting, the support roller 96 rolls on the downstream support wall 92, which, in this context, is again allowed by the gap between the roller 96 and the other support walls (in this case, the upstream support wall 90). On the other hand, if the turbojet engine is in operation, the gas thrust on the convergent wing 30 may be sufficient to pivot the convergent wing 30, even before the downstream support wall 92 contacts the lever 80.

[0076] Therefore, during the manipulation to increase the convergence of the retractable vanes, cylinder 56 operates along the deployment direction of its rod 56B, which is mechanically advantageous. In fact, at least in the preferred case where cylinder 56 is a hydraulic cylinder, the deployment of the rod is caused by hydraulic pressure applied to the entire surface of the piston, while the retraction of the rod is caused by hydraulic pressure applied to the piston surface minus the rod's cross-section. At least for this reason, the deployment of the rod generally provides increased power compared to its retraction.

[0077] Furthermore, all the components involved in controlling the internal flaps, including lever 80, upstream support wall 90 and downstream support wall 92, and the means for connecting these upstream and downstream support walls to the movable part of the drive unit, can therefore have limited size and mass.

[0078] Furthermore, the levers 80 of each controlled convergent vane 30 are advantageously arranged at the upstream end of the vane to again, as optimally as possible, limit the size and mass of the control system for the vane.

[0079] In this case, it is advantageous to arrange the synchronization ring 82 downstream of the lever 80 of each controlled convergent vane 30.

[0080] It should be noted that the body 56A of the cylinder 56 can be rigidly fastened to the stator structure 38 in the same manner as the known example shown in Figure 2 and described above.

[0081] In an alternative embodiment, the upstream support walls 90 may be connected to each other to form a single upstream support structure extending within 360 degrees. Similarly, the downstream support walls 92 may be connected to each other to form a single downstream support structure extending within 360 degrees.

[0082] Such a support structure can be directly integrated into the main body 86 of the synchronization ring 82.

[0083] The following describes an apparatus for controlling the position of the diffuser 32 based on the position of the convergent 30.

[0084] For each diffuser 32, these devices include a first link 120, a second link 122, and a third link 124. The first link 120 has a first end 120A hinged to the corresponding convergent wing 30 and an opposite second end 120B. The second link 122 has a first end 122A hinged to the corresponding diffuser 32 and an opposite second end 122B. The third link 124 has a first end 124A hinged to the synchronizing ring 82 and an opposite second end 124B. Furthermore, the second end 120B of the first link 120 is hinged to at least one of the second end 122B of the second link 122 and the second end 124B of the third link 124. Finally, the second end 124B of the third link 124 is hinged to the second end 122B of the second link 122.

[0085] exist Figures 3 to 7 In the example shown, the second end 120B of the first link 120 is hinged to the second end 122B of the second link 122 and the second end 124B of the third link 124 according to the common hinge axis AC.

[0086] Furthermore, according to the third hinge axis A3, the first end 120A of the first link 120 is hinged to the reinforcing structure 81 fixed to the panel 31 of the convergent wing, and the third hinge axis A3 is advantageously closer to the first hinge axis A1 than the second hinge axis A2. In the example shown, the third hinge axis A3 is also closer to the centerline LM than the first axis A1, which is located halfway between the first axis A1 and the second axis A2. According to the fourth hinge axis A4, the first end 122A of the second link 122 is hinged to the U-shaped clip 126 fixed to the panel 33 of the diffuser wing, and the fourth hinge axis A4 is advantageously located approximately halfway between the second hinge axis A2 and the downstream end 79B of the diffuser wing. Finally, the first end 124A of the third link 124 is hinged to the second U-shaped clip 128 of the synchronizing ring 82. This second U-shaped clip 128 is formed to protrude from the body 86 of the synchronizing ring 82, for example, extending downstream and radially outward from the body 86.

[0087] exist Figure 8 In the variant shown, the second end 124B of the third link 124 is hinged to the second end 122B of the second link 122 according to the fifth hinge axis A5, and the second end 120B of the first link 120 is hinged to the second end 122B of the second link 122 according to the sixth hinge axis A6 located between the fifth hinge axis A5 and the first end 122A of the second link 122.

[0088] In operation, the first link 120, the second link 122, and the third link 124 determine the orientation of the diffuser 32 according to an explicit law that is a function of the orientation of the convergent wing 30.

[0089] This mode of servo-controlled diffuser position allows for a wider range of values ​​for the ratio A9 / A8. Therefore, this range can typically be increased to 0.35 or even larger.

Claims

1. A rear portion of a turbojet engine, the rear portion of the turbojet engine comprising: -Upstream stator structure (38); - A convergent-diffracting nozzle (28) having a variable geometry, the convergent-diffracting nozzle comprising a set of convergent vanes (30) distributed around the longitudinal axis (11) of the rear portion of the turbojet engine, each convergent vane comprising a panel (31) intended to guide exhaust gas flow (F4) within the nozzle, and each convergent vane having an upstream end (36) hinged to the upstream stator structure, the convergent-diffracting nozzle further comprising a set of diffuser vanes (32) distributed around the longitudinal axis (11), each diffuser vane comprising a panel (33) intended to guide the exhaust gas flow within the nozzle, and each diffuser vane having an upstream end (44) hinged to the downstream end (46) of the corresponding convergent vane; - Synchronization ring (82), the synchronization ring being arranged around the set of convergent winglets (30) or the upstream stator structure (38); - A drive device configured to cause the synchronization ring (82) to translate relative to the upstream stator structure (38) along the longitudinal axis (11); - An actuation device, wherein the synchronization ring (82) acts on at least some convergent vanes (30) via the actuation device to control the geometric change of the nozzle, wherein the at least some convergent vanes are referred to as controlled convergent vanes; and - An apparatus for controlling the position of the diffuser based on the position of the convergent wing, the apparatus comprising a first link (120), a second link (122), and a third link (124), the first link having a first end (120A) hinged to the corresponding convergent wing (30) and an opposite second end (120B), the second link having a first end (122A) hinged to the corresponding diffuser (32) and an opposite second end (122B), and the third link having a first end (124A) hinged to the synchronization ring (82) and an opposite second end (124B). Wherein, the second end (120B) of the first link (120) is hinged to at least one of the second end (122B) of the second link (122) and the second end (124B) of the third link (124), and The second end (124B) of the third link (124) is hinged to the second end (122B) of the second link (122).

2. The rear portion of the turbojet engine according to claim 1, wherein, According to the common hinge axis (AC), the second end (120B) of the first link (120) is hinged to the second end (122B) of the second link (122) and the second end (124B) of the third link (124).

3. The rear portion of the turbojet engine according to claim 1, wherein, According to the fifth hinge axis (A5), the second end (124B) of the third link (124) is hinged to the second end (122B) of the second link (122), and according to the sixth hinge axis (A6) located between the fifth hinge axis (A5) and the first end (122A) of the second link (122), the second end (120B) of the first link (120) is hinged to the second end (122B) of the second link (122).

4. The rear portion of the turbojet engine according to any one of claims 1 to 3, wherein, The ratio A9 / A8 is equal to the ratio of the section A9 of the nozzle at the downstream end of the diffuser vane to the section A8 of the nozzle at the neck of the junction formed between the converging vane and the diffuser vane, and varies in the range greater than 0.

35.

5. A turbojet engine for an aircraft, the turbojet engine comprising a rear portion according to any one of claims 1 to 4.

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