Assembly for a turbomachine

Abstract

The invention relates to an assembly for a turbomachine, the assembly comprising: - a nacelle comprising an air intake lip defining an air intake, and - means for varying the geometry of the air intake, the means comprising: - a first nozzle, and - a second nozzle, the first nozzle and the second nozzle translating with respect to the nacelle between a first configuration and a second configuration: - in the first configuration, the first nozzle forms the air intake lip and the second nozzle extends inside the nacelle, and - in the second configuration, the first nozzle extends away from the air intake lip and the second nozzle forms the air intake lip so as to define a flow passage between a downstream surface of the first nozzle and an upstream surface of the second nozzle.

Description

TECHNICAL FIELD

[0001] The present invention relates to a nacelle of a gas turbine engine.

[0002] More specifically, the object of the present invention is a nacelle of a gas turbine engine comprising means for varying the geometry of the intake. BACKGROUND

[0003] Known gas turbine engines have a high bypass ratio in order to improve the propulsive efficiency of the gas turbine engine. However, these gas turbine engines have a large volume and mass which penalizes the efficiency of the gas turbine engine in propelling the aircraft and leads to an increase in the geometric installation constraints.

[0004] In order to eliminate this drawback, it is known to shorten the nacelle of the gas turbine engine. Nonetheless, the intake is deprived of the ability to generate and / or transmit a distortion of the intake which limits the passage of the airflow into the gas turbine engine. This distortion is detrimental to the stability and performance of the gas turbine engine, in particular of the fan of the gas turbine engine.

[0005] The known method to solve this problem consists in installing on the nacelle means for varying the geometry of the intake. Such means generally comprise a mobile part which moves between a retracted position, in which the intake has a nominal geometry, and a deployed position, in which the geometry of the intake is varied. In this way, during low speed advancement of the aircraft, the run of the stall point in the airflow can be reduced thanks to the opening of an additional supply section of the incoming airflow. Furthermore, the usual phenomenon of bypass of the flow on the intake lip is minimized. Indeed, this phenomenon usually occurs at low speed advancement, at which time the radially outer part of the flow is deflected towards the outside of the nacelle due to the convex shape of the intake lip. This leads to a loss of incoming airflow which can be detrimental to the performance of the gas turbine engine.

[0006] Nonetheless, the means for varying the geometry of the intake known from the prior art have many drawbacks. For example, the deployment of the means leads to unnecessary aerodynamic phenomena, i.e. perturbations of the flow (such as the delamination of the boundary layer in the flow), leading to a major charge loss in the incoming airflow. Furthermore, in the retracted position, the mobile part has many discontinuities which increase the drag in the flow. In addition, the mechanism required to pass from the retracted position to the deployed position is usually complex and bulky, which limits the integration of the mechanism within the nacelle of the gas turbine engine. Finally, such means do not have a redundancy system in the event of failure of the deployment mechanism.

[0007] There is therefore a need to eliminate at least one of the drawbacks of the prior art described above. SUMMARY

[0008] One of the objectives of the application is to reduce the creation and / or the transmission of distortions in the intake of a gas turbine engine, in particular a gas turbine engine comprising a nacelle with reduced axial dimensions and thickness.

[0009] Another objective of the application is to improve the aerodynamic properties of the nacelle of a gas turbine engine whatever the conditions of the airflow.

[0010] Another objective of the application is to propose a nacelle whose intake has a variable geometry.

[0011] Another objective of the application is to strengthen the robustness of the aerodynamic behavior of the nacelle over its entire field of operation.

[0012] Another objective of the application is to limit unnecessary aerodynamic phenomena inside and outside the nacelle during deployment of the device for varying the geometry of the intake.

[0013] To this end, an objective of the application is an assembly for a gas turbine engine, comprising:

[0014] - a nacelle of a gas turbine engine, the nacelle of the gas turbine engine having a longitudinal axis and comprising an intake lip defining an intake, and

[0015] - a device for varying the geometry of the intake, said device comprising:

[0016] o a first slat, and

[0017] o a second slat,

[0018] the first slat and the second slat being translationally mobile between a first configuration and a second configuration relative to the nacelle according to the longitudinal axis:

[0019] ■ in the first configuration, the first slat forms the intake lip and the second slat extends inside the nacelle without interacting with the outside airflow, and

[0020] ■ in the second configuration, the first slat extends from the intake lip by a distance and the second slat forms the intake lip so as to define an airflow passage between a downstream surface of the first slat and an upstream surface of the second slat.

[0021] In this assembly, thanks to the airflow channel in the second configuration, the air capture surface can be increased at low speed advancement of the aircraft on which the nacelle is mounted, i.e. the absolute speed of the aircraft relative to the ground. Moreover, still thanks to the second configuration, at low speed advancement, the length of the intake can be lengthened, which limits the generation and / or the transmission of distortions within the intake. Furthermore, thanks to the first configuration, at high speed advancement, the length of the intake remains short, which limits the aerodynamic losses of the gas turbine engine. Moreover, the transition from the first configuration to the second configuration ensures a stable aerodynamic behavior of the flow. Finally, whatever the operating conditions of the gas turbine engine, the geometry of the airflow channel can be designed by locally modifying, i.e. according to the circumferential direction around the longitudinal axis, the downstream surface of the first slat and / or the upstream surface of the second slat, to optimize the performance of the inlet flow.

[0022] Advantageously, and optionally, the assembly according to the application can also comprise at least one of the following features, which can be taken individually or in combination:

[0023] - the second slat has, in the radial plane, an upstream profile of the second slat with tangential continuity, in the second configuration, with the profile of the nacelle defined in the radial plane,

[0024] - the first slat has, in the radial plane relative to the longitudinal axis, an upstream profile of the first slat, the second slat having, in the radial plane, an upstream profile of the second slat identical to the upstream profile of the first slat,

[0025] - the second slat:

[0026] o has, in a first radial plane relative to the longitudinal axis, a first upstream profile of the second slat, and

[0027] o has, in a second radial plane relative to the longitudinal axis, a second upstream profile of the second slat.

[0028] the second upstream profile of the second slat being different from the first upstream profile of the second slat, so that the airflow channel has, in the first radial plane, a first longitudinal section and, in the second radial plane, a second longitudinal section different from the first longitudinal section,

[0029] - the first slat:

[0030] o has, in a first radial plane relative to the longitudinal axis, a first downstream profile of the first slat, and

[0031] o has, in a second radial plane relative to the longitudinal axis, a second downstream profile of the first slat.

[0032] the second downstream profile of the first slat being different from the first downstream profile of the first slat, so that the airflow passage has a first longitudinal section in a first radial plane and a second longitudinal section in a second radial plane, the second longitudinal section being different from the first longitudinal section,

[0033] - the first slat and the second slat are connected, on at least one corner portion of the device for varying the geometry of the intake, to form a single slat which also translates with respect to the nacelle according to the longitudinal axis,

[0034] - the first slat and the second slat extend on a first corner portion of the nacelle, a second corner portion of the nacelle being delimited by a fixed portion of the intake lip,

[0035] - the device for varying the geometry of the intake comprises a guide rod which extends parallel to the longitudinal axis and which connects the first slat to the second slat, so that the first slat is fixedly mounted with respect to the second slat on at least one corner portion of the device for varying the geometry of the intake,

[0036] - the device for varying the geometry of the intake comprises a guide rod which extends parallel to the longitudinal axis, the guide rod being telescopic so that the movement of the first slat is independent of the movement of the second slat on at least one corner portion of the device for varying the geometry of the intake,

[0037] - the device for varying the geometry of the intake comprises an anti-icing system of the first slat and / or of the second slat, the anti-icing system comprising an anti-icing fluid flow passage which extends at least partially inside the guide rod,

[0038] - the device for varying the geometry of the intake further comprises an actuator which is housed in the nacelle and which is configured to move the first slat and the second slat between the first configuration and the second configuration,

[0039] - the nacelle has a non-axially symmetrical form.

[0040] Another object of the application is a gas turbine engine comprising a nacelle as described previously. BRIEF DESCRIPTION OF DRAWINGS

[0041] Other characteristics, objects and advantages of the application will emerge from the following description, which is purely illustrative and non-limiting, and must be considered with reference to the attached drawings, in which:

[0042] Figure 1 is a partial view of a nacelle of a gas turbine engine,

[0043] Figure 2 is a perspective view of a first embodiment of an assembly for a gas turbine engine according to the application in a first configuration,

[0044] Figure 3 yes Figure 1 The first embodiment of the components for a gas turbine engine shown is in perspective view of a second configuration.

[0045] Figure 4 This is a cross-sectional view of a portion of a second embodiment of a component for a gas turbine engine according to the present invention, within the first configuration.

[0046] Figure 5 This is a cross-sectional view of a third embodiment of a component for a gas turbine engine according to the present invention, located in the first configuration.

[0047] Figure 6 yes Figure 4 The second embodiment of the component for a gas turbine engine shown is a cross-sectional view of a portion of the second configuration.

[0048] Figure 7 This is a cross-sectional view of a portion of a fourth embodiment of a component for a gas turbine engine according to the present invention, in the first configuration.

[0049] Figure 8 yes Figure 7 The diagram shows a cross-sectional view of the fourth embodiment of the components for a gas turbine engine, in the second configuration.

[0050] Figure 9 This is a front view of a fifth embodiment of a component for a gas turbine engine according to the present invention.

[0051] Figure 10 This is a cross-sectional view of the portion of the sixth embodiment of a component for a gas turbine engine according to the present invention, in the second configuration.

[0052] Figure 11 This is a cross-sectional view of the seventh embodiment of a component for a gas turbine engine according to the present invention, in the first configuration.

[0053] Figure 12 yes Figure 11 The seventh embodiment of the components for a gas turbine engine shown is a cross-sectional view of the portion in the second configuration.

[0054] Figure 13 This is a perspective side view of an eighth embodiment of a component for a gas turbine engine according to the present invention.

[0055] Figure 14 This is a cross-sectional view of a ninth embodiment of a component for a gas turbine engine according to the present invention.

[0056] Figure 15 yesFigure 14 the front view of a tenth embodiment of an assembly for a gas turbine engine according to the present application, as shown in the accompanying drawings, in which:

[0057] Figure 16 is a cross-sectional view of a tenth embodiment of an assembly for a gas turbine engine according to the present application, and

[0058] Figure 17 is Figure 14 the front view of a tenth embodiment of an assembly for a gas turbine engine according to the present application, as shown in the accompanying drawings, in which: DETAILED DESCRIPTION

[0059] Nacelle of a gas turbine engine

[0060] With reference to Figures 1 to 3 , the nacelle 1 of the gas turbine engine has a longitudinal axis X-X around which it extends, for example in a rotational manner, generally in a substantially cylindrical manner when rotating. However, this is not limiting as the nacelle 1 can also extend around the longitudinal axis X-X in a non-axisymmetric manner, i.e. in a manner in which the longitudinal axis X-X is not orthogonal to a plane passing through all points forming an upstream end of the nacelle 1.

[0061] The nacelle 1 further comprises an air intake lip 10 defining an air intake 100 through which air can enter into the gas turbine engine. At an upstream end of the nacelle 1, the air intake lip 10 joins an inner wall 12 of the nacelle 1 with an outer wall 14 of the nacelle 1, the inner wall of the nacelle being in contact with the airflow entering into the gas turbine engine, the outer wall of the nacelle being in contact with the air circulating around the gas turbine engine.

[0062] Throughout the following, upstream and downstream are defined with respect to the direction of the airflow within the gas turbine engine. Furthermore, “axial” defines a direction substantially parallel to the longitudinal axis X-X, while “radial” defines a direction substantially orthogonal to said longitudinal axis X-X.

[0063] The nacelle 1 generally encloses an engine body 2 of the gas turbine engine, a fan 20 extending at an upstream of the engine body. When in operation, the fan 20 sucks the incoming airflow through the air intake 100 and pushes it into the interior of the engine body 2. As Figure 1 shown, the fan 20 has a diameter D and the air intake 100 has a length L defined as the distance separating the upstream end of the air intake lip 10 from the upstream end of the fan 20. More precisely, as Figure 1As shown, the length L corresponds to the distance between the leading edge of the blade root of the fan 20 and the intersection of the longitudinal axis X-X and the air capture plane (i.e. the plane passing through the end of the intake lip 10), this distance being taken along the longitudinal axis X-X. In other words, the length L corresponds to the distance between the leading edge of the blade root of the fan 20 and the plane orthogonal to the leading edge of the intake lip 10, this distance being taken in the region of the median plane of the nacelle 1, the median plane containing the longitudinal axis X-X and being orthogonal to the plane of symmetry of the nacelle 1, i.e. the plane passing through the 90° and 270° azimuthal angles of the intake 100. In the preferred embodiment, the dimensions of the nacelle 1 are such that the ratio of the length of the intake L to the diameter of the fan D is between 0.15 and 0.25, and preferably 0.19. In such an embodiment, the aerodynamic properties of the gas turbine engine are optimized, in particular with respect to the main frame of the engine body 2, i.e. with respect to the dimensions of the engine body 2, the resistance being minimal.

[0064] Device for changing the geometry of an air intake

[0065] The device 3 for varying the geometry of the intake is in particular shown in Figures 2 to 8 , Figures 10 to 14 and Figure 16 , thanks to which the nacelle 1 can have at least two configurations.

[0066] In the first configuration shown in Figure 2 , Figure 4 , Figure 5 , Figure 7 and Figure 11 , the geometry of the intake 100 is particularly adapted to the conditions of cruise flight, resulting in an infinitely high airflow upstream of the nacelle 1 at the high speed of advance of the aircraft to which the nacelle 1 is fixed. This geometry of the first configuration in fact has a very uniform flow in the region of the intake 100. In this first configuration, the device 3 for varying the geometry of the intake is in a retracted or retracted position.

[0067] In the second configuration shown in Figure 3 , Figure 8 , Figure 10 , Figures 12 to 14 and Figure 16 , the geometry of the intake 100 is particularly adapted to the conditions of takeoff flight at lower speed airflows, landing and / or on the ground. This geometry of the second configuration provides a wider air suction surface and shifts the stop point of the flow entering downstream of the intake 100 to reduce the bypass of the flow on the intake lip 10. In addition, this geometry of the second configuration minimizes the distortion of the flow of the airflow in the region of the intake, especially in conditions of crosswind or high impact wind. In this second configuration, the device 3 for varying the geometry of the intake is in a deployed position.

[0068] In any case, the device 3 for varying the geometry of the intake improves the ability of the nacelle 1 to distort the airflow entering in the area of the intake 100, which in particular improves the aerodynamic performance of the fan 20.

[0069] As seen more precisely in Figures 4 to 8 , Figures 10 to 14 and Figure 16 , the device 3 for varying the geometry of the intake comprises a first slat 30 and a second slat 32. As shown on these figures, the first slat 30 and the second slat 32 translate according to a longitudinal axis X-X relative to the nacelle 1 between a first configuration and a second configuration:

[0070] - the first configuration, illustrated in Figure 2 , Figure 4 , Figure 5 , Figure 7 and Figure 11 , wherein:

[0071] o the first slat 30 forms the intake lip 10, and

[0072] o the second slat 32 extends inside the nacelle 1, and

[0073] - the second configuration, illustrated in Figure 3 , Figure 8 , Figure 10 , Figures 12 to 14 and Figure 16 , wherein:

[0074] o the first slat 30 extends from the intake lip 10 by a distance, and

[0075] o the second slat 32 forms the intake lip 10,

[0076] In the first configuration, the second slat 32 remains inside the nacelle 1 without interacting, i.e. interacting in a limited manner, with the airflow circulating around the intake lip 10.

[0077] In the second configuration, the first slat 30 and the second slat 32 define, between a downstream surface of the first slat 30 and an upstream surface of the second slat 32, an airflow passage 34.

[0078] In this way, in the first configuration, the intake 100 has a nominal geometry, i.e. adapted in particular to the conditions of cruise flight at high speed airflow, in the second configuration, the geometry of the intake 100 is varied to be adapted in particular to the conditions of take-off flight at lower speed airflow and / or on the ground. The first configuration and the second configuration of the first slat 30 and of the second slat 32 respectively correspond to a retracted position and to a deployed position of the device 3 for varying the geometry of the intake.

[0079] With reference to Figure 3 , Figures 7 to 9 , Figure 13 , Figure 15 and Figure 17 , the device 3 for varying the geometry of the intake can also comprise a guide rod 36 connecting the first slat 30 to the second slat 32, so that on at least one corner portion of the device 3 for varying the geometry of the intake, the first slat 30 is fixedly mounted with respect to the second slat 32. However, this is not limiting, as the first slat 30 and the second slat 32 can not be fixedly mounted with respect to each other on another corner portion of the device 3 for varying the geometry of the intake, wherein, for example, the first and second slats are connected by a telescopic guide rod 36, so that the movement of the first slat 30 is independent of the movement of the second slat 32. In this way, it is possible to adjust the width of the airflow channel 34 according to the flight conditions, and to maintain the continuity of the joint between the second slat 32 and the inner wall 12

[0080] Advantageously, as shown in Figure 13 , the guide rod 36 extends parallel to the longitudinal axis X-X, so as to facilitate the deployment and stowage movement of the slats 30, 32 within the nacelle 1.

[0081] Also advantageously, as shown in Figure 9 , Figure 13 , Figure 15 and Figure 17 , the device 3 for varying the geometry of the intake comprises a plurality of guide rods 36 distributed around the longitudinal axis X-X, for example uniformly spaced in an azimuthal manner. In this way, the device 3 for varying the geometry of the intake has enhanced robustness and can perform the deployment and stowage movement of the slats 30, 32 within the nacelle 1 quickly and precisely.

[0082] In an embodiment, an anti-icing system of the first slat 30 and / or of the second slat 32 is provided. Indeed, in high-altitude areas, the risk of icing in the area of the intake lip 10 is particularly high. This system therefore ensures the normal operation of the device 3 for varying the geometry of the intake. Advantageously, the anti-icing system comprises an anti-icing fluid flow passage extending at least partially inside the guide rod 36. Thus, the guide rod 36 can be at least partially hollow in order to convey an anti-icing fluid, for example hot air. This ensures anti-icing of the first slat 30 and / or of the second slat 32 regardless of the configuration of the device 3 for varying the geometry of the intake. In this regard, the embodiment in which the first slat 30 is fixedly mounted with respect to the second slat 32, as shown in Figure 4 , is particularly advantageous. Indeed, the presence of ice on the upstream surface of the second slat 32 does not prevent the retraction of the device 3 for varying the geometry of the intake.

[0083] Reference Figure 7 and Figure 8 The device 3 for changing the geometry of the air intake may also include an actuator 38 configured to be housed in the nacelle 1. The actuator 38 is also configured to move the first slat 30 and / or the second slat 32 between the first configuration and the second configuration.

[0084] Advantageously, the actuator 38 is configured to move the guide rod 36 to switch from the first configuration to the second configuration and vice versa.

[0085] In any case, refer to Figures 4 to 8 , Figures 10 to 12 , Figure 14 and Figure 16 The guide rail 380 extends parallel to the longitudinal axis XX and is configured to translatably guide the first slat 30 and / or the second slat 32. In this configuration, the first slat 30 and / or the second slat 32 can be advantageously translatably mounted on one or more guide rails 380. Figure 5 As shown, guide rail 380 has an axial length δ i Also like Figure 10 As shown, due to volume considerations, the axial length δ of guide rail 380 is... i Preferably, it is less than or equal to the distance between the inlet neck and the air inlet end of the divergence section 100.

[0086] Actuator 38 can be pneumatic. In this case, the anti-icing system (e.g., hot air system) of cabin 1 can be connected to actuator 38.

[0087] However, as Figures 4 to 6 As shown, the presence of actuator 38 ensures that the deployment and / or retraction of the device 3 for changing the geometry of the air inlet is not mandatory. In fact, given the structure of the device 3 for changing the geometry of the air inlet, the deployment and / or retraction of this structure can be performed automatically as follows:

[0088] - Under high-speed airflow, the dynamic pressure upstream of the gas turbine engine generates an infinitely large axial force towards the downstream. This axial force naturally pushes the device 3 used to change the geometry of the intake back to its retracted position, which corresponds to... Figure 4 , Figure 7 and Figure 11 The first configuration shown in the figure, and

[0089] - during the low-speed advancement of the aircraft fitted with the gas turbine engine, and under the high air flow entered by the gas turbine engine (for example, at take-off), during the bypassing of the intake lip 10, the acceleration of the flow around said intake lip 10 generates a pressure field that generates an axial force directed upstream, which naturally unfolds the device 3 for varying the geometry of the intake, as shown in Figure 6 , and keeps it in the unfolded position, which corresponds to the second configuration shown in Figure 8 , Figure 10 , Figures 12 to 14 and Figure 16 .

[0090] Therefore, alternatively or in addition to the actuator 38, a locking device can be provided, configured to fix the device 3 for varying the geometry of the intake in the first configuration and / or in the second configuration, the unfolding and stowing of the device also being able to be ensured naturally by the high-speed and / or low-speed air flow. In this way, the device 3 for varying the geometry of the intake has considerable robustness and it is no longer necessary to provide safety devices to correct any malfunction of the actuator 38.

[0091] Geometry of the first and second slat and of the airflow passage

[0092] With reference to Figure 10 , the geometry of the first slat 30, of the second slat 32 and of the air flow channel 34 will now be described in greater detail. Figure 10 is a sectional view of the device 3 for varying the geometry of the intake according to a radial plane P with respect to the longitudinal axis X-X, i.e. in a plane P comprising the longitudinal axis X-X.

[0093] As shown in Figure 10 , the nacelle 1 defines:

[0094] - a first radius R1, which corresponds to the radius of the inlet neck, i.e. to the distance with which the surface of the inner wall 12 of the nacelle 1 closest to the longitudinal axis X-X is spaced from said longitudinal axis X-X,

[0095] - a second radius R2, which corresponds to the radius of the end of the divergence of the intake 100, i.e. to the distance with which the surface of the inner wall 12 of the nacelle 1 farthest from the longitudinal axis X-X is spaced from said longitudinal axis X-X, and

[0096] - a thickness e of the intake lip 10, which corresponds to the distance with which the surface of the inner wall 12 of the nacelle 1 closest to the longitudinal axis X-X is spaced from the surface of the outer wall 14 of the nacelle 1 farthest from said longitudinal axis X-X.

[0097] As shown in Figure 10As illustrated, the first slat 30 defines a third radius R3 corresponding to the distance separating the longitudinal axis X-X from the outer radial face 301 of the first slat 30 and a fourth radius R4 corresponding to the distance separating the longitudinal axis X-X from the inner radial face 300 of the first slat 30.

[0098] Moreover, the first slat 30 defines an upstream profile Cl of the first slat. The upstream profile Cl of the first slat forms an upstream surface 302 of the first slat, which is three-dimensionally rotated around the longitudinal axis X-X. A section of the upstream profile Cl of the first slat in a radial plane P forms a geometric line connecting, upstream of the first slat 30, the geometric line formed by the section in the radial plane P of the inner radial face 300 of the first slat 30 to the geometric line formed by the section in the radial plane P of the outer radial face 301 of the first slat 30. This geometric line can be defined by any curve passing, upstream of the first slat 30, through the geometric line formed by the section in the radial plane P of the inner radial face 300 of the first slat 30 and the geometric line formed by the section in the radial plane P of the outer radial face 301 of the first slat 30. Depending on the form of the upstream surface 302 of the first slat, the upstream profile Cl of the first slat can have a different form of geometric line in another radial plane P’, which also comprises the longitudinal axis X-X and is different from the radial plane P.

[0099] Moreover, the first slat 30 defines a downstream profile C2 of the first slat. The downstream profile C2 of the first slat forms a downstream surface 303 of the first slat, which is three-dimensionally rotated around the longitudinal axis X-X. A section of the downstream profile C2 of the first slat in a radial plane P forms a geometric line connecting, downstream of the first slat 30, the geometric line formed by the section in the radial plane P of the inner radial face 300 of the first slat 30 to the geometric line formed by the section in the radial plane P of the outer radial face 301 of the first slat 30. This geometric line can be defined by any curve passing, downstream of the first slat 30, through the geometric line formed by the section in the radial plane P of the inner radial face 300 of the first slat 30 and the geometric line formed by the section in the radial plane P of the outer radial face 301 of the first slat 30. Depending on the form of the downstream surface 303 of the first slat, the downstream profile C2 of the first slat can have a different form of geometric line in another radial plane P’, which also comprises the longitudinal axis X-X and is different from the radial plane P.

[0100] Furthermore, the first slat 30 defines a radial span ε corresponding to the difference between the third radius R3 and the fourth radius R4, i.e. to the distance spaced apart from the inner radial face 300 of the first slat 30 and the outer radial face 301 of the first plate portion. The radial span ε is limited by the thickness e of the intake lip 10 so that the radial span can easily be deployed and retracted for changing the geometry of the intake 3. Preferably, the ratio between the radial span ε of the first slat 30 and the thickness e of the intake lip 10 is comprised between 0.4 and 0.6. In any case, the inner radial face 300 of the first slat 30 is as close as possible to the longitudinal axis X-X, i.e. extends substantially in the same plane as the surface of the inner wall 12 of the nacelle 1 which is farthest from the longitudinal axis X-X, as shown in Figure 10 In this way, the airflow channel 34 terminates as close as possible to the outer radial face through which the airflow inside the intake 100 passes. In this way, in the second configuration, the increased airflow injected by the airflow channel 34 has current lines which are as parallel as possible to the current lines of said airflow inside the intake 100.

[0101] Furthermore, the first slat 30 defines a length H1 of the inner radial face of the first slat corresponding to the distance spaced apart from the inner radial end of the upstream profile C1 of the first slat and the inner radial end of the downstream surface of the first slat 30. As shown in Figure 10 The length H1 of the inner radial face of the first slat is determined so that said inner radial face 300 can act as a guide when the device is converted from the first configuration to the second configuration and vice versa. Furthermore, the length H1 of the inner radial face of the first slat is determined so as to guarantee the compactness of the device 3 for changing the geometry of the intake in the retracted position by contact with the inner wall 12 of the nacelle 1. This maintains the same aerodynamic performance in particular regardless of the configuration of the device 3 for changing the geometry of the intake, i.e. in the deployed position, in the retracted position or in an intermediate position between the two.

[0102] Furthermore, the first slat 30 defines a length H2 of the outer radial face of the first slat corresponding to the distance spaced apart from the outer radial end of the upstream profile C1 of the first slat and the outer radial end of the downstream surface of the first slat 30. As shown in Figure 10 The length H2 of the outer radial face of the first slat is determined so that said outer radial face 301 can act as a guide when the device is converted from the first configuration to the second configuration and vice versa. Furthermore, the length H2 of the outer radial face of the first slat is determined so as to guarantee the compactness of the device 3 for changing the geometry of the intake in the retracted position by contact with the outer wall 14 of the nacelle 1. This maintains the same aerodynamic performance in particular regardless of the configuration of the device 3 for changing the geometry of the intake, i.e. in the deployed position, in the retracted position or in an intermediate position between the two.

[0103] Finally, with reference to Figures 4 to 6 , the second slat 32 defines a length H of an inner radial face 320 of the second slat 32 i , which corresponds to the distance by which the inner radial end of the upstream profile C3 of the second slat is spaced apart from the inner radial end of the downstream surface of the second slat 32. This length H i is also equal to the length of an outer radial face 321 of the second slat 32, which corresponds to the distance by which the outer radial end of the upstream profile C3 of the second slat is spaced apart from the outer radial end of the downstream surface of the second slat 32. As shown in these figures, the length H i of the inner and / or outer radial face of the second slat is determined so that the inner and / or outer radial face can act as a guide when the device is converted from the first configuration to the second configuration, and vice versa. Furthermore, the length H i of the inner and / or outer radial face of the second slat is determined so as to guarantee the tightness of the device 3 for changing the geometry of the intake in the retracted position by contact with the inner wall 12 of the nacelle 1. This maintains the same aerodynamic performance regardless of the configuration of the device 3 for changing the geometry of the intake, i.e. in the deployed position, in the retracted position or in an intermediate position between the two.

[0104] Also as Figure 10 shown, the second slat 32 also has an upstream profile C3 of the second slat. The upstream profile C3 of the second slat forms an upstream surface 322 of the second slat, which is three-dimensionally rotated around the longitudinal axis X-X. The section of the downstream profile C2 of the first slat in the radial plane P forms a geometric line which connects the geometric line formed by the section of the inner radial face 320 of the second slat 32 in the radial plane P to the geometric line formed by the section of the outer radial face 321 of the second slat 32 in the radial plane P. This geometric line can be defined by any curve which passes through the geometric line formed by the section of the inner radial face 320 of the second slat 32 in the radial plane P and the geometric line formed by the section of the outer radial face 321 of the second slat 32 in the radial plane P. Depending on the form of the upstream surface 322 of the second slat, the upstream profile C3 of the second slat can have a geometric line of different form in another radial plane P’, which also comprises the longitudinal axis X-X and is different from the radial plane P.

[0105] In any case, in the second configuration, the upstream profile C3 of the second slat has tangential continuity with the profile of the nacelle 1 defined in the radial plane P. More precisely, as Figure 10As shown, the nacelle 1 also has a profile in the radial plane P, which takes the form of two geometric lines respectively corresponding to the inner wall 12 and to the outer wall 14 of the nacelle 1. Now, at the intersection of the upstream profile C3 of the second slat with the profile of the nacelle 1, in the region of the inner wall 12 and of the outer wall 14 respectively, the local tangent of the upstream profile C3 of the second slat merges with the local tangent of the profile of the nacelle 1. In other words, in the second configuration, the second slat 32 precisely matches the respective edges of the inner wall 12 and of the outer wall 14 of the nacelle 1 so as not to disturb the airflow through the intake 100. This ensures that the additional disturbance of the airflow is minimized. This tangential continuity of the second configuration exists in any radial plane and all around the longitudinal axis X-X.

[0106] In a preferred embodiment, for example in Figure 10 As shown, the upstream profile C1 of the first slat is identical to the upstream profile C3 of the second slat. More precisely, in any radial plane P, the geometric line that defines the upstream profile C1 of the first slat is identical to the geometric line that defines the upstream profile C3 of the second slat, i.e. these two geometric lines can be perfectly superimposed over their entire length. In this way, the intake lip 10 has exactly the same form in the first configuration and in the second configuration.

[0107] Still with reference to Figure 10 Regardless of the position of the device 3 for changing the geometry of the intake, the airflow passage 34 has:

[0108] - an internal axial length δ1 defined as the distance separating the downstream end of the inner radial face of the first slat 300 from the upstream end of the inner radial face of the second slat 320, and

[0109] - an external axial length δ2 defined as the distance separating the downstream end of the outer radial face of the first slat 301 from the upstream end of the outer radial face of the second slat 321.

[0110] In an advantageous configuration of the device 3 for changing the geometry of the intake, the internal axial length δ1 of the airflow passage 34 is less than the external axial length δ2 of the airflow passage 34. This defines a converging geometry of the airflow passage 34, as Figure 10 shown.

[0111] Furthermore, in an embodiment in which the first slat 30 is fixedly mounted relative to the second slat 32, as Figure 4 shown, regardless of the position of the device 3 for changing the geometry of the intake, the internal axial length δ1 and the external axial length δ2 are constant.

[0112] Moreover, regardless of the position of the device 3 for varying the geometry of the intake, the device for varying the geometry of the intake has an axial span δ e , which is defined as:

[0113] the sum of the length H1 of the inner radial face of the first slat and the inner axial length δ1 of the airflow channel, or

[0114] the sum of the length H2 of the outer radial face of the first slat and the outer axial length δ2 of the airflow channel, or

[0115] Thereby, the device 3 for varying the geometry of the intake limits the maximum axial span δ to the maximum of the distance between the outer radial end of the upstream profile C1 of the first slat and the outer radial end of the upstream edge of the inner wall 12 of the nacelle 1 and the distance between the inner radial end of the upstream profile C1 of the first slat and the inner radial end of the upstream edge of the inner wall 12 of the nacelle 1 when the device 3 for varying the geometry of the intake is in the deployed position. In other words, the maximum axial span δ corresponds to the axial span δ e of the device 3 for varying the geometry of the intake when the device 3 for varying the geometry of the intake is in the deployed position.

[0116] When said device 3 for varying the geometry of the intake is in the first configuration, i.e. in the retracted position, the space available for the expansion of the device 3 for varying the geometry of the intake is limited within the nacelle 1. In this way, the axial length δ i of the guide rail 380 is preferably equal to the axial span δ e of the device 3 for varying the geometry of the intake in the retracted position and, in any case, is not less than the axial span δ e .

[0117] In any case, the inner axial length δ1 of the airflow channel and the outer axial length δ2 of the airflow channel are each less than the axial span δ e of the device 3 for varying the geometry of the intake. Moreover, advantageously, the inner axial length of the airflow channel and the outer axial length of the airflow channel are each non-zero to guarantee the presence of the airflow channel 34. Moreover, in the embodiment in which the first slat 30 and the second slat 32 can be provided for axial movement independently of each other, the inner axial length δ1 of the airflow channel and the outer axial length δ2 of the airflow channel can take a minimum value depending on the point of contact between the first slat 30 and the second slat 32. This is particularly evident in Figure 5 . As shown in this figure, this makes it possible to obtain an advantageous gain in space in the retracted position.

[0118] Therefore, the airflow channel 34 has, in the radial plane P:

[0119] - an upstream profile corresponding to a downstream profile C2 of the first slat, and

[0120] - a downstream profile corresponding to an upstream profile C3 of the second slat.

[0121] Advantageously, the airflow channel 34 can have a different geometry around the longitudinal axis X-X. In other words, as shown in Figure 11 and Figure 12 , the airflow channel 34 has:

[0122] - in the first radial plane PI:

[0123] o a first upstream profile corresponding to a first downstream profile C21 of the first slat, and

[0124] o a first downstream profile corresponding to a first upstream profile C31 of the second slat, and

[0125] - in the second radial plane P2:

[0126] o a second upstream profile corresponding to a second downstream profile C22 of the first slat, and

[0127] o a second downstream profile corresponding to a second upstream profile C32 of the second slat.

[0128] More precisely, since the second upstream profile C32 of the second slat is different from the first upstream profile C31 of the second slat, and / or the second downstream profile C22 of the first slat is different from the first downstream profile C21 of the first slat, as shown in Figure 11 and Figure 12 , the airflow channel 34 has a first longitudinal section in the first radial plane PI and a second longitudinal section in the second radial plane P2, the second longitudinal section being different from the first longitudinal section. By "longitudinal" section is meant a section of the airflow channel 34 according to a radial plane with respect to the longitudinal axis X-X.

[0129] Figures 11 to 13 The longitudinal sections of the airflow channel 34 that are different in the first configuration of the device 3 for varying the geometry of the intake and in the second configuration of said device 3 are shown respectively. It is clearly seen in each of these figures that the first longitudinal section converges in the first radial plane PI, i.e. the airflow channel narrows as it gets closer to the longitudinal axis X-X, while the second longitudinal section converges-diverges in the second radial plane P2, i.e. the airflow channel first narrows and then increases again as it gets closer to the longitudinal axis X-X, until the airflow channel ends at the intake 100.

[0130] This locally optimizes the aerodynamic effect of the modification of the geometry of the intake 100. This possibility is particularly advantageous when the nacelle 1 is not axisymmetric, as Figures 12 to 13 Indeed, since the upstream profile and / or the downstream profile of the airflow passage 34 defines a converging and / or converging-diverging section of the airflow passage 34 depending on the azimuthal position of the upstream profile and / or of the downstream profile of the airflow passage around the longitudinal axis X-X, it is possible to compensate for the asymmetry of the nacelle 1 by changing the flow of the airflow entering via the intake 100. For example, at takeoff, the stagnation point of the flow is generally located on the outside of the nacelle 1 and the airflow bypasses the intake lip 10 from the outside to the inside over the entire circumference of the nacelle 1. However, when the longitudinal axis X-X is inclined with respect to the horizontal (i.e. the aircraft is subjected to a shock), the stagnation point can be more towards the inside of the nacelle 1 in the second radial plane P2, while it is more towards the outside of the nacelle in the first radial plane PI. The different forms in the two radial planes PI and P2 make it so that in the second radial plane P2, the flow bypasses the intake lip 10 in both directions, while in the first radial plane PI, the flow always bypasses the intake lip 10 in the same direction.

[0131] In this way, it is possible to define the degree of geometric convergence of the airflow passage 34. More precisely, assuming that the internal radial surface Al and the external radial surface A2 of the airflow passage 34 are generatrices of a cylinder with a circular base, the convergence ratio τ of the airflow passage 34 can be defined as the ratio of the internal radial surface Al of the airflow passage 34 to the external radial surface A2 of the airflow passage 34:

[0132]

[0133] With the ratio of the third radius R3 to the fourth radius R4 always greater than 1, the airflow passage 34 is naturally converging. However, the ratio of the internal axial length δ1 of the airflow passage 34 to the external axial length δ2 of the airflow passage influences the convergence ratio τ. Indeed, this ratio is:

[0134]

[0135] In this way, the convergence can be defined by the respective choice of the length H1 of the internal radial face of the first slat and of the length H2 of the external radial face of the first slat, both of which are each limited by the axial span δ of the device 3 for modifying the geometry of the intake.

[0136] In the second configuration, the airflow passage 34 provides an additional change in the amount of movement within the fluid entering via the inlet 100. This change can include acceleration or deceleration of the incoming fluid. In low-speed flows, typically in subsonic fields, acceleration is generally preferred, and the convergence rate τ should be less than 1, meaning the longitudinal section of the airflow passage 34 is convergent. In any case, the structure of the device 3 for changing the geometry of the inlet results in a fairly large convergence rate τ, i.e., close to 0, but also results in a very low convergence rate τ, i.e., close to 1, or even divergent, i.e., where the convergence rate τ is greater than 1.

[0137] Embodiments of a device for changing the geometry of an air intake

[0138] In embodiments where the first longitudinal section of the airflow channel 34 in the first radial plane P1 differs from the second longitudinal section in the second radial plane P2, the axial span δ of the device 3 for changing the geometry of the air inlet is the same in both the first and second planes P1 and P2. Figure 11 and Figure 12 As shown. This ensures simple linear translation of the device 3 used to change the geometry of the air intake.

[0139] like Figure 9 , Figure 13 , Figure 15 and Figure 17 As shown, in the embodiment, the first slat 30 and the second slat 32 have rotational symmetry about the longitudinal axis XX in at least one given angular sector. This makes it easy to deploy and retract the device 3 for changing the geometry of the air intake, and ensures the compactness of the cabin 1 in both the first and second configurations. Furthermore, this minimizes disturbances in the flow entering via the air intake 100. In any case, the upstream profiles C1, C3 of the slats 30, 32 may or may not exhibit rotational symmetry about the longitudinal axis XX in at least one given angular sector. Since the inner radial surfaces 300, 320 and the outer radial surfaces 301, 321 of the slats 30, 32 are the generatrices of cylinders rotating about the longitudinal axis XX, which is also the direction of deployment and retraction, deployment and retraction of the device 3 for changing the geometry of the air intake are facilitated in all cases.

[0140] However, the different angular portions of the first slat 30 and / or the second slat 32 can be independently converted between the first configuration and the second configuration, and vice versa. In other words, the device 3 for changing the geometry of the air inlet can include a plurality of first slats 30 connected to a plurality of second slats 32, each first slat extending according to an angular portion about the longitudinal axis XX, and each first slat translating between the first configuration and the second configuration. In this way, the first slat 30 and the second slat 32 can be in the first configuration, while another first slat 30 and another second slat 32 are in the second configuration. This ensures modular use of the device 3 for changing the geometry of the air inlet according to different flows entering the air inlet 100. Advantageously, the plurality of first slats 30 and the plurality of second slats can also have upstream profiles C1, C3 and downstream profiles C2 that are different from each other.

[0141] Reference Figures 14 to 17 For the 100° angle portion of the air intake, it is usually the upper azimuth angle portion, such as... Figures 14 to 17 As shown, the presence of airflow passage 34 may be unnecessary or even undesirable. In fact, the presence of airflow passage 34 can generate aerodynamic disturbances, increasing the impact of the flow within the regions of these portions of the air intake 100.

[0142] Therefore, in Figure 14 and Figure 15 In the first embodiment shown, it can be ensured that the first slat 30 and the second slat 32 are connected at at least one corner portion of the device 3 for changing the geometry of the air intake, such that the internal axial length δ1 and the external axial length δ2 of the airflow passage 34 are zero. Then, at at least this corner portion, the first slat 30 and the second slat 32 form a single slat 33, which also translates relative to the nacelle 1 according to the longitudinal axis XX.

[0143] exist Figure 16 and Figure 17 In the second embodiment shown, the first slat 30 and the second slat 32 do not extend around the entire longitudinal axis XX, but only extend along the corner portions, for example, on the side and lower portions of the air intake 100. The upper portion of the air intake 100 is defined by an air intake lip 10 fixed relative to the nacelle 1. In other words, in this second embodiment, the first slat 30 and the second slat 32 extend along the first corner portion of the nacelle 1, and the second corner portion of the nacelle is defined by the fixed portion of the air intake lip 10. More precisely, at this second corner portion, the air intake lip 10 does not include means 3 for changing the geometry of the air intake.

Claims

1. An assembly for a gas turbine engine, the assembly comprising: - a nacelle (1) having a longitudinal axis (X-X) and comprising an air intake lip (10) defining an air intake (100), and - means (3) for varying the geometry of the air intake, the means (3) comprising: - a first slat (30), and - a second slat (32), the first slat (30) and the second slat (32) translating relatively to the nacelle (1) according to the longitudinal axis (X-X) between a first configuration and a second configuration: - in the first configuration, the first slat (30) forms the air intake lip (10) and the second slat (32) extends inside the nacelle (1), and - in the second configuration, the first slat (30) extends from the air intake lip (10) by a distance and the second slat (32) forms the air intake lip (10) so as to define a flow passage (34) between a downstream surface of the first slat (30) and an upstream surface of the second slat (32), wherein the first slat (30): - has a first downstream profile (C21) of the first slat in a first radial plane (PI) relative to the longitudinal axis (X-X), and - has a second downstream profile (C22) of the first slat in a second radial plane (P2) relative to the longitudinal axis (X-X), - the second downstream profile (C22) of the first slat is different from the first downstream profile (C21) of the first slat so that the flow passage (34) has a first longitudinal section in the first radial plane (PI) and a second longitudinal section in the second radial plane (P2), the second longitudinal section being different from the first longitudinal section.

2. An assembly for a gas turbine engine as recited in claim 1, wherein, the second slat (32) has an upstream profile (C3) of the second slat in a radial plane (P) relative to the longitudinal axis (X-X), the upstream profile (C3) of the second slat having a tangential continuity with a profile of the nacelle (1) defined in the radial plane (P) in the second configuration.

3. The assembly for a gas turbine engine of claim 1 or 2, wherein, the first slat (30) has an upstream profile (CI) of the first slat in a radial plane (P) relative to the longitudinal axis (X-X), the second slat (32) having an upstream profile (C3) of the second slat in the radial plane (P), the upstream profile (C3) of the second slat being identical to the upstream profile (CI) of the first slat.

4. An assembly for a gas turbine engine as recited in claim 1 or 2, wherein, the second slat (32): - has a first upstream profile (C31) of the second slat in a first radial plane (PI) relative to the longitudinal axis (X-X), and - has a second upstream profile (C32) of the second slat in a second radial plane (P2) relative to the longitudinal axis (X-X), - said second upstream profile (C32) of said second slat is different from said first upstream profile (C31) of said second slat, so that said airflow passage (34) has a first longitudinal section in said first radial plane (PI) and a second longitudinal section in said second radial plane (P2), said second longitudinal section being different from said first longitudinal section.

5. An assembly for a gas turbine engine as recited in either of claims 1 or 2, wherein, On at least one corner portion of said device (3) for varying the geometry of said air intake, said first slat (30) and said second slat (32) are connected to form a single slat (33) which is also translationally mobile with respect to said nacelle (1) according to said longitudinal axis (X-X).

6. An assembly for a gas turbine engine as recited in either of claims 1 or 2, wherein, Said first slat (30) and said second slat (32) extend on a first corner portion of said nacelle (1), a second corner portion of said nacelle (1) being delimited by a fixed portion of said air intake lip (10).

7. An assembly for a gas turbine engine as recited in claims 1 or 2, wherein, Said device (3) for varying the geometry of said air intake comprises a guide rod (36) which extends parallel to said longitudinal axis (X-X) and connects said first slat (30) to said second slat (32), so that on at least one corner portion of said device (3) for varying the geometry of said air intake, said first slat (30) is fixedly mounted with respect to said second slat (32).

8. An assembly for a gas turbine engine as recited in claims 1 or 2, wherein, Said device (3) for varying the geometry of said air intake comprises a guide rod (36) which extends parallel to said longitudinal axis (X-X), said guide rod (36) being telescopic so that on at least one corner portion of said device (3) for varying the geometry of said air intake, the movement of said first slat (30) is independent of the movement of said second slat (32).

9. Assembly for a gas turbine engine according to claim 7, said assembly comprising an anti-icing system of said first slat (30) and / or of said second slat (32), said anti-icing system comprising an anti-icing fluid flow passage which extends at least partially inside said guide rod (36).

10. An assembly for a gas turbine engine as recited in either of claims 1 or 2, wherein, Said device (3) for varying the geometry of said air intake further comprises an actuator which is housed in said nacelle (1) and is configured to move said first slat (30) and said second slat (32) between said first configuration and said second configuration.

11. An assembly for a gas turbine engine as recited in claims 1 or 2, wherein, Said nacelle (1) has a non-axially symmetric shape.

12. A gas turbine engine comprising an assembly according to any one of claims 1 to 11.

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