Nacelle inlet for an aircraft propulsion assembly for facilitating a thrust-reversal phase

Abstract

An air intake (1) for an aircraft propulsion assembly (8) is disclosed, which extends along a longitudinal axis (X) and comprises a turbojet engine (7) comprising a primary flow path (4) and a secondary flow path (5) for guiding, respectively, a primary flow (F1) and a secondary flow during thrust, and a thrust reverser device for changing the secondary flow into a reversed flow (F-INV) during thrust reversal (B), the air intake (1) comprising a peripheral casing (11) comprising, in each plane radial to the axis (X), a point of maximum curvature (P) of the reversed flow (F-INV), a osculating circle of the radius of curvature defined at each point of maximum curvature (P), the average value (Rmoy) of the radii of curvature satisfying: Rmoy < 0.028*R3, with R3 being the internal radius of the air intake (1).

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aircraft propulsion assemblies, and more specifically to a nacelle inlet for an aircraft propulsion assembly. BACKGROUND

[0002] With reference to figure 1A, in a known manner, an aircraft propulsion assembly 800 extends along a longitudinal axis X from upstream to downstream, and comprises a turbine engine 700 and a nacelle 200. The turbine engine 700 extends along the longitudinal axis X, and is configured to allow propulsion of the aircraft by acceleration of an internal airflow F-INT circulating in the turbine engine 700 from upstream to downstream. On the other hand, the nacelle 200 extends outwardly along the longitudinal axis X around the turbine engine 700, and is able to channel the internal airflow F-INT in the turbine engine 700. The terms "upstream" and "downstream" are then defined with respect to the direction of the longitudinal axis X. The terms "internal" and "external" are in turn defined with respect to the longitudinal axis X along a radial direction.

[0003] In a known manner, as illustrated in figure 1A, the turbine engine 700 is a bypass turbine engine, and comprises, upstream, a fan 300 rotatably mounted around the longitudinal axis X to accelerate the internal airflow F-INT from upstream to downstream. Downstream of the fan 300, the turbine engine 700 also comprises a radially internal primary flowpath 400 and a radially external secondary flowpath 500, separated by a casing 600. The casing 600 is configured to channel a first portion of the internal airflow F-INT, called primary airflow Fl, into the primary flowpath 400 for fuel combustion, and to channel a second portion of the internal airflow F-INT, called secondary airflow F2, into the secondary flowpath 500, so as to generate thrust of the turbine engine 700.

[0004] Still with reference to figure 1A, in a known manner, the nacelle 200 extends radially outwardly to the fan 300, and radially outwardly delimits the secondary flowpath 500. The nacelle 200 comprises, at its upstream end, an inlet 100 comprising a peripheral shell 110 delimiting an annular inner cavity 120. The peripheral shell 110 comprises an inner wall 130 turned towards the longitudinal axis X, and an outer wall 140 opposite the inner wall 130, which are connected together upstream by an inlet lip 150 comprising a leading edge. The inlet 100 has an aerodynamic circular profile, which makes it possible to divide an upstream airflow F into an internal airflow F-INT channeled by the inner wall 130, and an external airflow F-EXT channeled by the outer wall 140.

[0005] With reference to FIG. 1B , in order to reduce the braking distance of an aircraft, particularly during landing, it is known to change the direction of the airflow in secondary flow path 500 to implement a thrust reversal phase B. Hereinafter, a distinction is made between thrust phase A ( FIG. 1A ), in which secondary airflow F2 circulates from upstream to downstream in secondary flow path 500, and thrust reversal phase B ( FIG. 1B ), in which reverse airflow F-INV circulates from downstream to upstream. It is specified that during thrust reversal phase B, internal airflow F-INT from upstream airflow F circulates from upstream to downstream at the root of fan 300 to supply primary airflow F1, as is also the case during thrust phase A. Primary airflow F1 may also be supplied by a portion of reverse airflow F-INV that bypasses casing 600.

[0006] To perform the thrust reversal phase, patent application FR 2 120 172 A1 discloses at least partially blocking the secondary flow path 500 downstream of the fan 300 and jointly exposing a grille housed in the nacelle 200, so as to generate an inverted airflow F-INV directed opposite to the secondary airflow F2. However, this thrust reversal system has drawbacks that are detrimental to the mass, dimensions and drag of the aircraft propulsion assembly 800, in particular for large-diameter nacelles 200 used in high-dilution aircraft propulsion assemblies, i.e., for which the ratio of the mass of the secondary airflow F2 to the mass of the primary airflow F1 is greater than 16, in particular greater than 20.

[0007] Referring to FIG. 1B , for high-dilution aircraft propulsion assemblies, a variable-pitch fan 300, abbreviated as "VPF," is known. This fan includes blades whose pitch angle is controlled to reverse the direction of airflow circulation in the secondary flowpath 500. In practice, during the thrust reversal phase B, the reverse airflow F-INV circulates from downstream to upstream in the secondary flowpath 500, then passes through the fan 300 and is directed upstream by the upstream inner wall 110 of the air inlet 100. The reverse airflow F-INV then moves in the opposite direction of the upstream airflow F, particularly radially outward near the nacelle 200, which aids braking. The reverse airflow F-INV then merges with the external airflow F-EXT, which is allowed to enter the secondary flowpath 500 through the downstream end of the nacelle 200.

[0008] In fact, it can be observed that when the reverse airflow F-INV merges with the external airflow F-EXT, the reverse airflow F-INV remains attached to the outer shell 110 of the air inlet 100. In other words, it flows along the contour of the air inlet 200 without separating from the outer shell 110. This phenomenon is known as the "Coanda effect." This phenomenon tends to accelerate the reverse airflow F-INV at the air inlet 100, creating a local low pressure DP, thereby generating a force opposing the thrust reversal phase. This phenomenon can reduce the performance of the turbine engine 700 during the thrust reversal phase, which is undesirable.

[0009] To improve the performance of turbine engine 700 during the thrust reversal phase, patent applications FR1904087A1 and FR1904094A1 disclose the movable installation of deflectors and straightening vanes, respectively, on the air inlet 100, which are deployed during thrust reversal phase B and retracted during thrust phase A. Patent applications FR1904089A1 and FR1904096A1 alternatively teach making a portion of the air inlet 100 movable or elastically deformable during thrust reversal phase B. Patent application FR1904092A1 further teaches the formation of an internal duct in the air inlet 100 that opens during thrust reversal phase B to divert a portion of the reverse airflow F-INV. All of these solutions advantageously allow the circular profile of the air inlet to be modified only during thrust reversal phase B to prevent the formation of localized low pressure DP, without compromising performance during thrust phase A. However, this solution has the disadvantage of requiring a complex, expensive, and actuated dual-profile air inlet, which makes it prone to failure.

[0010] Incidentally, in order to improve the resistance during the propulsion phase, an air inlet is known from application US 10399687 B2, which includes a through-hole opening controlled by a movable cover member. This solution has the same drawbacks as the previous ones.

[0011] The invention thus relates to an air intake 100 of a nacelle 200 of an aircraft propulsion assembly 800 making it possible to increase the performance of a turbine engine 700 during thrust phases A and thrust reversal phases B, with a simple and robust structure. Summary of the Invention

[0012] The invention relates to an air intake for a nacelle of an aircraft propulsion assembly, the nacelle extending along a longitudinal axis oriented from upstream to downstream and comprising a turbine engine, the turbine engine comprising a radially inner primary flow path and a radially outer secondary flow path, the radially inner primary flow path and the radially outer secondary flow path being configured to guide the primary and secondary air flows, respectively, from upstream to downstream during a thrust phase, the turbine engine comprising, upstream, a fan mounted rotatably about the longitudinal axis, the aircraft propulsion assembly comprising a thrust reversal device configured to change the secondary air flow into a reversed air flow circulating in the secondary flow path from downstream to upstream during a thrust reversal phase, the nacelle extending outwardly around the turbine engine and comprising, at its upstream end, an air intake, the air intake comprising a peripheral casing oriented along the longitudinal axis, the casing comprising an inner wall turned toward the longitudinal axis and an outer wall opposite the inner wall, the latter being connected upstream by an air intake lip, the air intake lip comprising a leading edge for promoting the thrust phase, the air intake comprising an inner radius at the fan, referred to as the “fan radius R3”, the peripheral casing comprising a variable curvature in a plane radial to the longitudinal axis.

[0013] The notable feature of the present invention is that the outer shell of the air inlet includes a maximum curvature point in each radial plane so as to separate the reverse airflow from the outer shell, thereby facilitating the thrust reversal phase, an osculating circle including a curvature radius is defined at each maximum curvature point, and an average value Rmoy of the curvature radius of the outer periphery of the air inlet satisfies the following relationship: Rmoy<0.028*R3.

[0014] By geometric definition, the osculating circle at a point on a curve corresponds to the circle that is tangent to the curve at that point and that follows the curve as closely as possible. Considering that the outer shell describes a curve in each radial plane along the longitudinal axis, hereinafter referred to as the "inlet curve," the osculating circle at the point of maximum curvature of the inlet curve represents the circle that is tangent to the inlet curve at that point and that follows the inlet curve as closely as possible. The radius of curvature represents the radius of the osculating circle, i.e., the straight line segment connecting its center and circumference.

[0015] A mathematical definition of the radius of curvature R at the point of maximum curvature P of the intake curve defined by the parametric equation in Cartesian coordinates x(P), y(P) is also provided:

[0016] [Operation 1]

[0017] Thanks to the invention, the inlet includes a constant profile which makes it possible to direct the flow effectively during the thrust phase and the thrust reversal phase, which improves the performance of the aircraft turbine engine. More precisely, the profile of the inlet includes on the one hand a circular upstream end and on the other hand a set of maximum curvature points. The circular upstream end advantageously makes it possible to separate the upstream flow into an internal flow which feeds the aircraft turbine engine and an external flow during the propulsion phase. The maximum curvature points allow the reversal flow to be separated from the peripheral casing during the thrust reversal phase.

[0018] According to a first aspect of the invention, the peripheral casing of the inlet has no moving parts. According to another aspect independent of the first aspect, the peripheral casing is non-deformable. Such a fixed and non-deformable profile inlet advantageously has a simple, robust, durable and economical structure.

[0019] According to an aspect of the invention, the average of the radii of curvature satisfies the following relationship: Rmoy < 0.02 * R3, and preferably the following relationship: Rmoy < 0.01 * R3. In other words, the inlet includes very pronounced curvatures, i.e. small radial thicknesses, at the maximum curvature points, which promote the separation of the reversal flow.

[0020] According to a preferred aspect of the invention, the average of the radii of curvature satisfies the following relationship: Rmoy > 0.005 * R3. This guarantees a minimum radial thickness of the inlet to preserve the performance of the thrust phase.

[0021] According to an aspect of the invention, the maximum curvature points are located in an upstream portion of the peripheral casing which includes a longitudinal length L16 defined from the leading edge of the inlet lip and which satisfies the following relationship: L16 < 6 * Rmoy, preferably: L16 < 4 * Rmoy, and preferably: L16 < 2 * Rmoy. In other words, the longitudinal length is defined between the upstream end of the inlet, i.e. the leading edge, and a variable downstream point, depending on the average of the radii of curvature. Thus, the maximum curvature points are concentrated in the vicinity of the inlet lip. This advantageously makes it possible to start the separation of the reversal flow in the vicinity of the inlet lip in order to separate it from the outer wall and this in order to prevent any local low pressure which would be detrimental to the thrust reversal.

[0022] According to an aspect, the maximum curvature points are located on the leading edge of the inlet lip, which is the best point to start the separation of the reversal flow.

[0023] According to an aspect of the invention, the maximum curvature points together form a closed curve, called "separation curve", preferably belonging to a transverse plane on the longitudinal axis, preferably in the form of a circle centered on the longitudinal axis. In other words, the maximum curvature points extend continuously with respect to each other and all preferably belong to the same plane for separating the reversal flow globally and uniformly over the entire profile of the peripheral casing. The circular separation curve corresponds to an axisymmetric inlet and allows a perfectly uniform separation.

[0024] Preferably, the value of the radius of curvature of the osculating circle defined at two consecutive points of maximum curvature of the separation curve does not vary by more than 10%, preferably by more than 5%. In other words, the inlet lip has a radial thickness that gradually varies over its circumference, providing for uniform separation of the opposing air flows. Thin and thick areas are not adjacent to each other.

[0025] According to one aspect of the present invention, the value of the radius of curvature of the osculating circle defined in at least 25% of the points of maximum curvature is substantially the same, preferably at least 50% of the points of maximum curvature, and preferably at all points of maximum curvature. In other words, the inlet lip has a substantially constant radial thickness around its circumference, providing for uniform separation of the opposing air flows. Specifically, the expressions "substantially the same" or "substantially constant," as used herein and throughout this patent application, permit deviations of up to 10%.

[0026] The invention also relates to an assembly of an air intake for an aircraft propulsion assembly and a turbine engine fan, the assembly being rotatably mounted about a longitudinal axis, the air intake extending outwardly around a fan comprising a plurality of variable-pitch blades, thereby forming a thrust reversal device for the aircraft propulsion assembly. This advantageously facilitates the thrust reversal phase, in particular by avoiding the use of a grid system housed in the nacelle, which increases mass and size.

[0027] Preferably, the fan comprises a compression ratio, defined as the ratio of the pressure at the fan outlet to the pressure at the fan inlet, less than 1.4, in order to optimize the operation of the turbine engine with a high dilution ratio.

[0028] Preferably, the air inlet has a length L1, which is defined as the longitudinal distance separating the leading edge of the air inlet lip and the upstream edge of the fan blade, and satisfies the following relationship: L1 < 0.8 * R3, preferably the following relationship: L1 < 0.6 * R3, and preferably the following relationship: L1 < 0.4 * R3. With a short air inlet, the phenomenon of reverse airflow separating from the outer shell is more likely to occur, which is advantageous.

[0029] The invention also relates to a nacelle for an aircraft propulsion assembly, the aircraft propulsion assembly extending along a longitudinal axis oriented from upstream to downstream and comprising a turbine engine, the turbine engine comprising a radially inner primary flow path and a radially outer secondary flow path, the radially inner primary flow path and the radially outer secondary flow path being configured to guide the primary airflow and the secondary airflow, respectively, from upstream to downstream during a thrust phase, the turbine engine comprising a fan mounted rotatably about the longitudinal axis upstream, the aircraft propulsion assembly comprising a thrust reversal device configured to change the secondary airflow into a reverse airflow circulating in the secondary flow path from downstream to upstream during a thrust reversal phase, the nacelle extending outwardly around the turbine engine and comprising an air inlet as described above at its upstream end.

[0030] Preferably, the nacelle comprises a maximum outer radius Rmax, defined as the maximum radial distance from its center to its outer wall, which satisfies the following relationship: Rmax<1.2*R3, preferably: Rmax<1.15*R3, and preferably: Rmax<1.1*R3. In other words, such a nacelle comprises a small radial thickness, which makes it possible to facilitate the separation of the reverse airflow from the outer casing.

[0031] Preferably, the nacelle comprises a length L2, defined as the length of the outer shell of the nacelle, which satisfies the following relationship: L2<3*R3, and preferably the following relationship: L2<2.5*R3, and preferably the following relationship: L2<2*R3. A short-length nacelle can advantageously be provided with a short air inlet and comprise a small radial thickness.

[0032] The present invention also relates to an aircraft propulsion assembly extending along a longitudinal axis oriented from upstream to downstream and comprising a turbine engine and a nacelle, the turbine engine comprising a radially inner primary flow path and a radially outer secondary flow path, the primary flow path and the secondary flow path being configured to guide the primary and secondary flows, respectively, from upstream to downstream during a thrust phase, the turbine engine comprising an upstream fan rotatably mounted about the longitudinal axis, the aircraft propulsion assembly comprising a thrust reverser configured to change the secondary flow into a reverse flow that circulates from downstream to upstream in the secondary flow path during a thrust reversal phase, the nacelle extending outwardly around the turbine engine and comprising an air inlet as described above at its upstream end, the aircraft propulsion assembly preferably comprising a dilution ratio greater than 20. The specified dilution ratio is defined as the ratio of the mass of the secondary flow to the mass of the primary flow. Such a propulsion assembly having a very high dilution ratio advantageously improves propulsion efficiency, thereby improving the performance of the aircraft turbine engine during the thrust phase.

[0033] The present invention also relates to a method of using a cabin air inlet in an aircraft propulsion assembly as described above, wherein:

[0034] - During the propulsion phase, the intake lip divides the upstream airflow circulating from upstream to downstream into an external airflow guided by the outer wall and an internal airflow guided by the inner wall, and

[0035] -During the thrust reversal phase, the point of maximum curvature separates from the outer shell and the reverse airflow circulates from downstream to upstream in the air inlet to join the external airflow, thereby promoting the thrust reversal phase.

[0036] This method advantageously eliminates the need for any steps to move and / or deform the air inlet, which retains its shape during the thrust reversal phase, resulting in a simple, robust, durable and economical structure. It is also quick and easy to implement, with no delay in switching from one phase to another.

[0037] The present invention also relates to a method of using an aircraft propulsion assembly as described above, wherein:

[0038] - During the propulsion phase, the intake lip divides the upstream airflow circulating from upstream to downstream into an external airflow guided by the outer wall and an internal airflow guided by the inner wall, and

[0039] -During the thrust reversal phase, the thrust reversal device is activated, the point of maximum curvature separates from the outer shell, and the reverse airflow circulates from downstream to upstream in the air inlet to join the external airflow, thereby promoting the thrust reversal phase.

[0040] Preferably, during the thrust reversal phase, the pitch of the fan blades is changed to form a reverse airflow. BRIEF DESCRIPTION OF THE DRAWINGS

[0041] The invention will be better understood on reading the following description, given by way of example, and with reference to the following drawings, given by way of non-limiting example, in which like references indicate similar objects.

[0042] FIG. 1A is a longitudinal half-sectional view of an aircraft propulsion assembly according to the prior art during a propulsion phase; FIG. 1B is a longitudinal half-sectional view of the aircraft propulsion assembly of FIG. 1A during a thrust reversal phase;

[0043] [ FIG. 2A ] FIG. 2A is a longitudinal half-section view of an aircraft propulsion assembly according to one embodiment of the present invention during a propulsion phase;

[0044] [ FIG. 2B ] FIG. 2B is a longitudinal half-section view of the aircraft propulsion assembly of FIG. 2A during a thrust reversal phase;

[0045] [Figure 3] Figure 3 is a longitudinal half-section view of the aircraft propulsion assembly of Figure 2A;

[0046] [FIG. 4] FIG. 4 is a radial half-section view of the air inlet of the propulsion assembly of FIG. 2A;

[0047] [Figure 5] Figure 5 is a perspective view of the air inlet of Figure 4;

[0048] [Figure 6] Figure 6 is a cross-sectional view of the air intake of Figure 4;

[0049] [ FIG. 7 ] FIG. 7 is a radial half-section view of an air inlet of an aircraft propulsion assembly according to an alternative embodiment of the present invention;

[0050] [Figure 8] Figure 8 is a perspective view of the air inlet of Figure 7;

[0051] [FIG. 9] FIG. 9 is a perspective view of an air intake of an aircraft propulsion assembly according to an alternative embodiment of the present invention; and

[0052] [ Fig. 10] Fig. 10 is a cross-sectional view of an air intake of an aircraft propulsion assembly according to another alternative embodiment of the present invention.

[0053] It should be noted that the accompanying drawings illustrate the invention in detail in order to implement the invention and that said drawings can of course be used to better define the invention if necessary. DETAILED DESCRIPTION

[0054] 2A and 2B , the present invention relates to an innovative aerodynamic aircraft propulsion assembly 8 , in particular at the air inlet 1 , to improve performance during thrust phases A and thrust reversal phases B.

[0055] As shown in FIG2A and described in the introduction, an aircraft propulsion assembly 8 extends from upstream to downstream along a longitudinal axis X and includes a turbine engine 7 and a nacelle 2. The turbine engine 7 extends along the longitudinal axis X and is configured to propel the aircraft by accelerating an internal airflow F-INT that circulates from upstream to downstream within the turbine engine 7. The nacelle 2 extends outwardly around the turbine engine 7 along the longitudinal axis X and is capable of directing the internal airflow F-INT within the turbine engine 7. Subsequently, the terms "upstream" and "downstream" are defined relative to the direction of the longitudinal axis X. The terms "inner" and "outer" are, in turn, defined radially relative to the longitudinal axis X.

[0056] As shown in FIG2A and described in the introduction, the turbine engine 7 is of the bypass type and includes, upstream, a fan 3 rotatably mounted about a longitudinal axis X to accelerate an internal air flow F-INT from upstream to downstream. Downstream of the fan 3, the turbine engine 7 also includes a radially inner main flow path 4 and a radially outer secondary flow path 5, which are separated by a casing 6. The casing 6 is configured to guide a first portion of the internal air flow F-INT (referred to as the main flow F1) into the main flow path 4 for fuel combustion, and to guide a second portion of the internal air flow F-INT (referred to as the secondary flow F2) into the secondary flow path 5 in order to generate the thrust of the turbine engine 7.

[0057] Still referring to FIG. 2A , as described in the introduction, the nacelle 2 extends radially outward to the fan 3 and defines a secondary flow path 5 radially outward. At its upstream end, the nacelle 2 includes an air inlet 1 comprising an outer casing 11 defining an annular inner cavity 12. The outer casing 11 comprises an inner wall 13 directed toward the longitudinal axis X and an outer wall 14 opposite the inner wall 13, which are connected upstream by an air inlet lip 15 including a leading edge. The air inlet 1 has an aerodynamically circular profile, which allows the upstream airflow F to be divided into an inner airflow F-INT guided by the inner wall 13 and an outer airflow F-EXT guided by the outer wall 14.

[0058] With reference to figure 2B and as mentioned in the preamble, the fan 3 is of the variable-pitch type, shortly "VPF", i.e. it comprises blades whose pitch angle is controlled so as to reverse the circulation direction of the secondary flow F2 in the secondary flow path 5. In the following, a distinction is made between a thrust phase A (figure 2A) in which the secondary flow F2 circulates in the secondary flow path 5 from upstream to downstream, and a thrust-reversal phase B (figure 2B) in which a reverse flow F-INV circulates from downstream to upstream in order to reduce the braking distance of the aircraft, in particular during landing. Indeed, during the thrust-reversal phase B, the reverse flow F-INV circulates in the secondary flow path 5 from downstream to upstream, then passes through the fan 3 and is directed upstream by the inner wall 11 of the inlet 1 upstream. The reverse flow F-INV then opposes the upstream flow F, in particular radially outwards near the nacelle 2, so as to act as a brake. The reverse flow F-INV then merges with the external flow F-EXT which allows the passage of the secondary flow path 5 through the downstream end of the nacelle 2. It is provided that, during the thrust-reversal phase B, the internal flow F-INT from the upstream flow F flows from upstream to downstream at the root of the fan 3, to supply the primary flow F1, as during the thrust phase A. The primary flow F1 can also be supplied by a portion of the reverse flow F-INV which bypasses the casing 6.

[0059] Other thrust-reversal systems can be used as an alternative to the variable-pitch fan 3. In particular, it can be provided to at least partially obstruct the secondary flow path 5 downstream of the fan 3 and to jointly expose a grid housed in the nacelle 2 so as to form the reverse flow F-INV. In particular, a grid or flap reverser can be used.

[0060] According to the application and with reference to figures 2A and 2B, in order to improve the performance of the aircraft propulsion assembly 8 during the thrust phase A and the thrust-reversal phase B, the peripheral casing 11 of the inlet 1 comprises an adapted curvature:

[0061] - during the thrust phase A (figure 2A), the upstream flow F is divided into an internal flow F-INT and an external flow F-EXT, and

[0062] - during the thrust-reversal phase B (figure 2B), the reverse flow F-INV is separated from the peripheral casing 11 when it merges with the external flow F-EXT.

[0063] In other words, during the thrust reversal phase B, the outer casing 11 is configured to create a separation zone ZD ( FIG. 2B ) at the outer wall 14 , wherein the reverse airflow F-INV, instead of flowing from upstream to downstream along the contour of the outer wall 14 , has zero or even opposite velocity and forms a vortex. The reverse airflow F-INV generally flows through the separation zone ZD and merges with the external airflow F-EXT. This air inlet 1 advantageously makes it possible to prevent the occurrence of the Coanda effect, which is characterized by the local acceleration of the reverse airflow F-INV that remains attached to the outer wall 14, which generates a local low pressure DP that causes an undesirable force opposite to the thrust reversal B. The air inlet 1 preferably has no moving parts and is non-deformable, which makes it durable and economical compared to air inlets that include moving parts or a deformable casing.

[0064] With reference to FIG3 , in addition to the curvature of the outer casing 11 of the air inlet 1 , the present invention also includes other geometric criteria that the aircraft propulsion assembly 8 must meet in order to improve its performance in Phases A and B. The first criterion is based on the dilution ratio of the turbine engine 7 , which is preferably greater than 16, preferably greater than 20. With reference to FIG2A , the dilution ratio corresponds to the ratio of the mass of the secondary flow F2 to the mass of the primary flow F1 , and for a bypass turbine engine 7 , this dilution ratio is typically between 3 and 12. In other words, due to the acceleration of the secondary flow F2 , the turbine engine 7 has a very high dilution ratio and generates the majority of its thrust, in fact exceeding 80%. This high dilution ratio also makes it possible to generate a considerable reverse flow F-INV. It is stated that this turbine engine 7 with a very high dilution ratio is preferably associated with a more efficient variable-pitch fan 3 , without compromising the aircraft's mass and dimensions.

[0065] 3 , the second criterion relates to the compression ratio of the fan 3 , which is preferably less than 1.4. The compression ratio is defined as the ratio of the pressure at the fan outlet 3 to the pressure at the fan inlet 3 . The following independent standards relate to the nacelle 2 , including:

[0066] The length L2, defined as the length of the outer shell 11 of the nacelle 2 along the longitudinal axis X, preferably satisfies the following: L2 < 3*R3, where R3 here and in the future denotes the inner diameter of the air inlet 1 at the fan 3, referred to as the "fan radius." Furthermore, the length L2 preferably satisfies the following: L2 < 2.5*R3, and preferably: L2 < 2*R3.

[0067] The length L1 of the air inlet 1 defined between the leading edge of the air inlet lip 15 and the leading edge of the fan 3 preferably satisfies: L1<0.8*R3. Further preferably, the length L1 satisfies: L1<0.6*R3, and preferably: L1<0.4*R3.

[0068] The maximum outer radius Rmax of the nacelle 2 preferably satisfies: Rmax<1.2*R3. Further preferably, the maximum outer radius Rmax satisfies: Rmax<1.15*R3, and preferably: Rmax<1.1*R3.

[0069] In other words, these standards require that the nacelle 2 and the air inlet 1 have a thin radial thickness, which facilitates the separation of the reverse airflow F-INV during the thrust reversal phase B. For optimal performance, it is best to take into account all the published standards.

[0070] The curvature of the peripheral casing 11 of the air inlet 1 is defined more precisely below.

[0071] According to the invention, with reference to Figures 2A, 2B, and 3, the outer casing 11 describes a curve in each plane radial to the longitudinal axis X, subsequently referred to as the "inlet curve," which includes a variable curvature, particularly a point of maximum curvature P, i.e., the point where the curvature is most pronounced. Still according to the invention, each point of maximum curvature P is characterized by a radius of curvature, the average value Rmoy of which, for all points of maximum curvature P, satisfies: Rmoy < 0.028 * R3, where R3 is the fan radius. The radius of curvature of a point on the curve is defined as corresponding to the radius of the osculating circle at that point on the curve, i.e., the radius of the circle tangent to the curve at that point, which follows the curve as closely as possible.

[0072] In the example of FIG4 showing the radial half-section of the air inlet 1, the maximum curvature point P of the air inlet curve and the osculating circle C with a curvature radius R are shown. Also shown are two other points Z and Z' of the air inlet curve and their curvature radii R. Z 、R Z’ Oscillating circle C Z 、C Z’ As shown in FIG4 , the point of maximum curvature P corresponds to the point of the intake curve including the minimum curvature radius R, in particular, the curvature radius R smaller than that of points Z and Z′. Z 、R Z’ .

[0073] This point of maximum curvature P advantageously forms the separation point for the reverse airflow F-INV, which is directed from downstream to upstream by the inner wall 13, since its small radius of curvature tends to separate the reverse airflow F-INV from the outer shell 11. Thus, the inlet lip 15 and the point of maximum curvature P respectively facilitate the thrust phase A and the thrust reversal phase B. In fact, the inlet lip 15, due to its circular profile, facilitates the separation of the upstream airflow F into the inner airflow F-INT and the outer airflow F-EXT, while the point of maximum curvature P facilitates the separation of the reverse airflow F-INV.

[0074] In the example of FIG. 4 , the point of maximum curvature P is located at the upstream end of the air inlet 1, converging with the leading edge of the air inlet flange 15 . Advantageously, this point of maximum curvature P makes it possible to create a separation zone ZD ( FIG. 2B ) between the air inlet 1 and the upstream end of the outer wall 14 , where the Coanda effect can occur. However, it goes without saying that the point of maximum curvature P can be located elsewhere on the conical outer shell 11 , as will be seen later. In practice, regardless of the location of the point of maximum curvature P, the outer shell 11 comprises a blended profile compatible with both the thrust phase A and the thrust reversal phase B. In practice, this results in an average radius of curvature Rmoy that is sufficiently large for the thrust phase A, preferably satisfying: Rmoy>0.005*R3, and sufficiently small for the thrust reversal phase B, preferably satisfying: Rmoy<0.02*R3, and preferably: Rmoy<0.01*R3. Preferably, each radius of curvature R is greater than 0.005*R3 and preferably less than 0.05*R3. More preferably, each radius of curvature is smaller than 0.02*R3, preferably 0.01*R3.

[0075] Referring to FIG. 5 , which shows the air intake 1 of FIG. 4 in perspective, all points of maximum curvature P are located on the leading edge of the air intake lip 15 and together define a closed curve, referred to as the "separation curve Q." In this example, since the air intake 1 is axisymmetric, the separation curve Q takes the form of a circle whose center intersects the longitudinal axis X and extends in a plane transverse to the longitudinal axis X. This positioning of the points of maximum curvature P, i.e., their continuous extension and their location in the same plane transverse to the longitudinal axis X, advantageously allows for comprehensive and uniform separation of the opposing airflow F-INV over the entire circumference of the air intake 1. It goes without saying that the points of maximum curvature P can be positioned differently from one another, as will be seen later.

[0076] Referring to FIG. 6 , it is shown that, for the air inlet 1 of FIG. 4 , all points of maximum curvature P comprise the same radius of curvature R in a plane YY ( FIG. 5 ) transverse to the longitudinal axis X. In other words, the radius of curvature R at each point of maximum curvature P is equal to the average radius of curvature Rmoy. This results in an annular inner cavity 12 having a constant radial thickness E12 over the entire circumference of the air inlet 1 in the transverse plane YY. This advantageously allows for uniform separation over the entire circumference of the air inlet 1. It goes without saying that the radius of curvature R can vary from one point of maximum curvature P to another, as will be seen later.

[0077] Figure 7 shows an alternative embodiment of the application in which the points of maximum curvature P' are separated from the leading edge of the intake lip 15, i.e. in this example on the inner wall 13 close to the intake lip 15. The points of maximum curvature P' located on the inner wall 13 allow the reverse airflow F-INV to be pre-separated before the intake lip 15. Alternatively, the points of maximum curvature P' can also be located on the outer wall 14 for separating the reverse airflow F-INV later. In fact, for each intake curve, the points of maximum curvature P' are preferably located on the upstream portion 16 of the peripheral casing 11 which extends along the longitudinal axis X from the leading edge of the intake lip 15 by a longitudinal length L16 which satisfies: L16 < 6 * Rmoy. Preferably, the upstream portion 16 is reduced and satisfies: L16 < 4 * Rmoy, and preferably: L16 < 2 * Rmoy. In other words, the points of maximum curvature P' are preferably located close to the intake lip 15 to create a separation zone ZD in contact with the intake lip 15 and to prevent any occurrence of a Coanda effect.

[0078] In the example of Figure 8, all the points of maximum curvature P' are located on the inner wall 13 as in the example of Figure 7 and belong to the same plane perpendicular to the longitudinal axis x. Such points of maximum curvature P' describe a separation curve Q' extending in a plane transverse to the longitudinal axis X downstream of the intake lip 15, which in this example is circular with a diameter smaller than that of the intake lip 15, and therefore smaller than that of the separation curve Q of Figure 6. Such separation curves Q, Q' allow the reverse airflow to be uniformly separated over the entire circumference of the intake 1.

[0079] With reference to Figure 9, by way of example, for a non-axisymmetric intake 1 or to promote thrust phasing A, points of maximum curvature P" can be provided which are differently positioned with respect to each other on the intake curve thereof. In this example, certain points of maximum curvature P" are thus located on the intake lip 15, while others extend onto the outer wall 14. Preferably, the points of maximum curvature P" extend continuously with respect to each other, thus forming a separation curve Q" extending in a plane not transverse to the longitudinal axis x.

[0080] In the embodiment shown in FIG10 , the points of maximum curvature P, P′, P″ of the outer shell 11 of the air inlet 1 comprise different radii of curvature R, R′, so that the radial thickness of the annular cavity 12 varies in the transverse plane YY ( FIG5 ). In this example, the annular cavity 12 comprises a portion 17 of greater thickness E17 and a portion 18 of lesser thickness E18. The variation in the radial thickness of the annular cavity 12 advantageously makes it possible to promote the thrust phase A in the thick areas and the thrust reversal phase B in the thin areas. In practice, in order to maintain the aerodynamic characteristics of the air inlet 1, the radial thickness of the annular cavity 12 and the radii of curvature R, R′ of the points of maximum curvature P, P′, P″ vary gradually and slightly. In the example of FIG10 , the radial thickness of the annular cavity 12 varies gradually from the portion 17 of greater thickness E17 to the portion 18 of lesser thickness E18. Preferably, the variation of the radius of curvature R, R' from the point of maximum curvature P, P', P" to its adjacent points does not exceed 10%, preferably does not exceed 5%. More preferably, the radii of curvature R, R' of at least 25% of the points of maximum curvature P, P', P" are substantially the same, preferably, at least 50% of the points of maximum curvature P, P', P", and preferably, all the points of maximum curvature P, P', P" are substantially the same.

[0081] In summary, the aircraft propulsion assembly 8 of the present invention comprises an air inlet 1 having a robust and economical structure without moving parts and non-deformable, which facilitates both the thrust phase A and the thrust reversal phase B. More precisely, the outer casing 11 comprises an inlet lip 15 for effectively separating the upstream airflow F during the thrust A phase, and a set of points of maximum curvature P, P', P", for facilitating the separation of the reverse airflow F-INV near the inlet lip 15 during the thrust reversal B phase. To improve the performance of the aircraft propulsion assembly 8, it also comprises a variable pitch fan 3, a turbine engine 7 with a very high dilution rate, and a nacelle 2 with a reduced length and radial thickness.

[0082] A method of using an aircraft propulsion assembly 8 according to the invention during a thrust phase A and a thrust reversal phase B is described below.

[0083] 2A , during thrust phase A, such as during takeoff or cruise, the blades of variable-pitch fan 3 are oriented to allow airflow to circulate from upstream to downstream. Inlet lip 15 divides upstream airflow F into an external airflow F-EXT guided by outer wall 14 and an internal airflow F-INT guided by inner wall 13 to turbine engine 7, thereby ensuring thrust.

[0084] 2B , during thrust reversal phase B, such as during braking or landing, the pitch of the fan 3 blades is changed to reverse the direction of secondary airflow F2 in the secondary airflow path 5. Consequently, the reversed airflow F-INV circulates from downstream to upstream, passing through the fan 3 and being guided by the inner wall 13 until it separates from the inner wall at points of maximum curvature P, P', and P". A separation zone ZD is then formed, which contacts the inlet lip 15, preventing any localized depression that would be detrimental to thrust reversal B.

[0085] Switching from thrust phase A to thrust reversal phase B is advantageously simple and quick to implement, requiring only modification of the orientation of the fan blades 3. As in the prior art, no steps are required to move or deform the air inlet 1. The profile of the air inlet 1 remains unchanged between thrust phase A and thrust reversal phase B.

Claims

1. An assembly of an air intake (1) and a fan (3) of a nacelle (2) of a turbine engine (7) of an aircraft propulsion assembly (8), the aircraft propulsion assembly (8) extending along a longitudinal axis (X) oriented from upstream to downstream and comprising a turbine engine (7), the turbine engine (7) comprising a radially inner primary flow path (4) and a radially outer secondary flow path (5), the radially inner primary flow path (4) and the radially outer secondary flow path (5) being configured to guide a primary airflow (F1) and a secondary airflow (F2), respectively, from upstream to downstream during a thrust phase (A), the turbine engine (7) comprising a fan (3) rotatably mounted upstream about the longitudinal axis (X), the aircraft propulsion assembly (8) comprising a thrust reversal device configured to change the secondary airflow (F2) into a reverse direction circulating from downstream to upstream in the secondary flow path (5) during a thrust reversal phase (B). Airflow (F-INV), the fan (3) comprises a plurality of variable pitch blades so as to form a thrust reversal device of an aircraft propulsion assembly (8), the nacelle (2) extends outwards around a turbine engine (7) and comprises an air inlet (1) at its upstream end, the air inlet (1) extending outwards around the fan (3), the air inlet (1) comprising an outer casing (11) of a longitudinal axis (X), the outer casing (11) comprising an inner wall (13) turned towards the longitudinal axis (X) and an outer wall (14) opposite the inner wall (13), which are connected upstream by an air inlet lip (15), the air inlet lip (15) comprising a leading edge to promote a thrust phase (A), the air inlet (1) comprising an inner diameter at the fan (3), referred to as "fan radius (R3)", the outer casing (11) comprising a variable curvature in a plane radial to the longitudinal axis (X), the assembly having the following characteristics: the profile of the air inlet (1) is fixed and non-deformable, so that it remains unchanged between the thrust phase (A) and the thrust reversal phase (B), The outer casing (11) includes a point of maximum curvature (P, P', P") in each radial plane so as to separate the reverse flow (F-INV) from the outer casing (11) and thereby facilitate the thrust reversal phase (B), an osculating circle (C, C') including a radius of curvature (R, R') being defined at each point of maximum curvature (P, P', P"), and an average value (Rmoy) of the radius of curvature (R, R') at the periphery of the air inlet (1) satisfies: Rmoy<0.028*R3.

2. The assembly according to claim 1, wherein The average value (Rmoy) of the curvature radii (R, R') satisfies the following relationship: Rmoy<0.02*R3.

3. The assembly according to claim 1, wherein The point of maximum curvature (P, P', P") of the air inlet (1) is located in the upstream portion (16) of the outer shell (11), the upstream portion (16) including a longitudinal length (L16) defined from the leading edge of the air inlet lip (15) and satisfying the following relationship: L16<6*Rmoy.

4. The assembly according to claim 1, wherein The points of maximum curvature (P, P', P") are located on the leading edge of the air inlet lip (15).

5. The assembly according to claim 1, wherein The maximum curvature points (P, P', P") of the air inlet (1) together form a closed curve, called "separation curve (Q, Q', Q").

6. The assembly according to claim 1, wherein The values ​​of the radii of curvature (R, R') of the osculating circles (C, C') defined in at least 25% of the points of maximum curvature (P, P', P") are substantially the same.

7. An aircraft propulsion assembly (8) extending along a longitudinal axis (X) oriented from upstream to downstream and comprising a turbine engine (7) and a nacelle (2), the turbine engine (7) comprising a radially inner primary flow path (4) and a radially outer secondary flow path (5) configured to guide a primary airflow (F1) and a secondary airflow (F2), respectively, from upstream to downstream during a thrust phase (A), the aircraft propulsion assembly (8) comprising an assembly of an air inlet (1) and a fan (3) according to any one of claims 1 to 6, the turbine engine (7) comprising a fan (3) rotatably mounted upstream about the longitudinal axis (X), the aircraft propulsion assembly (8) comprising a thrust reversal device formed by the fan (3), the thrust reversal device configured to change the secondary airflow (F2) into a reverse airflow (F-INV) circulating in the secondary flow path (5) from downstream to upstream during a thrust reversal phase (B), the nacelle (2) extending outwardly around the turbine engine (7) and comprising the air inlet (1) at its upstream end.

8. A method for using an assembly of a nacelle (2) air inlet (1) and a turbine engine (7) fan (3) of an aircraft propulsion assembly (8) according to any one of claims 1 to 7, wherein: - during the thrust phase (A), the inlet lip (15) divides the upstream airflow (F) circulating from upstream to downstream into an external airflow (F-EXT) guided by the outer wall (14) and an internal airflow (F-INT) guided by the inner wall (13), and - During the thrust reversal phase (B) achieved by varying the pitch of the fan blades (3), the points of maximum curvature (P, P', P") separate from the outer casing (11), the reverse airflow (F-INV) circulates upstream in the air inlet (1) to join the external airflow (F-EXT) in order to facilitate the thrust reversal phase (B).

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