Aviation propulsion system with improved propulsion efficiency

Abstract

The invention relates to an aviation propulsion system (1), comprising a reduction mechanism (12) and an air intake duct (3), the reduction mechanism coupling a drive shaft (10) and a fan shaft (13), the air intake duct comprising an inner shroud (16) having a predetermined minimum radius (R2). The reduction mechanism (12) comprises two reduction stages (27, 32), the two reduction stages comprising planetary gears (28), each planetary gear having a first portion (38) meshing with the drive shaft (10) and a second portion (39) meshing with the fan shaft (13), each planetary gear (28) being mounted so as to be stationary relative to the inner shroud (16) of the air intake duct (3) and having a predetermined maximum radius (R1) which is greater than the minimum radius (R2) of the inner shroud (16) of the air intake duct (3), so that the air intake duct extends at least partially between the two planetary gears (28).

Description

Technical Field

[0001] The present invention relates to the field of aviation propulsion systems, and more precisely to a bypass propulsion system having a high or even very high bypass ratio and a high propulsion efficiency. Background Art

[0002] The bypass propulsion system generally includes a fan, a primary annular flow space and a secondary annular flow space from upstream to downstream in the direction of gas flow. Therefore, the mass of air sucked in by the fan is divided into a mainstream circulating in the primary flow space and a secondary flow concentric with the mainstream and circulating in the secondary flow space. The fan (or propeller) can be a ducted type and housed in a fan housing or a non-ducted variant of the unducted single fan (USF) type. The fan blades can be fixed or have a variable setting, which is adjusted according to the function of the flight phase by a pitch changing mechanism.

[0003] The main flow space passes through the main body, which includes one or more compressor stages (such as a low-pressure compressor (or supercharger) and a high-pressure compressor), a combustion chamber, one or more turbine stages (such as a high-pressure turbine and a low-pressure turbine), and a gas discharge nozzle. Typically, the high-pressure turbine rotationally drives the high-pressure compressor via a first shaft (the so-called high-pressure shaft), while the low-pressure turbine rotationally drives the low-pressure compressor and fan via a second shaft (the so-called low-pressure shaft). The low-pressure shaft is usually housed in the high-pressure shaft.

[0004] In order to improve the propulsion efficiency of the propulsion system and reduce the specific fuel consumption of the propulsion system and the noise emitted by the fan, a propulsion system with a high bypass ratio (i.e., the ratio of the flow velocity of the secondary flow to the flow velocity of the mainstream) has been provided. The term "high bypass ratio" will be understood herein to refer to a bypass ratio greater than or equal to 10 (e.g., between 10 and 80, including 10 and 80). In order to achieve such a bypass ratio, the fan is separated from the low-pressure turbine, so that the corresponding rotational speeds of the fan and the low-pressure turbine can be optimized independently. Typically, the separation is accomplished using a reduction gear (such as a planetary or planetary reduction gear) arranged between the upstream end of the low-pressure shaft and the fan. The fan is then driven by the low-pressure shaft via the reduction gear and by an additional shaft (the so-called fan shaft), which is attached between the reduction gear and the fan disc.

[0005] This separation therefore makes it possible to reduce the rotational speed and pressure ratio of the fan and to increase the power extracted by the low-pressure turbine. In particular, propulsion efficiency is the main determinant of the overall efficiency of an aviation propulsion system. Propulsion efficiency is favorably affected by minimizing the change in kinetic energy of the air during its passage through the propulsion system. In a propulsion system with a high bypass ratio, most of the flow rate that generates propulsion is constituted by the secondary flow of the propulsion system, the kinetic energy of which is mainly affected by the compression undergone by the secondary flow when passing through the fan, which acts as a low-pressure compressor. Therefore, propulsion efficiency and the pressure ratio of the fan are related: the lower the pressure ratio of the fan, the better the propulsion efficiency.

[0006] In addition to its primary influence on propulsion efficiency, the choice of the fan's pressure ratio also affects various technical characteristics of the propulsion system, including the fan diameter (and the external dimensions, mass and drag of the propulsion system and its pod), the fan's rotational rating and the reduction ratio of the reduction gear.

[0007] However, (i) the more the reduction ratio increases, the more the radial volume of the reduction mechanism increases, making it difficult to incorporate the reduction mechanism under the main flow path, and (ii) the higher the speed of the low-pressure shaft, the higher the speed of the supercharger driven by the low-pressure shaft, and the lower the average supercharger radius required to limit the circumferential speed at the blade tips of the supercharger.

[0008] These two combined limitations result in a steeper inlet channel in the main flow path upstream of the supercharger (often referred to as a swan neck due to its shape) when the radial volume of the reduction gear is large and the average supercharger radius is small. This leads to increased aerodynamic losses in this channel and poor supply to the supercharger, which adversely affects the propulsion efficiency of the propulsion system.

[0009] Currently, the preferred reduction gear mechanism is a planetary reduction gear, in which the fan is driven by a planetary pinion carrier, with a ring gear attached to the engine's stator. This is because this architecture enables higher reduction ratios than a sun-type reduction gear mechanism (where the ring gear drives the fan). However, the use of a planetary reduction gear does require the transfer of oil from the engine's fixed reference frame to the rotating reference frame of the planetary pinion carrier to supply the bearings and teeth of the reduction gear. Summary of the Invention

[0010] An object of the present invention is to provide an aviation propulsion system, such as a bypass turbine or a non-ducted propulsion system such as the USF type, the fan of which is ducted, with or without variable setting of the fan blades, which aviation propulsion system has a high bypass ratio and improved propulsion efficiency, and has a small and simplified transmission system with a high reduction ratio.

[0011] To this end, according to a first aspect of the present invention, there is provided an aviation propulsion system, the aviation propulsion system comprising:

[0012] a drive shaft, which is rotationally movable about an axis of rotation,

[0013] -fan,

[0014] a fan shaft configured to rotationally drive the fan about an axis of rotation,

[0015] - a speed reduction mechanism connecting the drive shaft and the fan shaft,

[0016] a compression section, which is driven in rotation by a drive shaft,

[0017] An air inlet passage extends between the fan and the compression section, the air inlet passage comprising an inner ring and an outer ring which together define an air inlet flow path in the compression section, the inner ring having a predetermined minimum radius.

[0018] Furthermore, the reduction mechanism includes two reduction stages, each of which includes at least two planetary gears distributed circumferentially about the rotation axis, each planetary gear including a first portion that meshes with the drive shaft and a second portion that meshes with the fan shaft. Each planetary gear is fixedly mounted relative to the inner ring of the intake housing, and each planetary gear has a predetermined maximum radius that is greater than a minimum radius of the inner ring of the intake passage, such that the air intake flow path extends at least partially between the two planetary gears.

[0019] The aircraft propulsion system may comprise a bypass turbine with a ducted fan (with or without variable setting of the fan blades), or an unducted propulsion system of the USF type.

[0020] Certain preferred but non-limiting features of the propulsion system are as follows, alone or in combination:

[0021] - the air inlet housing has an inlet adjacent to the fan and an outlet adjacent to the compression section, the section of the outer ring of the air inlet housing being substantially circular at the inlet in a radial plane to the axis of rotation;

[0022] - the passage of the air inlet flow path is discontinuous around the axis of rotation;

[0023] - the reduction mechanism is housed in a housing which forms a single unit with the inner ring of the intake channel;

[0024] The housing of the reduction mechanism comprises a series of projections at each planetary gear, each of the series of projections being configured to accommodate an associated planetary gear;

[0025] - the projections are connected in pairs by ring segments of the housing;

[0026] - the housing of the reduction mechanism is connected to the outer ring via a support arm extending from the protrusion;

[0027] The projections have the shape of a ring segment, the radius of which is greater than the radius of the housing portion extending between the projections;

[0028] - the portion of the inner ring of the intake housing extending between two adjacent planetary gears has a boss at the downstream end;

[0029] -The boss extends all the way to the compression section;

[0030] - the propulsion system further comprises a rotating multi-channel hydraulic seal located at the fan shaft upstream of the reduction gear mechanism;

[0031] - the planetary gears are mounted on a planetary pinion carrier, the propulsion system further comprising an oil reservoir and at least one conduit passing through the planetary pinion carrier, the conduit being configured to supply oil to the reduction mechanism;

[0032] - the reduction ratio of the reduction mechanism is greater than or equal to 2, preferably greater than or equal to 2.2 in the case of a ducted fan and greater than or equal to 5 in the case of a non-ducted fan, and / or

[0033] - The bypass ratio of the propulsion system is greater than or equal to 10 and less than or equal to 80.

[0034] According to a second aspect, the present invention provides an aircraft comprising an aircraft propulsion system according to the first aspect. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] Other features, objects and advantages of the present invention will become apparent from the following description, which is purely illustrative and non-limiting and must be read with reference to the accompanying drawings, in which:

[0036] Figure 1 An example of an aircraft propulsion system comprising a ducted fan with variable settings according to an embodiment of the invention is schematically shown.

[0037] Figure 2 An example of an aircraft propulsion system comprising a USF type unducted fan according to an embodiment of the invention is schematically shown.

[0038] Figure 3a is a partial schematic perspective view of a first example of a speed reduction mechanism and an air inlet of an aircraft propulsion system according to the present invention.

[0039] Figure 3b yes Figure 3a A schematic cross-sectional view of the propulsion system, in addition, Figure 3bAlso shown is the inlet cone of the propulsion system.

[0040] Figure 4a is a partial schematic perspective view of a second example of a speed reduction mechanism and an air inlet of an aircraft propulsion system according to the present invention.

[0041] Figure 4b yes Figure 4a A schematic cross-sectional view of the propulsion system, in addition, Figure 4b Also shown is the inlet cone of the propulsion system.

[0042] Figure 5 It is a partial schematic cross-sectional view showing a two-stage sun gear reduction mechanism (upper left), a single-stage planetary reduction mechanism (upper right), a two-stage planetary reduction mechanism (lower left) and the reduction mechanism of the present invention (lower right) and their corresponding air intake flow paths in four quadrants. These four reduction mechanisms all have the same reduction ratio.

[0043] Similar elements have the same reference numerals throughout the drawings. DETAILED DESCRIPTION

[0044] The propulsion system 1 generally comprises a fan 2 and a main body. The main body comprises an air intake duct 3 extending directly downstream of the fan 2 along the direction of gas flow within the propulsion system 1, a low-pressure compressor 4 (or supercharger 4), a high-pressure compressor 5, a combustion chamber 6, a high-pressure turbine 7, a low-pressure turbine 9, and a gas discharge nozzle. The high-pressure turbine 7 rotationally drives the high-pressure compressor 5 via a high-pressure shaft 8, while the low-pressure turbine 9 rotationally drives the low-pressure compressor 4 and the fan 2 via a low-pressure shaft 10.

[0045] The fan 2 comprises a fan disk 2a (particularly at Figure 3b and Figure 4b ), fan blades 11 are provided on the periphery of the fan disk. When the fan blades rotate, the fan blades drive the air flow into the primary flow space and the secondary flow space of the propulsion system 1.

[0046] The supercharger 4 comprises at least one compression stage comprising a wheel 14 of movable blades (rotor), driven by a low-pressure shaft 10, which rotates in front of a series of fixed vanes (stators or straighteners) distributed circumferentially around the X axis. Where applicable, the supercharger 4 may comprise at least two compression stages.

[0047] The air intake duct 3 extends directly downstream of the fan 2. The air intake duct has an inlet 18 and an outlet 20, the inlet being adjacent to the root 17 of the fan blades 11, perpendicular to the front fairing 19 of the primary and secondary flow spaces, and the outlet being adjacent to the supercharger 4. The air intake duct 3 has the general shape of a swan neck, so that the inlet 18 is radially further away from the axis of rotation X than the outlet 20. The air intake duct 3 comprises, in a manner known per se, a row of fixed blades, which are distributed circumferentially around the axis X between an inner ring 16 and an outer ring 17, which together define the air intake flow path in the supercharger 4. The inner ring 16 and the outer ring 17 of the air intake duct 3 extend circumferentially around the axis of rotation X and are fixed in the engine reference frame. Typically, the inner ring 16 and the outer ring 17 can be connected to a casing of the engine (such as the air intake casing 3).

[0048] The invention is applicable to any type of bypass aircraft propulsion system 1 , whether the fan is ducted or unducted, with fixed blades or a variable arrangement.

[0049] In this application, upstream and downstream are defined relative to the normal direction of the gas flow in the fan and through the propulsion system. In addition, the axial direction corresponds to the direction of the rotation axis X, and the radial direction is a direction perpendicular to the rotation axis X and passing through the rotation axis. In addition, the circumferential (or transverse) direction corresponds to a direction perpendicular to the rotation axis X and not passing through the rotation axis. Unless otherwise specified, the internal (or inward) and external (or outward) are used with reference to the radial direction, respectively, so that a part or an internal face of an element is closer to the rotation axis X than an external part or face of the same element.

[0050] The propulsion system 1 has a high bypass ratio. The term "high bypass ratio" is understood to mean a bypass ratio greater than or equal to 10 (for example, between 10 and 80). To this end, the fan 2 is separated from the low-pressure turbine 9 to independently optimize the respective rotational speeds of the fan and the low-pressure turbine using a reduction mechanism 12, which is arranged between the upstream end of the low-pressure shaft 10 (relative to the gas flow direction in the propulsion system 1) and the fan 2. The fan 2 is then driven by the low-pressure shaft 10 via the reduction mechanism 12 and a fan shaft 13, which is attached between the reduction mechanism 12 and the disk 2a of the fan 2. The fan shaft 13 is rotatably movable around a rotation axis X that is coaxial with the rotation axis X of the low-pressure shaft 10.

[0051] To calculate the bypass ratio, the secondary flow velocity and the primary flow velocity are measured when the propulsion system 1 is stabilized at takeoff ratings in standard atmosphere (as defined in International Civil Aviation Organization (ICAO) Document 7488 / 3, 3rd edition) and at sea level. In the remainder of this document, all parameters will be measured under these conditions.

[0052] In order to improve the propulsion efficiency of the propulsion system 1, the reduction mechanism is a two-stage static type. More precisely, the reduction mechanism 12 includes two reduction stages 27, 32, each stage 27, 32 including at least two planetary gears 28, the at least two planetary gears being distributed circumferentially around the axis of rotation X and including a first portion 38 meshing with the low-pressure shaft 10 and a second portion 39 meshing with the fan shaft 13. The first portions 38 of the planetary gears 28 extend in the same plane (the so-called first plane) and form the first stage 27 of the reduction mechanism 12, while the second portions 39 of the planetary gears 28 extend in the same plane (the so-called second plane), which is parallel to the plane of the first portions 38 and forms the second stage 32 of the reduction mechanism 12. The second plane is axially offset relative to the first plane. For example, the reduction mechanism 12 may include three or more planetary gears 28.

[0053] The planetary gears 28 are fixedly mounted relative to the inner ring 16 of the intake passage 3, and each planetary gear has a predetermined maximum radius R1 corresponding to the volume radius of the planetary gear 28. Furthermore, the maximum radius R1 of the planetary gears 28 is greater than the minimum radius R2 of the inner ring 16 of the intake passage 3, so that the air intake flow path extends at least partially between two adjacent planetary gears 28. The term "maximum radius R1 of the planetary gears 28" should be understood to refer to the maximum distance between the rotation axis X and the outer radial surface of the planetary gear 28 in a plane radial to the rotation axis X. Here, the maximum radius R1 is measured at the first portion 38 of the planetary gear 28, which has a larger diameter than the second portion 39. Furthermore, the term "minimum radius R2 of the inner ring 16 of the intake passage 3" should be understood to refer to the minimum distance between the rotation axis X and the inner ring 16 at the inlet 18 of the inner ring 16, in a plane radial to the rotation axis X. More precisely, the minimum radius R2 of the inner ring 16 is measured at the most upstream portion of the front fairing 19 .

[0054] Therefore, the reduction mechanism 12 does not include any radially bulky components that can move about the axis of rotation X, since the fan shaft 13 meshes directly with the second portion 39 of the planetary gears 28, and the planetary gears 28 are static in the engine reference frame (attached to the inner ring 16 of the intake housing 3). This makes it possible for the reduction mechanism 12 to not include any ring gear or planetary pinion carrier 21 that can move about the axis of rotation X. The centrifugal forces experienced by the reduction mechanism 12 are therefore much lower than those in reduction mechanisms of the sun gear or planetary gear type.

[0055] Furthermore, the air intake path partially passes between the planetary gears 28, whose inner ring 16 has a lower minimum radius R2 at the inlet 18 than in the prior art, resulting in a gentler slope of the intake passage 3 upstream of the supercharger 4, which improves the supply to the supercharger 4. In particular, the minimum radius R2 is at most 90% of the maximum radius R1 of the planetary gears 28, preferably at most 80% of the maximum radius R1.

[0056] Furthermore, the use of a two-stage static reduction mechanism 12 enables high, or even very high, reduction ratios to be achieved with a smaller radial volume compared to prior art engines. This reduction in the radial volume of the reduction mechanism 12 has the effect of making the slope of the intake passage 3, along the main flow path upstream of the supercharger 4, more gradual for the same reduction ratio. Consequently, the choice of a two-stage static reduction mechanism 12 doubles the supply to the supercharger 4 and, consequently, improves the propulsion efficiency of the propulsion system 1.

[0057] Where applicable, there is also the option of limiting the increase in the axial volume of the compression / reduction coupling by reducing the distance between the reduction unit and the compressor, thereby maintaining a steep inlet duct slope.

[0058] The reduction mechanism 12 is housed in a housing 23 that forms a single unit with the inner ring 16 of the intake passage 3. As long as the minimum radius R2 of the inner ring 16 is smaller than the maximum radius R1 of the planetary gears 28 so that the air flow path can pass between adjacent planetary gears 28, the housing 23 of the reduction mechanism 12 is not cylindrical.

[0059] More precisely, the housing 23 of the reduction mechanism 12 comprises, at the planetary gears 28, a series of projections 25, each configured to accommodate an associated planetary gear, the projections 25 being connected in pairs by ring segments 26 of the housing. The term "ring segment" should be understood to mean that these portions 26 of the housing 23 have a curved section corresponding to a semicircle in a plane radial to the axis of rotation X. The ring segments 26 all have the same radius. Where applicable, the radius of the ring segments 26 is equal to the minimum radius R2 of the inner ring 16.

[0060] exist Figure 4a and Figure 4b In the first embodiment shown, the inlet 18 of the intake passage 3 is segmented, i.e., it is not circumferentially continuous. The inlet 18 comprises a plurality of circumferential segments separated in pairs by projections 25 of the housing 23 of the reduction mechanism 12, each of which forms a channel for the flow path.

[0061] To this end, in an embodiment, the projections 25 have the shape of ring segments, the radius R3 of the projections being greater than the radius of the portion 26 of the housing 23 extending between the projections 25. In this embodiment, the radius R3 of the projections 25 can be substantially equal to the radius of the outer ring 17 of the housing 23 of the air inlet 3. The projections 25 are then connected to the adjacent ring segment portions by two walls 33, which extend substantially in the direction of flow at the inlet of the channel 3. The walls 33 can be planar or flat. Figure 4a The curved shape shown is modified to continuously form sections that converge and then diverge in the direction of gas flow, in the manner of a de Laval nozzle. The inlet 18 of the intake duct is then formed by passages 34, each of which is bounded by a portion 26 of the ring segment housing 23 and walls 33 extending on either side of the portion 26. Typically, in the case of a reduction gear 12 including three planetary gears 28, the housing 23 includes three portions 26 in the shape of ring segments (with three pairs of walls 33 extending on either side of each planetary gear 28), thereby forming three circumferential passages 34 for the airflow path into the intake duct 3.

[0062] In the variant, Figure 3a and Figure 3b In the second form of embodiment shown, the inlet 18 of the air inlet channel 3 is continuous over the entire circumference of this inlet around the axis of rotation X.

[0063] In this form of embodiment, the reduction mechanism 12 comprises a generally curved projection 25, the shape and dimensions of which are adapted to the shape and dimensions of the planetary gears 28. Furthermore, the housing 23 is connected to the outer ring 17 via support arms 36 extending from the projection 25. Typically, the propulsion system 1 may comprise as many support arms 36 as there are planetary gears 28, each extending from a vertex of the projection 25 housing a planetary gear 28. Consequently, the inlet of the air flow path into the intake duct 3 is formed by a single, continuous circumferential passage 35 about the axis of rotation X. Consequently, the outer ring 17 of the intake duct 3 housing 23 is generally circular at the projection 25, whereas the inner ring 16 is not cylindrical.

[0064] The presence of the projection 25 facilitates a change in the tangential velocity of the air flow at the inlet of the supercharger 4. In a form of embodiment, the inner ring 16 then comprises a boss 40 downstream of the inlet 18, in the extension of at least one portion 26 of the housing 23 of the reduction mechanism 12 (for example, in the extension of the ring segment portion 26 of the housing 23), which extends between two adjacent planetary gears 28. The boss 40 may extend over the entire length of the intake channel 3 or, in a variant, only over a portion of the intake channel 3.

[0065] In a form of embodiment, the inner ring 16 comprises as many bosses 40 as there are planetary gears 28 , each boss 40 extending in the extension of the portion 26 of the housing 23 situated between two adjacent planetary gears.

[0066] The high reduction ratio makes it possible to reduce the rotation speed and compression ratio of the fan 2 and optimize the size of the low-pressure turbine 9. Therefore, the propulsion efficiency of the propulsion system 1 is improved.

[0067] As a comparison, Figure 5 The radial volumes obtained for four reduction mechanisms with different technologies but the same reduction ratio are shown. Figure 5 The upper left quadrant of FIG is clockwise composed of the two-stage sun reduction mechanism, the planetary reduction mechanism, the two-stage static reduction mechanism 12, and the two-stage planetary reduction mechanism. This comparison clearly shows that, for the same reduction ratio, the reduction mechanism with the smallest local radial volume is the static two-stage reduction mechanism 12, as space is freed for the passage 34 of the airflow path between the fixed planetary gears 28.

[0068] The reduction ratio of the secondary solar reduction mechanism is at least equal to 2.

[0069] The propulsion system 1 includes a ducted fan 2 ( Figure 1 ) and variable setting fan blades 11 (where applicable), the reduction ratio is greater than or equal to 2.2 (for example, between 2.2 and 6).

[0070] The propulsion system 1 comprises, for example, a USF type unducted fan 2 ( Figure 2 ), the reduction ratio is greater than or equal to 5 (for example, between 5 and 9). The term "USF type fan" is understood here to mean an aircraft turbomachine comprising a non-ducted fan 2 (propeller), a gas generator intended to drive the non-ducted fan 2 in rotation about a longitudinal axis X, and at least one straightener comprising a plurality of variable-setting stator blades extending from a fixed casing along a radial axis perpendicular to the longitudinal axis X.

[0071] In an embodiment, the planetary gears 28 are mounted on the planetary pinion carrier 21 so as to be rotatably movable about respective rotation axes 37, for example, via journal bearings. On the other hand, the planetary pinion carrier 21 is fixed in the engine reference frame. For example, the planetary pinion carrier 21 may be attached to the inner ring 16 of the intake housing 23.

[0072] Each planet gear 28 is rotationally symmetrical relative to the axis of rotation 37 of the planet gear. The diameter of the first portion 38 is different from the diameter of the second portion 39. Thus, each portion 38, 39 of the planet gear 28 forms a stage 27, 32 of the reduction mechanism 12. More precisely, the first portion 38 of each planet gear 28 is cylindrical in rotation relative to the axis of rotation 37 of the first portion and has an outer radial surface configured to interact with the outer radial surface of the upstream end 35 of the low-pressure shaft 10. To this end, the inner radial surface of the first portion 38 of the planet gear 28 comprises meshing means (generally teeth 29) configured to mesh with the splines of the upstream end 35 of the low-pressure shaft 10.

[0073] The second portion 39 of each planet gear 28 is cylindrical in rotation relative to its axis of rotation 37 and has an inner radial surface configured to interact with the outer radial surface of the fan shaft 13. To this end, the inner surface of this second portion 39 also comprises meshing means (generally teeth 29) configured to mesh with splines formed on the upstream end 35 of the low-pressure shaft 10.

[0074] The meshing means may comprise straight teeth, helical teeth or herringbone teeth.

[0075] The first part 38 and the second part 39 of each planet gear 28 are a single unit. For example, the first part 38 and the second part 39 of the same planet gear 28 can be formed entirely from a single part (integrally). In a variant, the first part 38 and the second part 39 of the same planet gear 28 can be assembled.

[0076] Furthermore, the planetary gears 28 of the same reduction mechanism 12 are identical in shape and size.

[0077] Thus, in operation, the rotation of the low-pressure shaft 10 about the axis of rotation X drives the first part 38 of the planetary gear 28 to rotate about the axis of rotation 37 of the first part (the axis of rotation 37 is fixed relative to the inner ring 16 of the intake 3 housing 23), the first part of the planetary gear forming a single unit with the second part 39 of the planetary gear, and thus drives the fan shaft 13 to rotate about the axis of rotation X at different speeds.

[0078] The diameter of the second portion 39 of the planetary gear 28 is strictly smaller than the diameter of the first portion 38. Specifically, it is the difference in diameter between the first portion 38 and the second portion 39 of the planetary gear 29 that enables a higher reduction ratio to be achieved than the single-stage reduction mechanism 12 for a comparable radial volume. As a result, the diameter of the first portion 38 and the diameter of the second portion 39 of the planetary gear 28 can be sized so as to achieve a reduction ratio greater than or equal to 2 with a small radial volume, thereby making the slope of the intake passage 3 more gentle.

[0079] In this embodiment, since the planetary gears 28 are fixed in the engine reference frame, the hydraulic seals 15 can be arranged upstream of the reduction gear 12 and fluidically connected to the oil reservoir of the lubrication unit via a duct passing through the planetary pinion carrier 21. In other words, it is no longer necessary to convey the oil from the fixed reference frame of the engine to the rotating reference frame of the reduction gear 12 in order to supply the bearings and teeth of the reduction gear: it is sufficient to convey the oil directly into the ducts passing through the fixed planetary pinion carrier 21 and then, from these ducts, supply the bearings and teeth 29 of the reduction gear 12.

[0080] The design of the hydraulic seal 15 and the supply of the hydraulic seal are thereby simplified.

[0081] Since the hydraulic seal 15 is arranged upstream of the reduction mechanism 12 , it is more easily accessible, which simplifies maintenance operations.

[0082] In this first embodiment, the hydraulic seal 15 can be mounted on the fan shaft 13 , for example. For example, the hydraulic seal 15 includes a rotating portion mounted on the fan shaft 13 and a fixed portion mounted on the planet pinion carrier 21 .

[0083] Optionally, the propulsion system 1 further comprises a pitch changing mechanism 43 configured to modify the setting angle of the fan blades 11 as a function of the flight phase of the propulsion system. This pitch changing mechanism 43 then requires a lubricating device (such as a spray nozzle) supplied with oil by the hydraulic seal 15. Therefore, the propulsion system 1 further comprises an oil supply flow path 2 extending between the rotating portion of the hydraulic seal 15 and the pitch changing mechanism 43. These flow paths are rotationally fixed to the rotating portion of the hydraulic seal 15.

[0084] Advantageously, since the hydraulic seal 15 is located upstream of the reduction mechanism 12 , the distance between the hydraulic seal 15 and the lubrication device is shorter than in the prior art, and thus simpler.

[0085] Furthermore, only the portion of the hydraulic seal 15 supplying the lubrication device of the pitch changing mechanism 43 comprises a rotating part, the hydraulic seal 15 being supplied via a fixed line passing through the planet pinion carrier 21 .

[0086] In addition, where the propulsion system 1 includes a ducted fan 2, the diameter D of the fan can be between 105 inches (266.7 cm) and 135 inches (342.9 cm). Where the propulsion system 1 includes a non-ducted fan 2, the diameter D of the fan 2 can be between 150 inches (381 cm) and 180 inches (457.2 cm), for example, on the order of 167 inches (424.18 cm). The term "fan 2 diameter D" should be understood herein to mean twice the distance between the axis of rotation X and the apex 30 of the fan blade 11 measured in a plane radial to the axis of rotation X at the intersection between the leading edge 31 and the apex 30 of the blade 11. The term "leading edge 31" should be understood herein to mean the edge of the fan 11 that is configured to extend facing the airflow entering the fan 2. The leading edge corresponds to the front portion of the aerodynamic profile that faces the airflow and that separates the airflow into a pressure surface flow and a suction surface flow. Meanwhile, the trailing edge corresponds to the tail of the aerodynamic profile where the pressure surface flow and the suction surface flow meet.

[0087] Furthermore, for these fan 2 diameters D and rotational ratings, the compression ratio of the ducted fan 2 can be between 1.04 and 1.29, while the pressure ratio of the non-ducted fan 2 can be between 1.01 and 1.025. The compression ratio of the fan 2 is measured here under the same conditions as the bypass ratio, i.e., when the propulsion system 1 is held steady at takeoff ratings in standard atmosphere (as defined by International Civil Aviation Organization (ICAO) Document 7488 / 3, 3rd edition) and at sea level.

[0088] The bypass ratio of the propulsion system 1 may be between 10 and 31 in the case of a ducted fan, and between 40 and 80 in the case of a non-ducted fan 2 .

[0089] The circumferential velocity at the tip of the fan blade 11 at takeoff rating as defined above (i.e., measured at the apex 30 of the fan blade) is between 260 m / s and 330 m / s when the fan 2 is ducted and less than 225 m / s when the fan 2 is non-ducted.

Claims

1. An aviation propulsion system (1), comprising: - a drive shaft (10) which is movable in rotation about an axis of rotation (X), - fan (2), a fan shaft (13) configured to drive the fan in rotation about the axis of rotation (X), - a speed reduction mechanism (12), the speed reduction mechanism (12) coupling the drive shaft (10) and the fan shaft (13), - a compression section (4) which is driven in rotation by the drive shaft (10), - an air inlet channel (3) extending between the fan (2) and the compression section (4), the air inlet channel (3) comprising an inner ring (16) and an outer ring (17), the inner ring and the outer ring together defining an air inlet flow path in the compression section (4), the inner ring (16) having a predetermined minimum radius (R2), The aviation propulsion system (1) is characterized in that: - the reduction mechanism (12) comprises two reduction stages (27, 32), the two reduction stages comprising at least two planetary gears (28) distributed circumferentially around the axis of rotation (X), each planetary gear (28) comprising a first portion (38) meshing with the drive shaft (10) and a second portion (39) meshing with the fan shaft (13), - each planetary gear (28) is fixedly mounted relative to the inner ring (16) of the intake channel (3), and - Each planet gear (28) has a predetermined maximum radius (R1), which is greater than the minimum radius (R2) of the inner ring (16) of the intake channel (3), so that the air intake flow path extends at least partially between two of the planet gears (28).

2. The aviation propulsion system (1) according to claim 1, wherein: The air inlet channel (3) has an inlet (18) adjacent to the fan (2) and an outlet (20) adjacent to the compression section (4), and the section of the outer ring (17) of the air inlet channel (3) is generally circular at the inlet (18) in a radial plane of the rotation axis (X).

3. The aviation propulsion system (1) according to claim 1 or 2, wherein: The passageway (34) of the air intake flow path is discontinuous around the axis of rotation (X).

4. The aviation propulsion system (1) according to claim 3, wherein: The speed reduction mechanism (12) is housed in a housing (23), and the housing (23) forms a single unit with the inner ring (16) of the intake passage (3).

5. The aviation propulsion system (1) according to claim 4, wherein: The housing (23) of the reduction mechanism (12) includes a series of protrusions (25) at each planetary gear (28), each of the series of protrusions being configured to accommodate an associated planetary gear (28).

6. The aviation propulsion system (1) according to claim 5, wherein: The projections (25) are connected in pairs by ring segments (26) of the housing (23).

7. The aviation propulsion system (1) according to claim 5 or 6, wherein: The housing (23) of the speed reduction mechanism (12) is connected to the outer ring (17) via a support arm (36) extending from the protrusion (25).

8. The aviation propulsion system (1) according to claim 6, wherein: The protrusions (25) have a ring segment shape, the radius (R3) of the ring segment shape being larger than the radius of the ring segment portion (26) extending between the protrusions (25).

9. The aviation propulsion system (1) according to claim 2, wherein: The portion of the inner ring (16) of the intake passage (3) extending between two adjacent planetary gears (28) has a boss (40) downstream of the inlet (18).

10. The aviation propulsion system (1) according to claim 9, wherein: The boss (40) extends all the way to the compression section (4).

11. The aviation propulsion system (1) according to claim 1 or 2, further comprising a rotating multi-channel hydraulic seal (15), which is located at the fan shaft (13) upstream of the reduction gear (12).

12. The aviation propulsion system (1) according to claim 11, wherein: The planetary gears (28) are mounted on a planetary pinion carrier (21). The aircraft propulsion system (1) further comprises an oil reservoir and at least one pipe passing through the planetary pinion carrier (21), wherein the pipe is configured to supply oil to the reduction mechanism (12).

13. The aviation propulsion system (1) according to claim 1 or 2, wherein: The reduction ratio of the reduction mechanism (12) is greater than or equal to 2.

14. The aviation propulsion system (1) according to claim 1 or 2, wherein a bypass ratio of the aviation propulsion system (1) is greater than or equal to 10 and less than or equal to 80.

15. An aircraft comprising an aeronautical propulsion system (1) according to any one of claims 1 to 14.

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