Aeronautical propulsion system with improved propulsion efficiency

Abstract

An aeronautical propulsion system (1) comprises a drive turbine (8), a fan rotor (9) and a reduction mechanism (19) coupling the shaft (11) of the drive turbine (8) and the shaft (20) of the fan rotor (9). A rotational speed of the drive shaft (11) further satisfies formula (I) where: Ni is the rotational speed of the drive shaft (11); Te is the inlet temperature of the drive turbine (8); and Se is an inlet cross section of the drive turbine (8); and formula (II).

Description

[0001] DESCRIPTION

[0002] TITLE: Aeronautical propulsion system with improved propulsive efficiency

[0003] TECHNICAL FIELD

[0004] The present application relates generally to the field of propulsion systems, and more particularly to propulsion systems comprising a ducted or unducted fan and having high propulsive efficiency and fan efficiency.

[0005] STATE OF THE ART

[0006] A propulsion system generally comprises, from upstream to downstream in the direction of gas flow, a fan section, a compressor section which may include a low-pressure compressor and a high-pressure compressor, a combustion chamber and a turbine section which may include in particular a high-pressure turbine and a low-pressure turbine. The high-pressure compressor is rotated by the high-pressure turbine via a high-pressure shaft. The fan and, where appropriate, the low-pressure compressor are rotated by the low-pressure turbine via a low-pressure shaft.

[0007] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into consideration the impact factors in all phases of design and development to obtain less energy-intensive, more environmentally friendly aeronautical components and products whose integration and use in civil aviation have moderate environmental consequences with the aim of improving the energy efficiency of aircraft.

[0008] Thus, in order to improve the propulsive efficiency of the propulsion system and to reduce its specific consumption as well as the noise emitted by the fan section, propulsion systems have been proposed having a high bypass ratio (BPR, corresponding to the ratio between the flow rate of the secondary air flow and the flow rate of the primary air flow). To achieve such bypass ratios, the fan section can be decoupled from the low-pressure turbine, thus making it possible to independently optimize their respective rotation speeds. Generally, the decoupling is achieved using a reduction mechanism placed between the upstream end of the low-pressure shaft and a rotor of the fan section. The rotor of the fan section is then driven by the low-pressure shaft via the reduction mechanism at a rotation speed lower than that of the low-pressure shaft.

[0009] In order to further improve the overall efficiency of the turbomachine, the current trend is to increase the overall compression ratio of the propulsion system, which corresponds to the ratio between the pressure at the outlet of the high-pressure compressor and the pressure at the inlet of the fan. This therefore requires increasing the compression ratio of the high-pressure compressor and / or the low-pressure compressor, especially since at the same time we are trying to reduce the compression ratio of the fan to improve its efficiency. One of the possible consequences is that the low-pressure turbine, which drives the fan, is more mechanically loaded, particularly at the bottom of the blade.However, any modification to the low pressure turbine is likely to have an impact on the flow path within the turbine section which may have consequences on its efficiency, the possibilities of integrating the propulsion system into an aircraft or even on the integration of the bearings.

[0010] EXPOSED

[0011] An aim of the present application is to optimize the propulsion system in order to increase its efficiency without mechanically and / or thermally overloading the fan drive turbine, while taking into account the integration constraints of the propulsion system.

[0012] For this purpose, according to a first aspect, an aeronautical propulsion system is proposed comprising:

[0013] - a drive turbine connected to a drive shaft which can rotate around an axis of rotation;

[0014] - a fan section comprising a fan rotor connected to a fan shaft;

[0015] - a reduction mechanism coupling the drive shaft and the fan shaft in order to drive the fan shaft at a rotational speed lower than the rotational speed of the drive shaft; the propulsion system being configured so that a rotational speed of the drive shaft complies with the following formula:

[0016] -T |CZ*T e + ? .

[0017] 1 > / — S —ex 10 3 where: Ni is the rotational speed of the drive shaft when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in revolutions per minute (rpm);

[0018] Te is the temperature at the inlet of the drive turbine when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in degrees Celsius (°C) and is greater than or equal to 700 °C;

[0019] Se is an inlet section of the drive turbine (8), in square meters (m 2 ) ; and a = - 0.056 83.356 f \—mxm) 2 . in 7

[0020] Some preferred but non-limiting features of the propulsion system according to the first aspect are the following, taken individually or in combination:

[0021] - the propulsion system is further configured so that the drive speed of the drive turbine also complies with the following formula:

[0022] - the reduction mechanism has a reduction ratio greater than or equal to 2.5, preferably greater than or equal to 3.0 and less than or equal to 11.0;

[0023] - the drive turbine comprises at least 3 stages and at most 5 stages;

[0024] - the drive turbine has a hub-to-head ratio at the input greater than 0.75 and less than 0.90;

[0025] - the drive turbine has a hub-to-head output ratio greater than 0.55 and less than 0.75;

[0026] - the compressor comprises at least two stages and at most four stages;

[0027] - the propulsion system further comprises a second turbine and a second compressor connected via a second shaft, the second shaft rotating faster than the drive shaft, the second turbine being two-stage;

[0028] - the second compressor comprises at least eight stages and at most eleven stages;

[0029] - a thrust density per blade of the fan rotor of the propulsion system is greater than or equal to 5.0 x 10 4 and less than or equal to 17.0 x 10 4 N / m 2 where the thrust density per blade is defined by the following formula: 100 and where: FN is the thrust generated by the fan rotor and is measured when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level and is expressed in Newton (N); n is the number of blades in the fan rotor; and D is the diameter of the fan rotor, measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of the fan rotor blades, and is expressed in meters (m);

[0030] - a power density per fan rotor blade is greater than or equal to 3.65 x 10 6 and less than or equal to 22.0 x 10 6 W / m 2, where the power density per fan rotor blade is defined by the following formula: . 100 and where: fan power is the power of the fan rotor and is measured when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level and is expressed in Watts (W); n is the number of blades in the fan rotor; and D is the diameter of the fan rotor, measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of the fan rotor blades, and is expressed in meters (m);

[0031] - the fan section further has a fan compression ratio, corresponding to a pressure ratio between an outlet of the fan rotor and an inlet of the fan rotor less than or equal to 1.45;

[0032] - the diameter of the fan rotor is between 80 inches and 185 inches inclusive, preferably between 85 inches and 120 inches inclusive, for example of the order of 90 inches;

[0033] - the fan section is faired and a dilution ratio of the propulsion system is greater than or equal to 10, for example between 10 and 35 inclusive, preferably between 10 and 18 inclusive;

[0034] - the fan section is ducted and a peripheral speed at the tip of the fan rotor blades, when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level, is between 260 m / s and 400 m / s;

[0035] - the fan section is shrouded and a thrust density per blade of the fan rotor of the propulsion system is greater than or equal to 14.0 x 10 4 and less than or equal to 17.0 x 10 4 N / m 2where the thrust density per blade is defined by the following formula: 100 and where: FN is the thrust generated by the fan rotor and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in Newton; n is the number of blades in the fan rotor; and

[0036] D is the diameter of the fan rotor, measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of the fan rotor blades, and is expressed in meters (m);

[0037] - the fan section is unducted and a propulsion system bypass ratio is greater than or equal to 40, for example between 40 and 80 inclusive; and / or

[0038] - the fan section is unducted and a peripheral speed at the tip of the fan rotor blades, when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level, is between 210 m / s and 260 m / s; and / or

[0039] - the fan section is unducted and a thrust density per fan rotor blade of the propulsion system is greater than or equal to 5.0 x 10 4 and less than or equal to 10.0 x 10 4 N / m 2 where the thrust density per blade is defined by the following formula: 100 and where: FN is the thrust generated by the fan rotor and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in Newton; n is the number of blades in the fan rotor; and

[0040] D is the diameter of the fan rotor, measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of the fan rotor blades, and is expressed in meters (m).

[0041] According to a second aspect, the present application proposes an aircraft comprising at least one propulsion system according to the first aspect fixed to the aircraft by means of a mast.

[0042] According to a third aspect, the invention provides a method for dimensioning or manufacturing a propulsion system comprising a reduction mechanism coupling a drive turbine and a fan rotor to drive the fan rotor at a speed lower than a speed of the drive turbine, the drive turbine configured so that a rotational speed of the drive shaft complies with the following formula:

[0043] N 1 > / — S —ex 10 3where: Ni is the rotational speed of the drive shaft when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in revolutions per minute (rpm);

[0044] Te is the temperature at the inlet of the drive turbine when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in degrees Celsius (°C) and is greater than or equal to 700 °C;

[0045] Se is an inlet section of the drive turbine (8), in square meters (m 2 ) ; and a = - 0.056 83.356 f \—min xm) J 2 .

[0046] Some preferred but non-limiting aspects of the sizing or manufacturing method according to the third aspect are the following, taken individually or in combination:

[0047] - the fan section is dimensioned so that a thrust density per blade of the fan rotor of the propulsion system is greater than or equal to 5.0 x 10 4 and less than or equal to 17.0 x 10 4 N / m 2 where the thrust density per blade is defined by the following formula: 100 and where: FN is the thrust generated by the fan rotor and is measured when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level and is expressed in Newton (N); n is the number of blades in the fan rotor; and D is the diameter of the fan rotor, measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of the fan rotor blades, and is expressed in meters (m); and / or

[0048] - the fan is further dimensioned so that a power density per fan rotor blade is greater than or equal to 3.65 x 10 6 and less than or equal to 22.0 x 10 6 W / m 2 , where the power density per fan rotor blade is defined by the following formula: . 100 and where: fan power is the power of the fan rotor and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in Watts (W); n is the number of blades in the fan rotor; and

[0049] D is the diameter of the fan rotor, measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of the fan rotor blades, and is expressed in meters (m).

[0050] According to a fourth aspect, there is provided a method of manufacturing a propulsion system comprising the following steps:

[0051] - dimensioning the propulsion system according to the method according to the third aspect; and

[0052] - manufacture the propulsion system.

[0053] DESCRIPTION OF FIGURES

[0054] Other characteristics, aims and advantages will emerge from the following description, which is purely illustrative and not limiting, and which must be read in conjunction with the attached drawings in which:

[0055] Figure 1 is a schematic, partial, sectional view of an exemplary propulsion system according to one embodiment, in which the fan section is ducted;

[0056] Figure 2 is a schematic, partial, sectional view of an example propulsion system according to a first embodiment, in which the fan section is unducted; Figure 3 is a schematic sectional view of an example planetary reduction mechanism;

[0057] Figure 4 is a schematic sectional view of an exemplary epicyclic reduction mechanism;

[0058] Figure 5 is an example of an aircraft that may include at least one propulsion system according to one embodiment;

[0059] Figure 6 is a flowchart illustrating exemplary steps of a sizing or manufacturing method according to one embodiment.

[0060] Throughout the figures, similar elements have identical references.

[0061] DETAILED DESCRIPTION

[0062] A propulsion system 1 has a main direction extending along a longitudinal axis X and comprises, from upstream to downstream in the direction of flow of the gases in the propulsion system 1 when it is in operation, a fan section 2 and a primary body 3, often called a “gas generator”, comprising a compressor section 4, 5, a combustion chamber 6 and a turbine section 7, 8. The propulsion system 1 is here an aeronautical propulsion system 1 configured to be fixed to an aircraft 100 via a pylon (or mast).

[0063] The compressor section 4, 5 comprises a succession of stages each comprising a wheel of moving blades (rotor) 4a, 5a rotating in front of a wheel of fixed blades (stator) 4b, 5b. The turbine section 7, 8 also comprises a succession of stages each comprising a wheel of fixed blades (stator) 7a, 8a behind which a wheel of moving blades (rotor) 7b, 8b rotates.

[0064] In the present application, the axial direction corresponds to the direction of the longitudinal axis X, in correspondence with the rotation of the shafts of the gas generator, and a radial direction is a direction perpendicular to this axis X and passing through it. Furthermore, the circumferential (or lateral, or even tangential) direction corresponds to a direction perpendicular to the longitudinal axis X and not passing through it. Unless otherwise specified, internal (respectively, interior) and external (respectively, exterior), respectively, are used in reference to a radial direction so that the internal part or face of an element is closer to the axis X than the external part or face of the same element.

[0065] In operation, an air flow F entering the propulsion system 1 is divided between a primary air flow F1 and a secondary air flow F2, which circulate from upstream to downstream in the propulsion system 1.

[0066] The secondary airflow F2 (also called the "bypass airflow") flows around the primary body 3. The secondary airflow F2 cools the periphery of the primary body 3 and is used to generate most of the thrust provided by the propulsion system 1.

[0067] The primary air flow F1 flows in a primary vein inside the primary body 3, passing successively through the compressor section 4, 5, the combustion chamber 6 where it is mixed with fuel to serve as an oxidant, and the turbine section 7, 8. The passage of the primary air flow F1 through the turbine section 7, 8 receiving energy from the combustion chamber 6 causes rotation of the rotor of the turbine section 7, 8, which in turn drives rotation of the rotor of the compressor section 4, 5 as well as a rotor part 9 of the fan section 2.

[0068] In a twin-spool propulsion system 1, the compressor section 4, 5 may comprise a low-pressure compressor 4 and a high-pressure compressor 5. The turbine section 7, 8 may comprise a high-pressure turbine 8 and a low-pressure turbine 7. The rotor of the high-pressure compressor 5 is rotated by the rotor of the high-pressure turbine 8 via a high-pressure shaft 10. The rotor of the low-pressure compressor 4 and the rotor portion 9 of the fan section 2 are rotated by the rotor of the low-pressure turbine 7 via a low-pressure shaft 11. Thus, the primary body 3 comprises a high-pressure body comprising the high-pressure compressor 5, the high-pressure turbine 8 and the high-pressure shaft 10, and a low-pressure body comprising the fan section 2, the low-pressure compressor 4, the low-pressure turbine 7 and the low-pressure shaft 11.The rotational speed of the high-pressure spool is greater than the rotational speed of the low-pressure spool. In a triple-spool propulsion system 1, the turbine section 7, 8 further comprises an intermediate turbine, positioned between the high-pressure turbine 8 and the low-pressure turbine 7 and configured to drive the rotor of the low-pressure compressor 4 via an intermediate shaft. The fan rotor 9 and the rotor of the high-pressure compressor 5 remain driven by the low-pressure shaft 11 and the high-pressure shaft 10, respectively.

[0069] The low pressure shaft 11 is generally housed, over a section of its length, in the high pressure shaft 10 and is coaxial with the high pressure shaft 10. The low pressure shaft 11 and the high pressure shaft 10 may be co-rotating, i.e. driven in the same direction around the longitudinal axis X. Alternatively, the low pressure shaft 11 and the high pressure shaft are counter-rotating, i.e. driven in opposite directions around the longitudinal axis X. Where appropriate, the intermediate shaft is housed between the high pressure shaft 10 and the low pressure shaft 11. The intermediate shaft and the low pressure shaft 11 may be co-rotating or counter-rotating.

[0070] The fan section 2 comprises at least the fan rotor 9 adapted to be driven in rotation relative to a stator part of the propulsion system 1 by the turbine section 7, 8. Each fan rotor 9 comprises a hub 13 and blades 14 extending radially from the hub 13. The blades 14 of each rotor 9 may be fixed relative to the hub 12 or have a variable pitch. In this case, the root of the blades 14 of each rotor 9 is pivotally mounted along a pitch axis and is connected to a pitch change mechanism 15 mounted in the propulsion system 1, the pitch being adjusted according to the flight phases by a pitch change mechanism 15. The pitch change mechanism 15 is illustrated in broken lines in Figure 1 to show that this feature is optional.

[0071] The fan section 2 may further comprise a fan stator 16, or rectifier, which comprises blades 17 mounted on a hub 18 of the fan stator 16 and have the function of straightening the secondary air flow F2 which flows out of the fan rotor 9. The blades 17 of the fan stator 18 may be fixed relative to the hub 18 or have a variable pitch. In a similar manner to the rotor blades 14, the root of the stator blades 17 is pivotally mounted along a pitch axis X and is connected to a pitch change mechanism 15a, which is generally separate from that of the fan rotor 9, the pitch being adjusted according to the flight phases by the pitch change mechanism.

[0072] In order to improve the propulsive efficiency of the propulsion system 1 and to reduce its specific consumption as well as the noise emitted by the fan section 2, the propulsion system 1 has a high bypass ratio. By high bypass ratio, we mean here a bypass ratio greater than or equal to 10, for example between 10 and 80 inclusive. To calculate the bypass ratio, the mass flow rate of the secondary air flow F2 and the mass flow rate of the primary air flow F1 are measured when the propulsion system 1 is stationary, uninstalled, in takeoff mode in a standard atmosphere (as defined by the International Civil Aviation Organization (ICAO) manual, Doc 7488 / 3, 3 eedition) and at sea level (conditions known as SLS, for Seal Level Standard). It should be noted that, in the present application, the parameters (pressure, flow rate, thrust, speed, etc.) are systematically determined under these conditions. By "not installed", it is understood here that the measurements are carried out when the propulsion system 1 is in a test bench (and not installed on an aircraft 100), the measurements then being simpler to carry out. The distances (length, radius, diameter) are, on the other hand, measured at ambient temperature (approximately 20°C) when the propulsion system 1 is cold, that is to say when the propulsion system has been stopped for a sufficient period for the parts of the propulsion system to be at ambient temperature.

[0073] The fan rotor 9 is decoupled from the low-pressure shaft 11 using a reduction mechanism 19, placed between an upstream end of the low-pressure shaft 11 and the fan rotor 9, in order to independently optimize their respective rotational speed. In this case, the propulsion system 1 further comprises an additional shaft, called the fan shaft 20. The low-pressure shaft 11 connects the low-pressure turbine 7 to an inlet of the reduction mechanism 19 while the fan shaft 20 connects the outlet of the reduction mechanism 19 to the fan rotor 9. The fan rotor 9 is therefore driven by the low-pressure shaft 11 via the reduction mechanism 19 and the fan shaft 20 at a rotational speed lower than the rotational speed of the low-pressure turbine 7.

[0074] This decoupling makes it possible to reduce the rotational speed and the pressure ratio of the fan rotor 9 and to increase the power extracted by the low-pressure turbine 7. Indeed, the overall efficiency of the propulsion systems is conditioned to the first order by the propulsive efficiency, which is favorably influenced by a minimization of the variation in kinetic energy of the air when passing through the propulsion system 1. In a propulsion system 1 with a high bypass ratio, the majority of the flow generating the propulsive force is constituted by the secondary air flow F2 of the propulsion system 1, the kinetic energy of the secondary air flow F2 being mainly affected by the compression that the secondary air flow F2 undergoes when passing through the fan section 2. The propulsive efficiency and the pressure ratio of the fan section 2 are therefore linked: the lower the pressure ratio of the fan section 2, the better the propulsive efficiency will be.In order to optimize the propulsive efficiency of the propulsion system 1, the pressure ratio of the fan, which corresponds to the ratio between the average pressure at the outlet of the fan stator 17 (or, in the absence of a stator, of the fan rotor 9) and the average pressure at the inlet of the fan rotor 9, is less than or equal to 1.70, preferably less than or equal to 1.50, for example between 1.05 and 1.45. The average pressures are measured here over the height of the blade 14 (from the surface which radially delimits the flow path at the inlet of the fan rotor 9 to the tip 21 of the fan blade 14).

[0075] The propulsion system 1 is configured to provide a thrust of between 18,000 Ibf (80,068 N) and 51,000 Ibf (22,2411 N), preferably between 20,000 Ibf (88,964 N) and 35,000 Ibf (15,5688 N).

[0076] The fan section 2 may be ducted or unducted. In the case of a ducted fan section 2, the fan section 2 comprises a fan casing 12 and the fan rotor 9 is housed in the fan casing 12.

[0077] A ducted fan section 2 comprises a fan rotor 9 extending upstream of a fan stator. The vanes of the fan stator are then generally called outlet vanes (“Outlet Guide Vane” or “OGV” in English) and have a fixed pitch relative to the hub of the fan stator. Furthermore, the bypass ratio of the propulsion system 1 is preferably greater than or equal to 10, for example between 10 and 35 inclusive, preferably between 10 and 18 inclusive. Note that, when the bypass ratio is greater than or equal to 25, the fan rotor 9 is preferably of variable pitch. The peripheral speed at the tip 21 of the vanes of the fan rotor 9 may also be between 260 m / s and 400 m / s. The vanes 14 of the fan rotor 9 may be fixed or have a variable pitch. The blower pressure ratio can then be between 1.20 and 1.45.

[0078] In an unducted fan section 2, the fan section 2 is not surrounded by a fan casing. Since the fan section 2 is unducted, the blades 14 of the fan rotor 9 have a variable pitch. Propulsion systems comprising at least one unducted fan rotor 9 are known by the terms “open rotor” or “unducted fan”. The propulsion system 1 may comprise two unducted and counter-rotating fan rotors 9. Such a propulsion system 1 is known by the acronym CROR for “Contra-Rotating Open Rotor” or UDF for “Unducted Double Fan”. The fan rotor(s) 9 may be placed at the rear of the primary body 3 so as to be of the pusher type or at the front of the primary body 3 so as to be of the tractor type.Alternatively, the propulsion system 1 may comprise a single unducted fan rotor 9 and an unducted fan stator 16 (straightener). Such a propulsion system 1 is known by the English acronym USF for “Unducted Single Fan”. In the case of a propulsion system 1 of the USF type, the blades 17 of the straightener 16 are fixed in rotation relative to the axis X of rotation of the upstream fan rotor 9 and consequently do not undergo centrifugal force. The blades 17 of the straightener 16 are also variable-pitch.

[0079] The removal of the fairing around the fan section 2 makes it possible to increase the bypass ratio very significantly without the propulsion system 1 being penalized by the mass of the casings or nacelles intended to surround the fan section 2. The bypass ratio of the propulsion system 1 comprising an unfaired fan section 2 is thus greater than or equal to 40, for example between 40 and 80 inclusive. The peripheral speed at the tip 21 of the blades 14 of the fan rotor(s) 9 may also be between 210 m / s and 260 m / s. The fan pressure ratio may then preferably be between 1.05 and 1.20.

[0080] The reduction mechanism 19 may comprise, for example, a reduction mechanism 19 with an epicyclic gear train, for example of the “epicyclic” or “planetary” type, single-stage or two-stage.According to a first variant, the reduction mechanism 19 may be of the planetary type (“star” in English) (Figure 3) and comprise a sun gear 19a (input of the reduction mechanism 19), centered on an axis X of rotation of the reduction mechanism 19 (generally confused with the longitudinal axis X) and configured to be driven in rotation by the low pressure shaft 11, a ring gear 19b (output of the reduction mechanism 19) coaxial with the sun gear 19a and configured to drive the fan shaft 20 in rotation around the axis X of rotation, and a series of satellites 19c distributed circumferentially around the axis X of rotation between the sun gear 19a and the ring gear 19b, each satellite 19c being meshed internally with the sun gear 19a and externally with the ring gear 19b.The series of satellites 19c is mounted on a planet carrier 19d which is fixed relative to a stator part 19e of the propulsion system 1, for example relative to a casing of the compressor section 4, 5. According to a second variant, the reduction mechanism 19 can be epicyclic (“planetary” in English) (Figure 4), in which case the crown 19b is fixedly mounted on the stator part 19e of the propulsion system 1 and the fan shaft 20 is driven in rotation by the planet carrier 19d (which is therefore movable in rotation relative to a stator part 19e of the propulsion system 1, for example relative to a casing of the compressor section 4, 5).

[0081] Whatever the configuration of the reduction mechanism 19, the diameter of the crown 19b and the planet carrier 19d are greater than the diameter of the sun gear 19a, so that the rotation speed of the fan rotor 9 is lower than the rotation speed of the low pressure shaft 11.

[0082] In order to optimize the propulsion system 1 so as to increase its efficiency without mechanically and / or thermally overloading the low pressure turbine 8, the propulsion system is dimensioned so that a rotation speed of the low pressure shaft 11 complies with the following formula: 1.0 3 where: Ni is the rotational speed of the low pressure shaft 11, in revolutions per minute (rpm);

[0083] Te is the inlet temperature of the low pressure turbine 8, in degrees Celsius (°C);

[0084] Se is the inlet section of the low pressure turbine 8, in square meters (m 2) ; and a = - 0.056 83.356 f— xm) 2 . \min 7

[0085] Preferably, formula (1) applies when T e > 700°C to take into account the creep of the low pressure turbine rotor 7. The inlet temperature of the low pressure turbine Te can be between 950°C and 1230°C. The inlet section of the low pressure turbine can be between 0.08 m 2 and 0.35 m 2 . As previously stated, all of these parameters are determined when the propulsion system 1 is stationary, uninstalled, in takeoff mode in a standard atmosphere (as defined by the International Civil Aviation Organization (ICAO) manual, Doc 7488 / 3, 3 e edition) and at sea level.

[0086] The Applicant has in fact noticed that it was necessary to take into account the size of the low pressure turbine 8 to set the rotation speed of the low pressure shaft 11 if it was not desired to overload the low pressure turbine 8. The formula (1) presented below thus makes it possible to obtain a compromise between a high rotation speed of the low pressure shaft 11, and therefore an efficient low pressure turbine 8 and low pressure compressor 4, and acceptable mechanical integration and loading constraints of the low pressure turbine 8.

[0087] Thus, and as is apparent from formula (1), when the inlet section Se of the low pressure turbine 8 is high, the rotational speed of the low pressure shaft 11 is reduced, and therefore less efficient. This also generates connection problems between the low pressure turbine 8 and the high pressure turbine 7 and therefore pressure losses which further reduce the efficiency of the low pressure turbine 8. In addition, since the low pressure turbine 8 is bulky, the propulsion system 1 is more difficult to integrate under the wing of an aircraft 100. It is indeed necessary to ensure a flow between the wing and the nacelle of the primary body 3. However, when the low pressure turbine 8 is very bulky, it becomes necessary to axially advance the propulsion system 1 relative to the wing in order to maintain this flow, which increases the moment applied by the propulsion system 1 on its attachment to the aircraft 100 (pylon) and creates flutter phenomena.

[0088] Preferably, to ensure the integration of the propulsion system 1 under the wing.

[0089] Conversely, when the inlet section Se of the low pressure turbine 8 is reduced, the rotational speed of the low pressure shaft 11 is very high. The low pressure turbine 8 and the low pressure compressor 4 are, of course, more efficient. However, this gain in efficiency is obtained at the expense of, on the one hand, the mechanical strength of the attachment of the blades of the low pressure turbine 8 to the rotor disc 8b depending on the number of blades, and on the other hand, the feasibility of the low pressure turbine 8. The mechanical load in fact becomes too high at the attachment of the disc of the low pressure turbine 8 to the low pressure shaft 11. In addition, since the inlet section Se is small, the average radius of the low pressure turbine 8 is necessarily reduced, which reduces the space available under the low pressure turbine 8. The integration of the bearings and the corresponding enclosure(s) under the low pressure turbine 8 then becomes critical.

[0090] Thus, in order to guarantee the mechanical strength and feasibility of the low pressure turbine 8, the propulsion system 1 can further be configured so that the drive speed of the low pressure turbine 8 also complies with the following formula: knowing that 55 is expressed in xm 2 and that the speed limit (“redline speed” in English) of the low pressure shaft 11, which corresponds to the absolute maximum speed likely to be encountered by the low pressure shaft 11 during the entire flight (according to the European certification regulation EASA CS-E 740 (or according to the American certification regulation 14-CFR Part 33.87)), is between 8500 rpm and 12000 rpm, preferably between 9000 rpm and 11000 rpm. The limiting speed corresponds to the maximum rotation speed when the propulsion system is healthy (and potentially at the end of its life). It is therefore likely to be reached by the low pressure shaft 11 in flight conditions. This limiting speed is part of the data declared in the engine certification ("type certificate data sheet" in English).Indeed, this rotation speed is usually used as a reference speed for the dimensioning of propulsion systems 1 and in certain certification tests (such as blade loss or rotor integrity tests).

[0091] Low pressure turbines 8 complying with formula (1) may then have a hub-to-head ratio at the inlet, which corresponds to the ratio between the external radius R1 of the low pressure turbine 8 and the external radius of the hub R2 of the low pressure turbine 8 at the inlet of the low pressure turbine, greater than 0.75 and less than 0.90. The external radius R1 of the low pressure turbine 8 and the external radius R3 of the hub of the low pressure turbine 8 are measured here in a plane normal to the longitudinal axis X passing through the intersection between the leading edge 8c and the tip 8d of the blades of the most upstream rotor of the low pressure turbine 8 (i.e., of the first stage of the low pressure turbine 8). The external radius R1 of the low pressure turbine 8 corresponds to the distance, in this plane, between the tip 8b of the blades of the rotor of the low pressure turbine 8 and the axis of rotation X of the low pressure turbine 8.The external radius of the hub R2 corresponds to the distance, in this same plane, between the external radial surface of the hub (which radially delimits the flow vein in the low pressure turbine 8) and the axis of rotation X of the low pressure turbine 8. As indicated previously, these radii R1, R2 are measured when the propulsion system 1 is cold.

[0092] A low pressure turbine 8 respecting formula (1) and whose hub-head ratio at the inlet is strictly greater than 0.75 facilitates the integration of the bearings of the low pressure shaft 11.

[0093] Low pressure turbines 8 complying with formula (1) may also have a hub-to-head ratio at the outlet, which corresponds to the ratio between the external radius R1' of the low pressure turbine 8 and the external radius of the hub R2' of the low pressure turbine 8 at the outlet of the low pressure turbine, greater than 0.55 and less than 0.75. The external radius R1 of the low pressure turbine 8 and the external radius R3 of the hub of the low pressure turbine 8 are measured here in a plane normal to the longitudinal axis X passing through the intersection between the leading edge 8c and the tip 8d of the blades of the most downstream rotor of the low pressure turbine 8 (i.e., of the last stage of the low pressure turbine 8). The external radius R1' of the low pressure turbine 8 corresponds to the distance, in this plane, between the tip 8b of the blades of the rotor of the low pressure turbine 8 and the axis of rotation X of the low pressure turbine 8, when the propulsion system 1 is at rest.The external radius of the hub R2' corresponds to the distance, in this same plane, between the external radial surface of the hub (which radially delimits the flow vein in the low pressure turbine 8) and the axis of rotation X of the low pressure turbine 8, when the propulsion system 1 is at rest.

[0094] Such low-pressure turbines 8 then have an optimized inlet section and an outlet section. Indeed, the smaller the hub-to-head ratio at the inlet and outlet, the smaller the external diameter of the low-pressure turbine 8 (at iso-section). A hub-to-head ratio at the inlet and outlet included in the intervals described above thus makes it possible to optimize the low-pressure turbine 8, and in particular its rotation speed Ni (which is adapted to the inlet Se and outlet sections which expand the gases at the outlet of the combustion chamber 5), in a suitable size. The sizing of the low-pressure turbine 8 so as to obtain hub-to-head ratios included in these intervals therefore makes it possible to make the low-pressure turbine 8 more efficient, and therefore to reduce the specific consumption of the propulsion system 1, without penalizing its mechanical or thermal load of the low-pressure turbine 8.

[0095] Low pressure turbines 8 respecting formula (1) and whose rotation speed Ni remains less than or equal to x 10 3 can furthermore be made from conventional materials, such as at least one of the following materials: a nickel-base alloy with directed solidification such as a DS200 + Hf alloy, a nickel-base alloy without directed solidification such as a Rene77 alloy, a nickel-base alloy with a monocrystalline structure such as a first generation superalloy (AM1), an intermetallic alloy such as a TiAI alloy obtained by casting, additive manufacturing or forging.

[0096] A propulsion system 1 complying with formula (1) may comprise a dual-body propulsion system in which the low-pressure turbine 8 comprises at least three stages and at most five stages and the low-pressure compressor 4 comprises at least two stages and at most four stages. The high-pressure turbine 7 may then be two-stage and the high-pressure compressor 5 may comprise at least eight stages and at most eleven stages.

[0097] In order to obtain a rotation speed Ni in accordance with formula (1), the reduction ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11. In the case of a propulsion system 1 comprising a shrouded fan rotor 9, the reduction ratio may be greater than or equal to 2.7 and less than or equal to 6.0, typically around 3.0. In the case of a propulsion system 1 comprising an unshrouded fan rotor, the reduction ratio may be between 9.0 and 11.0.

[0098] The propulsion system 2 further has an overall compression ratio, which corresponds to the pressure ratio between the pressure at the outlet of the high-pressure compressor 5 and the pressure at the inlet of the fan rotor 9 (measured at the level of the foot of the fan rotor 9), may be greater than or equal to 40 and less than or equal to 70, preferably greater than or equal to 44 and less than or equal to 55.

[0099] In order to optimize the performance of the propulsion system 1 in terms of specific consumption, mass and drag, while guaranteeing the possibility of integrating the propulsion system 1 into an aircraft 100, the thrust density per blade 14 of the fan rotor 9 is greater than or equal to 5.0 x 10 4 and less than or equal to 17.0 x 10 4 N / m 2 where the thrust density per blade 14 is defined by the following formula: 100 and where: FN is the thrust generated by the fan rotor 9 and is expressed in Newton (N); n is the number of blades 14 in the fan rotor 9; and D is the diameter of the fan rotor 9, measured in a plane normal to the axis X of rotation at an intersection between a tip 21 and a leading edge 22 of the blades 14 of the fan rotor 9, and is expressed in meters (m). Note that since Figures 1 and 2 are partial views, the diameter D is only partially visible.

[0100] When the propulsion system 1 comprises a ducted fan, the thrust density per blade of the fan rotor 9 is preferably greater than or equal to 14.0 x 10 4 N / m 2 and less than or equal to 17.0 x 10 4 N / m 2 . When the propulsion system 1 comprises an unducted fan, the thrust density per blade of the fan rotor 9 is preferably greater than or equal to 5.0 x 10 4 N / m 2and less than or equal to 10.0 x 10 4 N / m 2 .

[0101] Indeed, the Applicant has noticed that, when the thrust density is less than 5.0 x 10 4 N / m 2 , it was difficult to integrate propulsion system 1 because it was too bulky, had too much mass and generated excessive drag. Furthermore, when the thrust density is greater than 17.0 x 10 4 N / m 2 , the performance of the propulsion system 1 in terms of specific consumption is degraded. The sizing of the propulsion system 1 so that the thrust density per blade 14 of the fan rotor 9 is between 14.0 x 10 4 and 17.0 x 10 4 N / m 2 in the case of a ducted fan and between 5.0 x 10 4 and 10.0 x 10 4 N / m 2in the case of an unducted fan therefore makes it possible to obtain a compromise between the integration and the performance of the propulsion system 1 when the propulsion system 1 comprises a reduction mechanism 19 and has a high bypass ratio. Such a thrust density range per blade 14 is furthermore compatible with a fan pressure ratio of less than 1.45, which makes it possible to optimize the propulsive efficiency of the propulsion system 1.

[0102] For example, a propulsion system 1 according to the invention comprising a shrouded fan rotor and whose thrust density per fan blade 14 is equal to 15 x 10 4 N / m 2 has a specific consumption lower by 5% compared to the same propulsion system whose thrust density per fan blade is equal to 21 x 10 4 N / m 2The thrust density per blade of the propulsion system 1 is influenced to the first order by the diameter D of the fan rotor and the pressure ratio of the fan section 2. The bypass ratio (at a defined diameter D), the overall compression ratio and the number of stages in the compression and turbine sections generally have little or no impact on the thrust density per blade 14.

[0103] Thus, the dimensioning and manufacturing of the propulsion system 1 comprising a shrouded fan so as to obtain a thrust density per blade 14 of between 14.0 x 10 4 N / m 2 and 17.0 x 10 4 N / m 2can be achieved by first setting the thrust (FN) that is to be generated with the fan section 2 and by modifying the diameter (D) of the fan rotor (and therefore the pressure ratio of the fan section 2). Compared to a propulsion system with a conventional reduction mechanism, the diameter D can for example be increased and the pressure ratio of the fan 2 can be reduced. The number of fan blades 14 (n) and the rotational speed of the fan rotor 9 can also be adapted in order to meet performance, acoustic and integration requirements. Depending on the aerodynamic characteristics of the fan section 2, the propulsion system 1 can be modified so as to integrate a pitch change mechanism 15, 15a making it possible to adapt the pitch of the blades 14 of the rotor 9 (and possibly the blades 16 of the stator 17) of the fan section 2.Finally, the thermodynamic cycle is adapted to the various parameters thus dimensioned (fan diameter, number of blades, pressure rate of the fan section 2, etc.) of the propulsion system 1: in particular the flow rate of the gas generator can be reduced and the reduction rate of the reduction mechanism 19 can be increased.

[0104] In the case of an unducted fan, the sizing and manufacturing of the propulsion system 1 so as to obtain a thrust density per blade 14 of between 5.00 x 10 4 N / m 2 and 10.0 x 10 4 N / m 2can be achieved by first setting the thrust (FN) that is desired to be generated with the fan section 2 and by modifying the diameter (D) of the fan rotor (and therefore the pressure ratio of the fan section 2) so as to obtain such thrust. Compared to a propulsion system with a conventional unducted fan and reduction mechanism, the diameter D can for example be slightly reduced to allow the integration of the propulsion system under the wing, and the pressure ratio of fan 2 can be adapted accordingly in order to obtain the desired thrust. In a similar manner to the ducted propulsion system, the number of fan blades 14 (n) and the rotational speed of the fan rotor 9 can then be adapted in order to meet performance, acoustic and integration requirements.Depending on the aerodynamic characteristics of the fan section 2, the propulsion system 1 can be modified so as to integrate a pitch change mechanism 15, 15a making it possible to adapt the pitch of the blades 14 of the rotor 9 (and possibly the blades 16 of the stator 17) of the fan section 2. Furthermore, depending on the integrated performance balance (fuel consumption balance of the propulsion system 1 integrated in the aircraft (mass, specific consumption, drag)) and the aircraft constraints (in terms of integration and program constraints), the fan section 2 can comprise a single fan rotor 9 or two counter-rotating fan rotors 9. Finally, the thermodynamic cycle is adapted to the various parameters thus dimensioned (fan diameter, number of blades, pressure ratio of the fan section 2, overall compression ratio, etc.) of the propulsion system 1: in particular the flow rate of the gas generator can be reduced and the reduction rate of the reduction mechanism 19 can be increased.

[0105] For thrust densities per blade 14 between 5.0 x 10 4 and 17.0 x 10 4 N / m 2 , the diameter D of the fan rotor can then be between 80 inches (203.2 cm) and 185 inches (469.9 cm) inclusive. When the fan rotor 9 is shrouded, the diameter D is preferably between 85 inches (215.9 cm) and 120 inches (304.8 cm) inclusive, for example of the order of 90 inches (228.6 cm), which allows the propulsion system 1 to be integrated in a conventional manner, in particular under the wing of an aircraft. When the fan rotor 9 is unshrouded, the diameter D is preferably greater than or equal to 100 inches (254 cm), for example between 120 inches (304.8 cm) and 156 inches (396.2 cm).

[0106] The fan rotor 9 further comprises at least twelve blades 14 and at most twenty-four blades 14, preferably at least sixteen blades 14 and at most twenty-two blades 14. The number of blades 16 in the fan stator 17 depends on the acoustic criteria defined for the propulsion system 1 and is at least equal to the number of blades 14. In order to further improve the propulsive efficiency of the propulsion system 1, the power density per blade 14 of the fan rotor 9 is greater than or equal to 3.65 x 10 6 and less than or equal to 22.0 x 10 6 W / m 2 , where the power density per blade 14 of the fan rotor 9 is defined by the following formula: 100 where the blower power corresponds to the power of the fan rotor 9 and is expressed in Watts (W).

[0107] In the case of a propulsion system 1 comprising a shrouded fan section 2, the power density per fan blade is preferably greater than or equal to 16.0 x 10 6 and less than or equal to 22.0 x 10 6 W / m 2 ,

[0108] In the case of a propulsion system 1 comprising an unducted fan section 2, the power density per fan blade is preferably greater than or equal to 3.65 x 10 6 and less than or equal to 7.50 x 10 6 W / m 2 ,

[0109] The fan rotor 9 also has a hub-to-head ratio of between 0.22 and 0.32. In the case of a fixed-pitch fan rotor, the hub-to-head ratio may be between 0.24 and 0.32. In the case of a variable-pitch fan rotor, the hub-to-head ratio is preferably between 0.26 and 0.32 to allow the integration of the pitch change mechanism 15. The hub-to-head ratio corresponds to the ratio between the internal radius Ri and the external radius R e of the fan rotor 9. The internal radius Ri corresponds to the distance between the axis of rotation X and the point of intersection between the leading edge 22 and the surface which radially delimits on the inside the flow vein at the inlet of the fan rotor 9 (and corresponds to the point of connection of the leading edge 22 with the aerodynamic surface of a platform of the fan rotor 9). The external radius R ecorresponds to the distance between the axis of rotation X and the point of intersection between the leading edge 22 and the tip 21 of the fan blades (and corresponds to half the fan diameter D). The lower the hub-to-head ratio, the more efficient the fan rotor 9 is. However, reducing the hub-to-head ratio of the fan rotor 9 implies an increase in the mechanical load of the hub 13 of the fan rotor 9. The sizing of the fan rotor 9 so that its hub-to-head ratio is between 0.22 and 0.32 makes it possible in particular to obtain a thrust density and a power density per blade 14 in the intervals defined above.

[0110] (a) Propulsion systems with ducted fan:

[0111] Engine 1 is a twin-body propulsion system corresponding to the current technical standard (at the date of filing of this application) which is sought to be improved and which includes a ducted fan section.

[0112] Engine 2 is a double-body propulsion system in accordance with the teaching of the present application which includes a ducted fan.

[0113] The limiting speed Ni of the low pressure shaft of engine 1 does not comply with the claimed formula since it is less than 9,833 rpm. On the other hand, the limiting speed Ni of the low pressure shaft 11 of engine 2 does comply with the claimed formula. It appears that engine 2 has better specific consumption than engine 1, without requiring a significant increase in the inlet temperature of the low pressure turbine. The efficiency of fan section 2 has therefore been improved without needing to further thermally load the low pressure turbine and without modifying the number of stages in the low pressure compressor.

[0114] Similarly, the thrust density per fan blade of engine 1 is greater than 1.70 x 10 5 N / m 2 while engine 2 has a thrust density per fan blade of between 1.40 and 1.70 x 10 5 N / m 2Engine 2 therefore has lower specific consumption than engine 1.

[0115] To move from engine 1 (reference) to engine 2 (compliant with the disclosure), the overall compression ratio was increased as well as the inlet temperature of the high pressure turbine 7, which made it possible to increase the thermal efficiency of the propulsion system 1. In addition, the diameter D of the fan and the bypass ratio BPR were increased, which made it possible to improve the propulsive efficiency and maintain comparable thrust given the reduction in the pressure ratio of the fan 2. Finally, since the temperature of the low pressure turbine 8 was kept stable, it was possible to increase its mechanical loading (Ni 2 S) and therefore reduce its number of stages. (b) Propulsion systems with unducted fan:

[0116] Engine 3 is a twin-spool propulsion system corresponding to the current technical standard (at the date of filing of this application) which is sought to be improved and which includes an unducted fan section. Engine 4 is a twin-spool propulsion system in accordance with the teaching of this application which includes an unducted fan section.

[0117] The limiting speed Ni of the low pressure shaft of engine 3 does not comply with the claimed formula since it is less than 11,453.73 rpm. On the other hand, the limiting speed Ni of the low pressure shaft 11 of engine 4 does comply with the claimed formula. It appears that engine 4 has better compactness than engine 3, without requiring an increase in the inlet temperature of the low pressure turbine.

[0118] Similarly, the thrust density per fan blade of engine 3 is less than 5.0 x 10 4 N / m 2while engine 4 has a thrust density per fan blade of between 5.0 and 10.0 x 10 4 N / m 2 . Engine 4 is therefore more compact than engine 3.

[0119] To move from engine 3 (reference) to engine 4 (compliant with the disclosure), the overall compression ratio was increased as well as the inlet temperature of the high pressure turbine 7, which made it possible to increase the thermal efficiency of the propulsion system 1. The diameter D of the fan and the bypass ratio BPR were reduced, which made it possible to improve the integration of engine 2. The compression ratio of the fan of engine 2 was, however, slightly increased (while remaining below 1.45) to maintain equivalent thrust. Finally, since the temperature of the low pressure turbine 8 was kept stable, it was possible to increase its mechanical loading (Ni 2S) and therefore reduce its number of stages. Since the fan section represents the majority of the mass of an unducted propulsion system 1, its greater compactness allows for a reduction in mass and easier installation on aircraft. These elements allow the aircraft comprising an engine 4 to have lower fuel consumption.

Claims

CLAIMS 1. Aeronautical propulsion system (1) comprising: - a drive turbine (8) connected to a drive shaft (11) rotatable around an axis of rotation (X); - a fan section (2) comprising a fan rotor (9) connected to a fan shaft (20); - a reduction mechanism (19) coupling the drive shaft (11) and the fan shaft (20) in order to drive the fan shaft (20) at a rotational speed lower than the rotational speed of the drive shaft (11); the propulsion system (1) being configured so that a rotational speed of the drive shaft (11) complies with the following formula: . x 10 3 where: Ni is the rotational speed of the drive shaft (11) when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in revolutions per minute (rpm); Te is the temperature at the inlet of the drive turbine (8) when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in degrees Celsius (°C) and is greater than or equal to 700°C; Se is an inlet section of the drive turbine (8), in square meters (m 2 ) ; and a = - 0.056 f \—min7 2 x — °C and fi = 83.356 f \—min xm) 7 2 .

2. Propulsion system (1) according to claim 1, wherein the propulsion system (1) is further configured such that the drive speed of the drive turbine (8) also complies with the following formula:

3. Propulsion system (1) according to one of claims 1 and 2, in which the reduction mechanism (19) has a reduction ratio greater than or equal to 2.5, preferably greater than or equal to 3.0 and less than or equal to 11.

0.

4. Propulsion system (1) according to one of claims 1 to 3, in which the drive turbine (8) comprises at least three stages and at most five stages.

5. Propulsion system (1) according to one of claims 1 to 4, in which the drive turbine (8) has a hub-to-head ratio at the input greater than 0.75 and less than 0.

90.

6. Propulsion system (1) according to one of claims 1 to 5, in which the drive turbine (8) has a hub-to-head output ratio greater than 0.55 and less than 0.

75.

7. Propulsion system (1) according to one of claims 1 to 6, further comprising a compressor (4) connected directly to the drive turbine (8) by the drive shaft (11) so that a rotational speed of the compressor (4) is equal to the rotational speed of the drive shaft (11), the compressor (4) comprising at least two stages and at most four stages.

8. Propulsion system (1) according to one of claims 1 to 7, further comprising a second turbine (7) and a second compressor (5) connected via a second shaft (10), the second shaft (10) rotating faster than the drive shaft (11), the second turbine (7) being two-stage.

9. Propulsion system (1) according to claim 8, wherein the second compressor (5) comprises at least eight stages and at most eleven stages.

10. Propulsion system (1) according to one of claims 1 to 9, wherein a thrust density per blade (14) of the fan rotor (9) of the propulsion system (1) is greater than or equal to 5.0 x 10 4 and less than or equal to 17.0 x 10 4 N / m 2 where the thrust density per blade (14) is defined by the following formula: 100 and where: FN is the thrust generated by the fan rotor (9) and is measured when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in Newton (N); n is the number of blades (14) in the fan rotor (9); and D is the diameter of the fan rotor (9), measured in a plane normal to the axis of rotation (X) at an intersection between a tip (21) and a leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in meters (m).

11. Propulsion system (1) according to one of claims 1 to 10, in which a power density per blade (14) of the fan rotor (9) is greater than or equal to 3.65 x 10 6 and less than or equal to 22.0 x 10 6 W / m 2 , where the power density per blade (14) of the fan rotor (9) is defined by the following formula: . 100 and where: the fan power corresponds to the power of the fan rotor (9) and is measured when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in Watts (W); n is the number of blades (14) in the fan rotor (9); and D is the diameter of the fan rotor (9), measured in a plane normal to the axis of rotation (X) at an intersection between a tip (21) and a leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in meters (m).

12. Propulsion system (1) according to one of claims 1 to 11, in which the fan section (2) further has a fan compression ratio, corresponding to a pressure ratio between an outlet of the fan rotor (9) and an inlet of the fan rotor (9) less than or equal to 1.45, preferably less than or equal to 1.

30.

13. Propulsion system (1) according to one of claims 1 to 12, wherein the diameter of the fan rotor (9) is between 80 inches (203.2 cm) and 185 inches (469.9 cm) inclusive, preferably between 85 inches (215.9 cm) and 120 inches (304.8 cm) inclusive, for example of the order of 90 inches (228.6 cm).

14. Propulsion system (1) according to one of claims 1 to 13, in which the fan section (2) is faired and a dilution ratio of the propulsion system (1) is greater than or equal to 10, for example between 10 and 35 inclusive, preferably between 10 and 18 inclusive.

15. Propulsion system (1) according to one of claims 1 to 14, in which the fan section (2) is ducted and a peripheral speed at the tip of the blades of the fan rotor (9), when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level, is between 260 m / s and 400 m / s.

16. Propulsion system (1) according to one of claims 1 to 15, in which the fan section is ducted and a thrust density per blade (14) of the fan rotor (9) of the propulsion system (1) is greater than or equal to 14.0 x 10 4 and less than or equal to 17.0 x 10 4 N / m 2 where the thrust density per blade (14) is defined by the following formula: 100 and where: FN is the thrust generated by the fan rotor (9) and is measured when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in Newton (N); n is the number of blades (14) in the fan rotor (9); and D is the diameter of the fan rotor (9), measured in a plane normal to the axis of rotation (X) at an intersection between a tip (21) and a leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in meters (m).

17. Propulsion system (1) according to one of claims 1 to 13, in which the fan section (2) is unducted and a bypass ratio of the propulsion system (1) is greater than or equal to 40, for example between 40 and 80 inclusive.

18. Propulsion system (1) according to one of claims 1 to 13 or 17, in which the fan section (2) is unducted and a peripheral speed at the tip of the blades of the fan rotor (9), when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level, is between 210 m / s and 260 m / s.

19. Propulsion system (1) according to one of claims 1 to 13, 17 or 18, in which the fan section is unducted and a thrust density per blade (14) of the fan rotor (9) of the propulsion system (1) is greater than or equal to 5.0 x 10 4 and less than or equal to 10.0 x 10 4 N / m 2 where the thrust density per blade (14) is defined by the following formula: 100 and where: FN is the thrust generated by the fan rotor (9) and is measured when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in Newton (N); n is the number of blades (14) in the fan rotor (9); and D is the diameter of the fan rotor (9), measured in a plane normal to the axis of rotation (X) at an intersection between a tip (21) and a leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in meters (m).

20. Aircraft (100) comprising at least one propulsion system (1) according to one of claims 1 to 19 fixed to the aircraft (100) via a mast.

21. Method for dimensioning a propulsion system (1) comprising a reduction mechanism (19) coupling a drive turbine (8) and a fan rotor (9) to drive the fan rotor (9) at a speed lower than a speed of the drive turbine (8), the drive turbine (8) configured so that a rotational speed of the drive shaft (11) complies with the following formula: x 10 3 where: Ni is the rotational speed of the drive shaft (11) when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in revolutions per minute (rpm); Te is the temperature at the inlet of the drive turbine (8) when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in degrees Celsius (°C) and is greater than or equal to 700°C; Se is an inlet section of the drive turbine (8), in square meters (m 2 ) ; and a = - 0.056 f \—min7 2 x — °C and fi = 83.356 f \—min xm) 7 2 .

22. A sizing method according to claim 21, wherein the fan section (2) is sized such that a thrust density per blade (14) of the fan rotor (9) of the propulsion system (1) is greater than or equal to 5.5 x 10 4 and less than or equal to 17.0 x 10 4 N / m 2 where the thrust density per blade (14) is defined by the following formula: 100 and where: FN is the thrust generated by the fan rotor (9) and is measured when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in Newton (N); n is the number of blades (14) in the fan rotor (9); and D is the diameter of the fan rotor (9), measured in a plane normal to the axis of rotation (X) at an intersection between a tip (21) and a leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in meters (m).

23. A sizing method according to one of claims 21 and 22, wherein the fan is further sized such that a power density per blade (14) of the fan rotor (9) is greater than or equal to 3.65 x 10 6 and less than or equal to 22.0 x 10 6 W / m 2 , where the power density per blade (14) of the fan rotor (9) is defined by the following formula: . 100 and where: the fan power corresponds to the power of the fan rotor (9) and is measured when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in Watts (W); n is the number of blades (14) in the fan rotor (9); and D is the diameter of the fan rotor (9), measured in a plane normal to the axis of rotation (X) at an intersection between a tip (21) and a leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in meters (m).

24. Method for manufacturing a propulsion system (1) comprising the following steps: dimensioning the propulsion system (1) according to one of claims 21 to 23; and manufacturing the propulsion system (1).

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