Aero-propulsion system with improved propulsion efficiency

Abstract

The invention relates to an aero-propulsion system (1) having a thrust density per blade (14) of the fan rotor (9) greater than or equal to 5.0 × 104 and less than or equal to 10.0 × 104 N / m2, wherein the thrust density per blade (14) is defined by the following formula: (Formula I) and wherein: FN is the thrust generated by the fan rotor (9), measured when the propulsion system (1) is stationary in take-off rating in a standard atmosphere and at sea level, expressed in Newton (N); n is the number of the blades (14) in the fan rotor (9); and D is the diameter of the fan rotor (9), measured in a plane normal to the rotation axis (X) at an intersection between a vertex (21) and a leading edge (22) of the blades (14) of the fan rotor (9), expressed in metres.

Description

[0001] DESCRIPTION

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

[0003] TECHNICAL FIELD

[0004] The present application generally relates to the field of propulsion systems, and more particularly to aeronautical propulsion systems comprising an unducted fan and having a high, or even very high, bypass ratio.

[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] However, improving the BPR bypass ratio and the fan pressure ratio involves increasing the fan diameter and, by extension, the external dimensions of the propulsion system (and therefore its mass and drag), which makes the integration of the propulsion system more difficult in addition to increasing its mass and specific consumption. The flow rate in the high-pressure body and the size of the high-pressure body are also reduced, which imposes limitations on the low-pressure body. In particular, since the low-pressure shaft is housed in the high-pressure shaft, reducing the size of the low-pressure body (and therefore the high-pressure shaft) involves reducing the diameter of the high-low-pressure shaft, which can therefore become supercritical. However, a supercritical shaft includes a bending deformation mode in an operating range of the propulsion system.The dynamics of the low pressure shaft must then be controlled so that the deformation mode does not appear in a stabilized range to avoid the risk of damaging the propulsion system.

[0010] EXPOSED

[0011] An aim of the present application is to optimize the performance of the propulsion system in terms of specific consumption, mass and drag, while ensuring the possibility of integrating the propulsion system into an aircraft.

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

[0013] - a drive shaft movable in rotation around an axis of rotation;

[0014] - a blower shaft;

[0015] - a fan section comprising an unducted fan rotor rotated by the fan shaft, the fan rotor comprising a plurality of blades;

[0016] - 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; a thrust density per blade of the fan rotor of the propulsion system being 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); n is the number of blades (14) in the fan rotor; and

[0017] 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).

[0018] Some preferred but non-limiting features of the aeronautical propulsion system according to the first aspect are the following, taken individually or in combination:

[0019] - a power density per fan rotor blade is greater than or equal to 3.65 x 10 6 and less than or equal to 7.50 x 10 6 W / m 2 , where the power density per fan rotor blade is defined by the following formula: . 100 and where the fan power corresponds to the fan rotor power 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);

[0020] - 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, preferably less than or equal to 1.30;

[0021] - the diameter of the fan rotor is between 80 inches (203.2 cm) and 185 inches (469.9 cm) inclusive, preferably between 120 inches (304.8 cm) and 170 inches (431.8 cm) inclusive, for example of the order of 156 inches (396.2 cm);

[0022] - a dilution ratio of the propulsion system is greater than or equal to 40, for example between 40 and 80 inclusive;

[0023] - 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;

[0024] - a fan rotor hub-to-head ratio is between 0.22 and 0.34;

[0025] - the propulsion system further comprises a drive turbine and a compressor connected directly by the drive shaft, the drive turbine comprising at least three and at most five stages;

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

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

[0028] - the high-pressure compressor comprises at least eight and at most eleven stages; and / or

[0029] - the fan rotor comprises at least ten fan blades and at most eighteen fan blades, preferably at least twelve fan blades and at most sixteen fan blades.

[0030] 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.

[0031] According to a third aspect, the present application provides a method for dimensioning a propulsion system comprising a reduction mechanism coupling a drive shaft and a fan rotor to drive the unducted fan rotor at a speed lower than a speed of the drive shaft. The fan rotor is dimensioned such 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 10.0 x 10 4 N / m 2where the thrust density per fan rotor blade is defined by the following formula: 100 and where: FN is the fan rotor thrust 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

[0032] D is the fan diameter, 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). Optionally, the fan rotor (9) may further be dimensioned such 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 7.50 x 10 6 W / m 2 , where the power density per fan rotor blade is defined by the following formula , .: 100 and where the fan rotor power 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).

[0033] According to a fourth aspect, there is provided a method of manufacturing a propulsion system comprising the following steps:

[0034] - dimension the propulsion system in accordance with the third aspect; and

[0035] - manufacture the propulsion system.

[0036] DESCRIPTION OF FIGURES

[0037] 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:

[0038] Figure 1 is a schematic, partial, sectional view of an exemplary propulsion system according to one embodiment, in which the fan section is unducted;

[0039] Figure 2 is a schematic sectional view of an exemplary planetary reduction mechanism;

[0040] Figure 3 is a schematic sectional view of an exemplary epicyclic reduction mechanism;

[0041] Figure 4 is an example of an aircraft that may include at least one propulsion system according to one embodiment;

[0042] Figure 5 is a flowchart illustrating exemplary steps of a sizing or manufacturing method according to one embodiment.

[0043] Throughout the figures, similar elements have identical references.

[0044] DETAILED DESCRIPTION

[0045] 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).

[0046] 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) 7b, 8b behind which a wheel of moving blades (rotor) 7a, 8a rotates. 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, inside) and external (respectively, outside), respectively, are used in reference to a radial direction such that the internal part or face of an element is closer to the X axis than the external part or face of the same element.

[0047] 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.

[0048] 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.

[0049] 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.

[0050] 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 7 and a low-pressure turbine 8. The rotor of the high-pressure compressor 5 is rotated by the rotor of the high-pressure turbine 7 via a high-pressure shaft 10. The rotor of the low-pressure compressor 4 and the rotor portion 9 (propeller) of the fan section 2 are rotated by the rotor of the low-pressure turbine 8 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 7 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 8 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 7 and the low-pressure turbine 8 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.

[0051] 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.

[0052] 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.

[0053] 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.

[0054] 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.

[0055] The fan rotor 9 is decoupled from the low-pressure shaft 11 by means of 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 8 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 8.

[0056] 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 8. 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).

[0057] 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).

[0058] The fan section 2 may be 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.

[0059] 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. 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.

[0060] 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 2) 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 3), 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).

[0061] 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.

[0062] The reduction ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11, preferably between 9.0 and 11.0.

[0063] The redline speed 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 redline 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 redline 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).

[0064] 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 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 expressed in Newton (N); n is the number of blades 14 in the fan rotor 9; and

[0065] 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 Figure 1 is a partial view, the diameter D is only partially visible.

[0066] When the propulsion system 1 comprises two fan rotors 9, the thrust density per blade 14 of the fan rotor is less than or equal to 4.0 x 10 4 N / m 2 .

[0067] Indeed, the Applicant has noticed that, when the thrust density is less than 5.0 10 4 N / m 2 , it was difficult to integrate the propulsion system 1 because it was too bulky, had too much mass and generated excessive drag. Furthermore, when the thrust density is greater than 10.0 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 5.0 x 10 4 and 10.0 x 10 4 N / m 2 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 interval 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.

[0068] For example, a propulsion system 1 according to the invention comprising an unducted fan rotor and whose thrust density per fan blade 14 is equal to 6.5 x 10 4 N / m 2may have a fan diameter 13% smaller than the same propulsion system with a thrust density per fan blade of 4.5 x 10 4 N / m 2 , which facilitates its integration under the wing without compromising its efficiency. The 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, 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.

[0069] Thus, the dimensioning and manufacturing of the propulsion system 1 so as to obtain a thrust density per blade 14 of between 5.0x 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 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 the fan 2 can be adapted accordingly in order to obtain the desired thrust. 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. 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.

[0070] For thrust densities per blade 14 between 5.0 x 10 4 and 10.0 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, preferably greater than or equal to 100 inches (254 cm), for example between 120 inches (304.8 cm) and 170 inches (431.8 cm), for example 156 inches (396.2 cm), which makes it possible to integrate the propulsion system 1 in a conventional manner, in particular under the wing of an aircraft.

[0071] The fan rotor 9 further comprises at least ten blades 14 and at most eighteen blades 14, preferably at least twelve blades 14 and at most sixteen 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 less than the number of blades 14 of the fan rotor 9.

[0072] 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 7.50 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).

[0073] The fan rotor 9 also has a hub-to-head ratio of between 0.22 and 0.34, which allows 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 0.34 makes it possible in particular to obtain a thrust density and a power density per blade 14 in the intervals defined above.

[0074] A dual-body propulsion system 1 having a thrust density and a power density per blade 14 of the fan rotor 9 in the intervals defined above may in particular comprise a two-stage high-pressure turbine 7, a high-pressure compressor 5 comprising at least eight stages and at most eleven stages, a low-pressure turbine 8 comprising at least three stages and at most five stages and a low-pressure compressor 4 comprising at least two stages and at most four stages.

[0075] Comparative example:

[0076] 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 an unducted fan section.

[0077] Engine 2 is a twin-body propulsion system 1 in accordance with the teaching of the present application comprising an unducted fan section, two shafts and a thrust density per fan blade equal to 6.53 x 10 4 N / m 2 .

[0078] Engine 1 has a thrust density per fan blade greater than 4.7 x 10 4 N / m 2 while engine 2 has a thrust density per fan blade of between 5.0 N / m 2 and 10.0 x 10 4 N / m 2. It appears that, thanks to its improved thrust density per fan blade, the engine 2 has greater compactness and a lower mass than the engine 1 . In the present comparative example, the mass reduction of the fan section 2 is estimated at around twenty percent. Since the fan section represents a third of the mass of the propulsion system 1 , this amounts to reducing the mass of the propulsion system 1 by around 7% with induced effects on the aircraft (overhang mass, reduction in the diameter of the fan rotor 9, etc.). The engine 2 can therefore be more easily installed on an aircraft 100, at equivalent low-pressure turbine inlet temperature 8 and fan thrust.

[0079] To move from engine 1 (reference) to engine 2 (compliant with the disclosure), the fan diameter D and the bypass ratio BPR were reduced, which improved the integration of engine 2. The pressure ratio of the fan section of engine 2 was, however, slightly increased (while remaining below 1.45) to maintain equivalent thrust. In addition, the overall compression ratio was increased without increasing the inlet temperature of the high-pressure turbine 7, which improves the efficiency of the primary body without increasing the thermal load of the low-pressure turbine 8. Finally, since the temperature of the low-pressure turbine 8 was kept stable, it was possible to increase its mechanical loading (Ni 2 S) in order to reduce its number of floors.

Claims

CLAIMS 1. Aeronautical propulsion system (1) comprising: - a drive shaft (11) movable in rotation around an axis of rotation (X); - a fan shaft (20); - a fan section (2) comprising an unducted fan rotor (9) rotated by the fan shaft (20), the fan rotor (9) comprising a plurality of blades (14); - 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); a thrust density per blade (14) of the fan rotor (9) of the propulsion system (1) being 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).

2. Propulsion system (1) according to claim 1, wherein 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 7.50 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).

3. Propulsion system (1) according to one of claims 1 and 2, wherein 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.

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4. Propulsion system (1) according to one of claims 1 to 3, wherein the diameter of the fan rotor (9) is between 80 inches (203.2 cm) and 185 inches (469.9 cm) inclusive, preferably between 120 inches (304.8 cm) and 170 inches (431.8 cm) inclusive, for example of the order of 156 inches (396.2 cm).

5. Propulsion system (1) according to one of claims 1 to 4, in which a dilution ratio of the propulsion system (1) is greater than or equal to 40, for example between 40 and 80 inclusive.

6. Propulsion system (1) according to one of claims 1 to 5, in which a peripheral speed at the tip (21) of the blades (14) 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.

7. Propulsion system (1) according to one of claims 1 to 6, in which a hub-to-head ratio of the fan rotor (9) is between 0.22 and 0.

34.

8. Propulsion system (1) according to one of claims 1 to 7, further comprising a drive turbine (8) and a compressor (4) directly connected by the drive shaft (11), the drive turbine (8) comprising at least three and at most five stages.

9. Propulsion system (1) according to claim 8, wherein the compressor (4) comprises at least two and at most four stages.

10. Propulsion system (1) according to one of claims 1 to 9, further comprising a high pressure turbine (7) and a high pressure compressor (5) connected via a high pressure shaft (10), the high pressure shaft (10) rotating faster than the drive shaft (11), the high pressure turbine (7) being two-stage.

11. Propulsion system (1) according to claim 10, wherein the high pressure compressor (5) comprises at least eight and at most eleven stages.

12. Propulsion system (1) according to one of claims 1 to 11, wherein the fan rotor (9) comprises at least ten fan blades (14) and at most eighteen fan blades (14), preferably at least twelve fan blades (14) and at most sixteen fan blades (14).

13. Aircraft (100) comprising at least one propulsion system (1) according to one of claims 1 to 12 fixed to the aircraft by means of a mast.

14. A method of dimensioning a propulsion system (1) comprising a reduction mechanism (19) coupling a drive shaft (11) and a fan rotor (9) to drive the unducted fan rotor (9) at a speed lower than a speed of the drive shaft (11), wherein the fan rotor (9) is dimensioned such that a thrust density per blade (14) of the fan rotor (9) of the propulsion system (1) is greater than or equal to 5.0 x 10 4 and less than or equal to 10.0 x 10 4 N / m 2 where the thrust density per blade (14) of the fan rotor (9) is defined by the following formula: 100 and where: FN is the thrust 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 Newton (N); n is the number of blades (14) in the fan rotor (9); and D is the fan diameter, measured in a plane normal to the axis 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).

15. A sizing method according to claim 14, wherein the fan rotor (9) 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 7.50 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 power of the fan rotor (9) 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).

16. Method for manufacturing a propulsion system (1) comprising the following steps: dimensioning the propulsion system (1) according to one of claims 14 and 15; and manufacturing the propulsion system (1).

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