Aeronautical propulsion system with improved propulsive efficiency

Abstract

The present invention relates to an aeronautical propulsion system (1) having a thrust density per blade (14) of the fan rotor (9) greater than or equal to 3.3 x 104 and less than or equal to 20.0 x 104 N / m² where the thrust density per blade (14) is defined by the following formula: and where: FN is the thrust generated by the fan rotor (9) and is measured when the propulsion system (1) is stationary in takeoff regime in a standard atmosphere and at sea level and is expressed in Newtons; 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 the level of an intersection between a vertex (21) and a leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in meters.

Description

Title of the invention: Aeronautical propulsion system with improved propulsive efficiency technical field

[0001] The present application relates generally to the field of propulsion systems, and more particularly to aeronautical propulsion systems comprising a shrouded or unshrouded fan and exhibiting a high, or even very high, dilution ratio. STATE OF THE ART

[0002] 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 driven in rotation by the high-pressure turbine via a high-pressure shaft. The fan and, where applicable, the low-pressure compressor are driven in rotation by the low-pressure turbine via a low-pressure shaft.

[0003] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into account the impactful factors in all phases of design and development in order to obtain aeronautical components and products that are less energy-intensive, more environmentally friendly, and whose integration and use in civil aviation have moderate environmental consequences, with the aim of improving the energy efficiency of aircraft.

[0004] Thus, in order to improve the propulsive efficiency of the propulsion system and reduce its specific fuel consumption as well as the noise emitted by the fan section, propulsion systems with a high bypass ratio (BPR, corresponding to the ratio between the secondary airflow rate and the primary airflow rate) have been proposed. To achieve such bypass ratios, the fan section can be decoupled from the low-pressure turbine, thereby allowing their respective rotational speeds to be optimized independently. Generally, 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 rotational speed lower than that of the low-pressure shaft.

[0005] However, improving the BPR dilution ratio and the fan pressure ratio requires increasing the fan diameter and, by extension, the external dimensions of the propulsion system (and therefore its mass and drag), which makes integrating the propulsion system more difficult in addition to increasing its mass and specific fuel consumption. Furthermore, the flow rate in the high-pressure body and the size of the high-pressure body are reduced, which imposes limitations on the low-pressure body. In particular, since the low-pressure shaft is housed within the high-pressure shaft, reducing the size of the low-pressure body (and therefore the high-pressure shaft) requires reducing the diameter of the high-low-pressure shaft, which can consequently become supercritical. A supercritical shaft exhibits a bending deformation mode within a range of propulsion system operation.The dynamics of the low-pressure shaft must then be controlled so that the deformation mode does not appear within a stabilized range, thus avoiding the risk of damaging the propulsion system. Description of the invention

[0006] One 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.

[0007] To this end, according to a first aspect, an aeronautical propulsion system is proposed comprising: - a movable drive shaft rotating around an axis of rotation; - a blower shaft; - a blower section comprising a blower rotor driven in rotation by the blower shaft, the blower rotor comprising a plurality of blades; - a reduction mechanism coupling the drive shaft and the blower shaft in order to drive the blower shaft at a rotational speed lower than the rotational speed of the drive shaft; a thrust density per blade of the blower rotor of the propulsion system being greater than or equal to 3.3 x 104 and less than or equal to 20.0 x 104 N / m2 where the thrust density per blade is defined by the following formula: Thrust density = 7-57 * 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 Newtons (N); n is the number of blades in the blower rotor; and D is the diameter of the blower rotor, measured in a plane normal to the axis of rotation at the intersection between a vertex and a leading edge of the blades of the blower rotor, and is expressed in meters (m).

[0008] Some preferred but non-limiting features of the propulsion system according to the first aspect are the following, taken individually or in combination: - a power density per blade of the blower rotor is greater than or equal to 3.65 x 106 and less than or equal to 22.0 x 106 W / m2, where the power density per blade of the blower rotor is defined by the following formula: . power of the except flante Power density = -------~---*100 and where the power of the blower corresponds to the power of the blower rotor and is measured when the propulsion system is stationary in takeoff regime in a standard atmosphere and at sea level and is expressed in Watts (W); - the blower section also has a blower compression ratio, corresponding to a pressure ratio between a blower rotor outlet and a blower rotor inlet less than or equal to 1.45, preferably less than or equal to 1.30; - The diameter of the blower rotor is between 80 inches and 185 inches inclusive, preferably between 85 inches and 120 inches inclusive, for example, in the order of 90 inches; - the blower 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; - the fan section is shrouded and a peripheral speed at the top of the fan rotor blades, when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level, is between 260 m / s and 400 m / s; - the blower section is not shrouded and a propulsion system dilution ratio is greater than or equal to 40, for example between 40 and 80 inclusive; - the fan section is unfaired and a peripheral speed at the top of the fan rotor blades, when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level, is between 210 m / s and 260 m / s; - a hub-to-head ratio of the blower rotor is between 0.22 and 0.32; - the propulsion system further comprises a drive turbine and a compressor connected directly by the drive shaft, the compressor having a power density greater than or equal to 0.64 and less than or equal to 0.97, where the power density of the compressor is defined by the following formula: Compressor power density = *^^7^ and where: BPR is the dilution rate of the propulsion system and is measured when the propulsion system is stationary in takeoff mode in an atmosphere standard and at sea level and is expressed in pounds of thrust; net is the number of compressor stages; Um is the average peripheral speed of the compressor rotor measured when the propulsion system is stationary in takeoff regime in a standard atmosphere and at sea level and is expressed in pounds of thrust and is expressed in meters per second (m / s); and Rext_moy is the average external radius of the compressor, which is equal to an arithmetic mean of the external radii of the rotors of the low-pressure compressor, the external radius corresponding to a distance between a rotor apex and the axis of rotation, halfway at the apex between a leading edge and a trailing edge of the rotor, and is expressed in meters (m); - the drive turbine comprises at least 3 and at most 5 stages; - the compressor comprises at least 2 and at most 4 stages; - 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; and / or - the high-pressure compressor comprises at least 8 and at most 11 stages.

[0009] 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.

[0010] According to a third aspect, the present application proposes a method for dimensioning or manufacturing a propulsion system comprising a reduction mechanism coupling a drive shaft and a blower rotor to drive the blower rotor at a speed lower than a speed of the drive shaft, in which the blower rotor is dimensioned such that a thrust density per blade of the blower rotor of the propulsion system is greater than or equal to 3.3 x 104 and less than or equal to 20.0 x 104 N / m2 where the thrust density per blade of the blower rotor is defined by the following formula: Thrust density = *100 and where: FN is the blower 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 Newtons (N); n is the number of blades in the blower rotor; and D is the blower diameter, measured in a plane normal to the axis of rotation at the intersection between a vertex and a leading edge of the blower rotor blades, and is expressed in meters (m).

[0011] Some preferred but non-limiting features of the sizing or manufacturing process according to the first aspect are the following, taken individually either dually or in combination: - the blower rotor is further dimensioned so that a power density per blade of the blower rotor is greater than or equal to 3.65 x 106 and less than or equal to 22.0 x 106 W / m2, where the power density per blade of the blower rotor is defined by the following formula: r, . . , . power of the suff lanie Power density = --------M 00 and where the power of the fan rotor 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); - The propulsion system further includes a compressor driven directly by the drive shaft, the compressor being sized such that a compressor power density greater than or equal to 0.64 and less than or equal to 0.97, where the compressor power density is defined by the following formula: Compressor power density = and where: BPR is the propulsion system's dilution ratio and is measured when the propulsion system is stationary in takeoff regime in a standard atmosphere and at sea level and is expressed in pounds of thrust; net is the number of compressor stages; Um is the average peripheral speed of the compressor rotor measured when the propulsion system is stationary in takeoff regime in a standard atmosphere and at sea level and is expressed in pounds of thrust and is expressed in meters per second (m / s); and Rext_moy is the average external radius of the compressor, which is equal to an arithmetic mean of the external radii of the rotors of the low-pressure compressor, the external radius being a distance between a rotor tip and the axis of rotation, halfway at the tip between a leading edge and a trailing edge of the rotor, and is expressed in meters (m). DESCRIPTION OF THE FIGURES

[0012] Other features, objectives and advantages of the invention will become apparent from the following description, which is purely illustrative and not limiting, and which should be read in conjunction with the accompanying drawings on which:

[0013] Fig. 1 is a schematic, partial, cross-sectional view of an example of a propulsion system according to a first embodiment of the invention, in which the blower section is faired;

[0014] Fig. 2 is a schematic, partial, cross-sectional view of an example of a propulsion system according to a first embodiment of the invention, in which the blower section is unfaired;

[0015] Fig. 3 is a schematic cross-sectional view of an example of a planetary reduction mechanism;

[0016] Fig. 4 is a schematic cross-sectional view of an example of an epicycloidal reduction mechanism;

[0017] The [Fig.5] is an example of an aircraft that may include at least one propulsion system conforming to the first or second embodiment of the invention;

[0018] Fig. 6 is a flowchart illustrating examples of steps in a sizing or manufacturing process according to one embodiment of the invention.

[0019] Throughout the figures, similar elements bear identical references. DETAILED DESCRIPTION OF THE INVENTION

[0020] A propulsion system 1 has a principal direction extending along a longitudinal axis X and comprises, from upstream to downstream in the direction of the flow of gases in the propulsion system 1 when in operation, a blower 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 here is an aeronautical propulsion system 1 configured to be fixed on an aircraft 100 by means of a pylon (or mast).

[0021] The compressor section 4, 5 comprises a series of stages, each comprising a rotating blade wheel (rotor) 4a, 5a in front of a fixed blade wheel (stator) 4b, 5b. The turbine section 7, 8 also comprises a series of stages, each comprising a fixed blade wheel (stator) 7a, 8a behind which a rotating blade wheel (rotor) 7b, 8b rotates.

[0022] In the present application, the axial direction corresponds to the direction of the longitudinal axis X, corresponding to the rotation of the shafts of the gas generator, and a radial direction is a direction perpendicular to and passing through this axis X. Furthermore, the circumferential (or lateral, or tangential) direction corresponds to a direction perpendicular to and not passing through the longitudinal axis X. Unless otherwise specified, internal (respectively, inner) and external (respectively, outer), respectively, are used with reference to a radial direction such that the inner part or face of an element is closer to the axis X than the outer part or face of the same element.

[0023] In operation, an airflow F entering the propulsion system 1 is divided between a primary airflow Fl and a secondary airflow F2, which flow from upstream to downstream in the propulsion system 1.

[0024] 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.

[0025] The primary airflow Fl flows into a primary vein inside the body primary 3, passing successively through the compressor section 4, 5, the combustion chamber 6 where it is mixed with fuel to serve as an oxidizer, and the turbine section 7, 8. The passage of the primary airflow Fl through the turbine section 7, 8 receiving energy from the combustion chamber 6 causes a rotation of the rotor of the turbine section 7, 8, which in turn drives the rotation of the rotor of the compressor section 4, 5 as well as a rotor part 9 of the blower section 2.

[0026] In a twin-body propulsion system 1, the compressor section 4, 5 may include a low-pressure compressor 4 and a high-pressure compressor 5. The turbine section 7, 8 may include a high-pressure turbine 8 and a low-pressure turbine 7. The rotor of the high-pressure compressor 5 is driven in rotation 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 blower section 2 are driven in rotation 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 including the high-pressure compressor 5, the high-pressure turbine 8 and the high-pressure shaft 10, and a low-pressure body including the blower 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 body is greater than the rotational speed of the low-pressure body. In a three-body propulsion system 1, the turbine section 7, 8 further includes 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 blower 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.

[0027] The low-pressure shaft 11 is generally housed, along a portion of its length, within 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, that is, driven in the same direction around the longitudinal axis X. Alternatively, the low-pressure shaft 11 and the high-pressure shaft may be counter-rotating, that is, driven in opposite directions around the longitudinal axis X. If applicable, 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.

[0028] The fan section 2 comprises at least the fan rotor 9 adapted to be driven in rotation relative to a stator portion 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 variable pitch. In this case, the base The blades 14 of each rotor 9 are mounted pivoting about a pitch axis and are 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 dashed lines on [Fig.1] to show that this feature is optional.

[0029] The fan section 2 may further include a fan stator 16, or rectifier, which comprises blades 17 mounted on a hub 18 of the fan stator 16 and whose function is to rectify the secondary airflow F2 exiting the fan rotor 9. The fan stator blades 17 may be fixed relative to the hub 18 or have a variable pitch. Similar to the rotor blades 14, the base of the stator blades 17 is pivotally mounted about a pitch axis X and is connected to a pitch-changing 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-changing mechanism.

[0030] In order to improve the propulsive efficiency of propulsion system 1 and reduce its specific fuel consumption as well as the noise emitted by the fan section 2, propulsion system 1 has a high bypass ratio. A high bypass ratio is understood here to be a 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 airflow F2 and the mass flow rate of the primary airflow Fl are measured when 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, 3rd edition) and at sea level. It should be noted that, in this application, the parameters (pressure, flow rate, thrust, speed, etc.) are systematically determined under these conditions.By "not installed", we will understand 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.

[0031] The fan rotor 9 is decoupled from the low-pressure shaft 11 by means of a reduction mechanism 19, positioned between an upstream end of the low-pressure shaft 11 and the fan rotor 9, in order to independently optimize their respective rotational speeds. In this case, the propulsion system 1 further comprises an additional shaft, referred to as 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 thus 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.

[0032] This decoupling makes it possible to reduce the rotational speed and pressure ratio of the fan rotor 9 and to increase the power extracted by the low-pressure turbine 7. Indeed, the overall efficiency of propulsion systems is primarily determined by the propulsion efficiency, which is favorably influenced by minimizing the variation in the kinetic energy of the air as it passes through the propulsion system 1. In a propulsion system 1 with a high bypass ratio, the majority of the flow generating the propulsive force consists of the secondary airflow F2 of the propulsion system 1, the kinetic energy of the secondary airflow F2 being mainly affected by the compression that the secondary airflow F2 undergoes as it passes through the fan section 2. The propulsion 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 propulsion efficiency.In order to optimize the propulsive efficiency of the propulsion system 1, the blower pressure ratio, which corresponds to the ratio between the average pressure at the outlet of the blower stator 17 (or, in the absence of a stator, of the blower rotor 9) and the average pressure at the inlet of the blower 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 duct inside the inlet of the blower rotor 9 to the apex 21 of the blower blade 14).

[0033] The propulsion system 1 is configured to provide a thrust of between 18,000 Ibf (80,068 N) and 51,000 Ibf (22,241 N), preferably between 20,000 Ibf (88,964 N) and 35,000 Ibf (15,568 N).

[0034] The blower section 2 may be shrouded or unshrouded. In the case of a shrouded blower section 2, the blower section 2 comprises a blower housing 12 and the blower rotor 9 is housed in the blower housing 12.

[0035] A shrouded fan section 2 comprises a fan rotor 9 extending upstream of a fan stator. The fan stator blades are then generally called outlet guide vanes (OGVs) and have a fixed pitch relative to the fan stator hub. 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. The peripheral velocity at the tip 21 of the fan rotor blades 9 can also be between 260 m / s and 400 m / s. The fan rotor blades 14 can be fixed or have a variable pitch. The fan pressure ratio can then be between 1.20 and 1.45.

[0036] In an unshrouded fan section 2, the fan section 2 is not enclosed by a fan housing. Since the fan section 2 is unshrouded, the fan blades 14 of the fan rotor 9 have variable pitch. Systems Propulsion systems comprising at least one unshod fan rotor 9 are known by the English terms "open rotor" or "unducted fan." The propulsion system 1 may comprise two unshod, counter-rotating fan rotors 9. Such a propulsion system 1 is known by the English acronym CROR for "Contra-Rotating Open Rotor" or UDF for "Unducted Double Fan." The fan rotor(s) 9 may be positioned at the rear of the primary casing 3 to be of the pusher type or at the front of the primary casing 3 to be of the tractor type. Alternatively, the propulsion system 1 may comprise a single unshod fan rotor 9 and an unshod fan stator 16 (rectifier). Such a propulsion system 1 is known by the English acronym USF for "Unducted Single Fan."In the case of a USF-type propulsion system 1, the blades 17 of the stator 16 are fixed in rotation relative to the X-axis of rotation of the upstream fan rotor 9 and consequently do not experience centrifugal force. The blades 17 of the stator 16 also have variable pitch.

[0037] Removing the fairing around the fan section 2 significantly increases the bypass ratio without the propulsion system 1 being negatively impacted by the mass of the housings or nacelles intended to surround the fan section 2. The bypass ratio of the propulsion system 1, including an unfaired fan section 2, is thus greater than or equal to 40, for example, between 40 and 80 inclusive. Furthermore, the peripheral velocity at the tip 21 of the fan blades 14 of the fan rotor(s) 9 can be between 210 m / s and 260 m / s. The fan pressure ratio can then preferably be between 1.05 and 1.20.

[0038] The reduction mechanism 19 may include an epi-cycloidal or planetary reduction mechanism 19, single-stage or two-stage. For example, the reduction mechanism 19 may be of the planetary type (“star” in English) ([Fig.3]) and include a sun pinion 19a (input of the reduction mechanism 19), centered on an X axis of rotation of the reduction mechanism 19 (generally coincident with the longitudinal axis X) and configured to be driven in rotation by the low pressure shaft 11, a ring 19b (output of the reduction mechanism 19) coaxial with the sun pinion 19a and configured to drive in rotation the blower shaft 20 around the X axis of rotation, and a series of satellites 19c distributed circumferentially around the X axis of rotation between the sun pinion 19a and the ring 19b, each satellite 19c being internally meshed with the sun pinion 19a and externally with the ring 19b.The series of satellites 19c is mounted on a satellite carrier 19d which is fixed relative to a stator part 19e of the propulsion system 1, for example relative to a housing of the compressor section 4, 5. Alternatively, the reduction mechanism 19 can be epicyclic (“planetary” in English) ([Fig.4]), in which case the ring 19b is fixedly mounted on the stator part 19e of the propulsion system 1 and the shaft of . blower 20 is driven in rotation by the satellite carrier 19d.

[0039] Regardless of the configuration of the reduction mechanism 19, the diameter of the ring gear 19b and the planet carrier 19d are greater than the diameter of the solar pinion 19a, so that the rotational speed of the blower rotor 9 is less than the rotational speed of the low pressure shaft 11.

[0040] 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 can 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 can be between 9.0 and 11.0.

[0041] The redline speed of the low-pressure shaft 11, which corresponds to the absolute maximum speed that the low-pressure shaft 11 is likely to encounter during the entire flight (according to European certification regulation EASA CS-E 740 (or according to US certification regulation 14-CFR Part 33.87)), is between 8,500 rpm and 12,000 rpm, preferably between 9,000 rpm and 11,000 rpm. The redline speed corresponds to the maximum rotational speed when the propulsion system is functioning correctly. It is therefore likely to be reached by the low-pressure shaft 11 under flight conditions. This redline speed is part of the data declared in the engine certification (type certification data sheet).Indeed, this rotational speed is usually used as a reference speed for the sizing of propulsion systems 1 and in certain certification tests (such as blade loss or rotor integrity tests).

[0042] In order to optimize the performance of the propulsion system 1 in terms of specific consumption, mass and drag, while ensuring 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 3.3 x 104 and less than or equal to 20.0 x 104 N / m2 where the thrust density per blade 14 is defined by the following formula: Thrust density - *100 and where: FN is the thrust generated by the blower rotor 9 and is expressed in Newtons (N);

[0043] n is the number of blades 14 in the blower rotor 9; and

[0044] D is the diameter of the blower rotor 9, measured in a plane normal to the X axis of rotation at the intersection between a vertex 21 and a leading edge 22 of the blades 14 of the blower rotor 9, and is expressed in meters (m). Note that the figures Since 1 and 2 are partial views, diameter D is only partially visible.

[0045] When the propulsion system 1 includes two blower rotors 9, the thrust density per blade 14 of the blower rotor is less than or equal to 4.0 x 104 N / m2.

[0046] Indeed, the Applicant observed that, when the thrust density is less than 3.3 x 10⁴ N / m², it was difficult to integrate propulsion system 1 because it was too bulky, had too high a mass, and generated excessive drag. Furthermore, when the thrust density is greater than 20.0 x 10⁴ N / m², the performance of propulsion system 1 in terms of specific fuel consumption is degraded. Therefore, sizing propulsion system 1 so that the thrust density per blade 14 of the fan rotor 9 is between 3.3 x 10⁴ and 20.0 x 10⁴ N / m² makes it possible to obtain a compromise between integration and the performance of propulsion system 1 when propulsion system 1 includes a reduction mechanism 19 and has a high bypass ratio.Such a blade thrust density range 14 is also compatible with a blower pressure ratio of less than 1.45, which allows the propulsive efficiency of the propulsion system 1 to be optimized.

[0047] By way of example, a propulsion system 1 according to the invention comprising a shrouded fan rotor and having a thrust density per fan blade 14 of 11 x 10⁴ N / m² has a specific consumption 15% lower than the same propulsion system having a thrust density per fan blade of 21 x 10⁴ N / m². The propulsion system 1 can be sized to obtain a thrust density per blade 14 between 3.3 x 10⁴ and 20.0 x 10⁴ N / m² by first fixing the thrust (FN) generated by the fan section 2 and 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 blower pressure rate 2 can be reduced.The number of fan blades 14 (n) and the rotational speed of the fan rotor 9 can also be adapted to meet performance, acoustic, and integration requirements. Depending on the aerodynamic characteristics of the fan section 2, the propulsion system 1 can be modified to incorporate a pitch-change mechanism 15, 15a, allowing adjustment of the pitch of the rotor blades 14 (and possibly the stator blades 16) of the fan section 2. Furthermore, depending on the integrated performance balance (fuel consumption balance of the propulsion system 1 integrated into the aircraft (mass, specific fuel consumption, drag)) and aircraft constraints (in terms of integration and program constraints), the fan section 2 can be ducted or unducted (and include a single fan rotor 9 or two counter-rotating fan rotors 9). Finally, the thermodynamic cycle is suitable for . different parameters thus dimensioned (blower diameter, number of blades, pressure rate of the blower 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.

[0048] For thrust densities per blade 14 between 3.3 x 10⁴ and 20.0 x 10⁴ N / m², 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 on 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 blower 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).

[0049] The fan rotor 9 also comprises at least 12 blades 14 and at most 24 blades 14, preferably at least 16 blades 14 and at most 22 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.

[0050] In order to further improve the propulsive efficiency of the propulsion system 1, the power density per blade 14 of the blower rotor 9 is greater than or equal to 3.65 x 106 and less than or equal to 22.0 x 106 W / m2, where the power density per blade 14 of the blower rotor 9 is defined by the following formula:

[0051] Power density = Z™— *100

[0052] where the power of the blower corresponds to the power of the blower rotor 9 and is expressed in Watts (W).

[0053] In the case of a propulsion system 1 comprising two blower rotors 9, the power density per blade 14 of the blower rotor is preferably less than or equal to 5.0 x 106 W / m2.

[0054] To further improve the efficiency of the propulsion system 1 and facilitate the integration of the reduction mechanism 19 into the propulsion system 1, the low-pressure compressor 4 has a power density greater than or equal to 0.64 and less than or equal to 0.97, where the power density of the low-pressure compressor 4 is defined by the following formula: Compressor power density = and where: net is the number of compressor stages;

[0055] Um is the average peripheral speed of the rotor of the low-pressure compressor 4 and is expressed in meters per second (m / s); and

[0056] Rext_moy is the average external radius of the low-pressure compressor 4 and is expressed in meters (m).

[0057] The average external radius of the compressor is equal to the arithmetic mean of the external radii RI of the rotors (moving blade wheels) of the low pressure compressor 4. In a given stage, the external radius RI of the rotor corresponds to the distance between the top of the moving blades of the rotor 4a and the X axis of rotation, at mid-distance (at the top) between the leading edge and the trailing edge of the moving blades of the rotor 9a (50% of the chord at the top).

[0058] Since the rotational speed of the blower rotor 9 is decoupled from the rotational speed of the low-pressure shaft 11 by means of the reduction mechanism 19, the rotational speed of the low-pressure shaft 11 can be increased so that the low-pressure turbine 8, and therefore the low-pressure compressor 4, rotates at an optimal speed. This implies an increase in the power density of the low-pressure compressor 4. However, the present application proposes to use the increased rotational speed of the low-pressure shaft 11 to reduce the average external radius Rext_moy of the low-pressure compressor 4.This reduction of the average external radius Rext_moy of the low pressure compressor 4 allows, where appropriate, for an increase in the overall compression ratio of the propulsion system 1, since it depends on the flow rate of the primary airflow FL. Thus, the overall compression ratio of the propulsion system 1, 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 blower rotor 9 (measured at the base of the blower rotor 9), can 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.

[0059] When the integration of the reduction mechanism 19 allows it, the reduction of the average external radius Rext_moy of the low pressure compressor 4 also allows the length and downward curvature (gooseneck) of the conduit 23 separating the low pressure compressor 4 and the high pressure compressor 5 to be reduced. The transition between the low pressure compressor 4 and the high pressure compressor 5 is therefore improved, which improves the efficiency of the compressor section 4, 5.

[0060] The fan rotor 9 also has a hub-to-head ratio of between 0.22 and 0.32. In the case of a fan rotor with fixed pitch, the hub-to-head ratio can be between 0.24 and 0.32. In the case of a fan rotor with variable pitch, the hub-to-head ratio is preferably between 0.24 and 0.32 to allow for 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 Rei 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 that radially delimits the inlet flow channel of the fan rotor 9 (and corresponds to the point of connection of the leading edge 22 with the aerodynamic surface of a (fan rotor platform 9). The external radius Re corresponds 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. However, decreasing the hub-to-head ratio of the fan rotor 9 implies an increase in the mechanical load on the hub 13 of the fan rotor 9. Dimensioning 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 within the ranges defined above.

[0061] A twin-body propulsion system 1 having a thrust density and a power density per blade 14 of the blower rotor 9 and a power density of the low-pressure compressor 4 in the ranges defined above may in particular include a two-stage high-pressure turbine 8, a high-pressure compressor 5 comprising at least eight stages and at most eleven stages, a low-pressure turbine 7 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.

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 dilution ratio of the propulsion system (1) being greater than or equal to 40; 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, being between 210 m / s and 260 m / s; and a thrust density per blade (14) of the fan rotor (9) of the propulsion system (1) being greater than or equal to 3.3 x 104 and less than or equal to 20.0 x 104 N / m2 where the thrust density per blade (14) is defined by the following formula: Thrust density = *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. A 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 106 and less than or equal to 22.0 x 106 W / m2, wherein the power density per blade (14) of the fan rotor (9) is defined by the following formula: . , - . blower power Power density = -------—----*100 and where the fan power corresponds to the fan rotor power (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.

30.

4. Propulsion system (1) according to one of claims 1 to 3, in which the diameter of the fan rotor (9) is between 80 inches (203.2cm) and 185 inches (469.9cm) 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).

5. Propulsion system (1) according to one of claims 1 to 4, wherein a dilution ratio of the propulsion system (1) is between 40 and 80 inclusive.

6. Propulsion system (1) according to one of claims 1 to 5, in which a hub-to-head ratio of the fan rotor (9) is between 0.22 and 0.

32.

7. Propulsion system (1) according to one of claims 1 to 6, further comprising a drive turbine (8) and a compressor (4) directly connected by the drive shaft (11), the drive turbine (8) comprising at least 3 and at most 5 stages.

8. Propulsion system (1) according to claim 7, wherein the compressor (4) comprises at least 2 and at most 4 stages.

9. Propulsion system (1) according to one of claims 1 to 8, 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.

10. Propulsion system (1) according to claim 9, wherein the high pressure compressor (5) comprises at least 8 and at most 11 stages.

11. Aircraft (100) comprising at least one propulsion system (1) according to one of claims 1 to 10 fixed to the aircraft via a mast.

12. A method for 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 bypass ratio of the propulsion system (1) is greater than or equal to 40; a peripheral speed at the tip (21) of the blades (14) of the fan rotor (9), when the propulsion system (1) is stationary in takeoff mode in a standard atmosphere and at sea level, is between 210 m / s and 260 m / s;and a thrust density per blade (14) of the fan rotor (9) of the propulsion system (1) is greater than or equal to 3.3 x 104 and less than or equal to 20.0 x 104 N / m2 where the thrust density per blade (14) of the fan rotor (9) is defined by the following formula: Thrust density = * 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).

13. A sizing method according to claim 12, 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 106 and less than or equal to 22.0 x 106 W / m2, wherein the power density per blade (14) of the fan rotor (9) is defined by the following formula: . j. , . power of the except flante e-,n Power density = ----------- *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).

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