Aeronautical propulsion system comprising an optimized fan section

Abstract

The present disclosure relates to a fan section having a performance coefficient of between 1.05 and 1.3, the performance coefficient being defined as follows: (Cp), where: Cp is the performance coefficient; FPR is the pressure ratio of the fan section (2); Cs is the strength of the fan rotor (9); and U is the peripheral speed at the blade tip of the fan rotor (9); the pressure ratio and the peripheral speed being measured at cruising speed.

Description

[0001] DESCRIPTION

[0002] Aircraft propulsion system including an optimized fan section

[0003] TECHNICAL FIELD

[0004] This application generally concerns the field of propulsion systems, and more particularly aeronautical propulsion systems having a high, or even very high, dilution 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] Improving the system's propulsive efficiency can also be achieved by sizing the fan section. Indeed, due to its large diameter (in particular to achieve high bypass ratios and low fan pressure ratios), the fan section represents a significant part of the propulsion system in terms of mass and therefore specific consumption. At the same time, the fan section produces a very large part of the propulsion system's thrust. EXPOSE

[0010] One aim of this application is to optimize the fan section of the propulsion system in order to make it more efficient without excessively penalizing the mass and therefore the specific consumption of the propulsion system.

[0011] To this end, according to a first aspect, there is provided a fan section of an aeronautical propulsion system, the fan section comprising a fan rotor having a coefficient of performance greater than or equal to 1.05 and less than or equal to 1.3, where the coefficient of performance is defined as follows: i

[0012] FPR3.5

[0013] Cp = — - x 10

[0014] Cs 5 where: Cp is the coefficient of performance and is expressed in seconds per meter squared (s / m) 2 ;

[0015] FPR is the blower section pressure ratio;

[0016] Cs is the strength of the fan rotor and ; and

[0017] U is the peripheral speed at the fan rotor blade tip and is expressed in meters per second (m / s); the pressure ratio and peripheral speed are measured at cruising speed.

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

[0019] - the blower pressure ratio is greater than or equal to 1.1 and less than or equal to 1.45;

[0020] - the solidity is greater than or equal to 0.9 and less than or equal to 1.3, preferably greater than or equal to 0.9 and less than or equal to 1.1;

[0021] - the peripheral speed is greater than or equal to 260 m / s and less than or equal to 400 m / s, for example greater than or equal to 270 m / s and less than or equal to 350 m / s;

[0022] - an elongation of the fan blades, which corresponds to a ratio between a height of the fan blades and a chord at the top of the fan blades, is greater than or equal to 2.4 and less than or equal to 3.0;

[0023] - a fan rotor diameter is greater than or equal to 177.8 cm and less than or equal to 304.8 cm, for example less than or equal to 304.8 cm, for example of the order of 228.6 cm; and / or

[0024] - the fan section has a hub-to-head ratio greater than or equal to 0.22 and less than or equal to 0.32, for example greater than or equal to 0.235 and less than or equal to 0.30, for example still less than or equal to 0.27.

[0025] According to a second aspect, an aeronautical propulsion system is proposed comprising:

[0026] - a drive shaft movable in rotation around an axis of rotation;

[0027] - a blower shaft;

[0028] - a fan section according to the first aspect, the fan rotor being rotated by the fan shaft; and - a reduction mechanism coupling the drive shaft and the fan shaft to drive the fan shaft at a rotational speed lower than the rotational speed of the drive shaft.

[0029] Optionally, the aeronautical propulsion system according to the second aspect may further comprise a fan casing surrounding the fan blades.

[0030] The dilution ratio of the propulsion system according to the second aspect may further be greater than or equal to 10, for example between 10 and 35 inclusive, for example between 10 and 18 inclusive.

[0031] According to a third aspect, there is provided an aircraft comprising at least one propulsion system according to the second aspect to the aircraft via a mast.

[0032] According to a fourth aspect, there is provided a method for dimensioning a fan section of an aeronautical propulsion system comprising a step of dimensioning a pressure ratio of the fan section, a solidity of a rotor of the fan section and a peripheral speed of the rotor such that a coefficient of performance of the fan section is greater than or equal to 1.05 and less than or equal to 1.3, where the coefficient of performance is defined as follows: i FPR3.5

[0033] Cp = — - x 10

[0034] Cs 5 where: Cp is the coefficient of performance; FPR is the fan section pressure ratio; Cs is the fan rotor strength; and U is the fan rotor blade tip peripheral speed; the pressure ratio and peripheral speed being measured at cruising speed.

[0035] Optionally, the sizing method according to the fourth aspect further comprises a step of sizing the fan blades such that their aspect ratio is greater than or equal to 2.4 and less than or equal to 3.0, where the aspect ratio corresponds to a ratio between a height of the fan blades and a chord at the top of the fan blades.

[0036] According to a fifth aspect, there is provided a method of manufacturing a fan section comprising the following steps:

[0037] - dimensioning a blower section in accordance with the dimensioning method according to the fourth aspect; and

[0038] - manufacture the blower section thus dimensioned.

[0039] DESCRIPTION OF FIGURES

[0040] Other characteristics, aims and advantages of the present disclosure will emerge from the following description, which is purely illustrative and non-limiting, and which must be read in conjunction with the appended drawings in which: Figure 1 is a schematic, partial and sectional view of an example of a propulsion system according to a first embodiment;

[0041] Figure 2a is a partial and schematic sectional view of an example of a fan rotor of a propulsion system according to one embodiment, the section being made in a plane passing through an upstream point of intersection between a tip and a leading edge of two adjacent blades;

[0042] Figure 2b is a perspective view of an exemplary fan rotor of a propulsion system according to one embodiment;

[0043] Figure 3 is a schematic sectional view of an example of a reduction mechanism according to a first variant;

[0044] Figure 4 is a schematic sectional view of an example of an epicyclic reduction mechanism according to a second variant;

[0045] Figure 5 is an example of an aircraft that may include at least one propulsion system according to one embodiment;

[0046] Figure 6 is a flowchart illustrating exemplary steps of a sizing or manufacturing method according to one embodiment.

[0047] Throughout the figures, similar elements have identical references.

[0048] DETAILED DESCRIPTION

[0049] 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).

[0050] The compressor section 4, 5 comprises a succession of stages each comprising a moving blade wheel (rotor) 4a, 5a rotating in front of a fixed blade wheel (stator) 4b, 5b. The turbine section 7, 8 also comprises a succession of stages each comprising a fixed blade wheel (stator) 7b, 8b behind which a moving blade wheel (rotor) 7a, 8a rotates.

[0051] 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.

[0052] 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. The secondary air flow F2 (also called “bypass air flow”) flows around the primary body 3. The secondary air flow F2 makes it possible to cool the periphery of the primary body 3 and is used to generate the majority of the thrust provided by the propulsion system 1.

[0053] 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.

[0054] In a twin-body 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 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.

[0055] 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.

[0056] The fan section 2 comprises at least the fan rotor 9 capable of being 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 13.

[0057] The fan section 2 may further comprise a fan stator 16, or rectifier, which comprises vanes 17 mounted on a hub of the fan stator 16 and whose function is to straighten the secondary air flow F2 which flows out of the fan rotor 9. The vanes 17 of the fan stator 18 may be fixed relative to the hub or have a variable pitch.

[0058] 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 e edition) and at sea level. By "not installed" we mean 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.

[0059] It will be noted that, in the present application, certain parameters are determined in cruising conditions, i.e. at 10668 m altitude (35000 feet), 0.8 Mach and in ISA conditions (International Standard Atmosphere) defined by the ISO2533 / 1975 edition / 1985 addendum. In addition, the distances (length, radius, diameter, chord, etc.) are measured at ambient temperature (approximately 20°C) when the propulsion system 1 is cold, i.e. when the propulsion system 1 has been stopped for a sufficient period for the parts of the propulsion system to be at ambient temperature, it being understood that these dimensions vary little compared to the conditions in which the propulsion system 1 is in takeoff mode.

[0060] 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.

[0061] 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.

[0062] The propulsion system 1 is configured to provide thrust of between 18,000 Ibf (80,068 N) and 51,000 Ibf (222,411 N), for example between 20,000 Ibf (88,964 N) and 35,000 Ibf (155,688 N), 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.

[0063] The fan section 2 may comprise a fan housing 12, the fan rotor 9 being housed in the fan housing 12.

[0064] The fan rotor 9 extends 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 for example greater than or equal to 10, for example between 10 and 35 inclusive, for example between 10 and 18 inclusive.

[0065] Each fan blade 14 has a leading edge 14a and a trailing edge 14b (see for example figures 2a and 2b). The leading edge 14a is configured to extend opposite the flow of gases entering the fan rotor 9. It corresponds to the front part of an aerodynamic profile which faces the air flow and which divides the air flow into a pressure flow and an extrados flow. The trailing edge 28b corresponds to the rear part of the aerodynamic profile, where the pressure and extrados flows meet.It should be noted here that, when the blades 14 comprise a leading edge and / or trailing edge shield, the leading edge 14a (respectively the trailing edge 14b) of the blades 14 corresponds to the front part of the shield profile which reconstitutes the leading edge (respectively the rear part of the shield profile which reconstitutes the trailing edge 14b) and whose function is to divide the flow into an intrados flow and an extrados flow (respectively to join the flows).

[0066] The fan rotor 9 further comprises a series of platforms each extending between two adjacent blades 14 and configured to radially delimit inside the air flow F passing through the rotor 9.

[0067] The fan blade 14 further has a blade tip chord ci. The blade tip chord ci corresponds to the straight line segment that connects an upstream intersection point P between the leading edge 14a and the tip 21 of a blade 14 and a downstream intersection point between the trailing edge 14b and the tip 21 of the blade 14.

[0068] The fan rotor 9 has between sixteen and twenty-two blades 14.

[0069] The reduction mechanism 19 may comprise a reduction mechanism 19 with an epicyclic gear train, for example a reduction mechanism of the “epicyclic” or “planetary” type according to the terminology sometimes encountered by those skilled in the art, 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 may be of the epicyclic type (“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 19. e of the propulsion system 1, for example in relation to a casing of the compressor section 4, 5).

[0070] 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.

[0071] The reduction ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11, for example greater than or equal to 2.7 and less than or equal to 6.0, for example around 3.0.

[0072] The dual-body propulsion system 1 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.

[0073] 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, for example between 9000 rpm and 11000 rpm, 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. The limiting speed corresponds to the maximum rotation speed when the propulsion system 1 is healthy. 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 certification data sheet" in English). Indeed, this rotation speed is usually used as a reference speed for the dimensioning and manufacturing of propulsion systems 1 and in certain certification tests (such as blade loss or rotor integrity tests, typically CS-E-800 certification - bird strike and ingestion). In order to optimize the performance of the propulsion system 1 , a coefficient of performance of the fan section is between 1.05 and 1.3, where the coefficient of performance is defined as follows: i FPR3.5

[0074] Cp = — - x 10

[0075] Cs 5where: Cp is the coefficient of performance and is expressed in seconds per meter squared (s / m) 2 ;

[0076] FPR is the pressure ratio of fan section 2 (dimensionless);

[0077] Cs is the strength of the fan rotor 9 (dimensionless); and

[0078] U is the peripheral speed at the tip of the fan rotor blade and is expressed in meters per second (m / s).

[0079] Note that the pressure ratio FPR of the fan section 2 and the peripheral speed U are measured here in cruising mode since this is the flight phase in which we wish to obtain the maximum efficiency of the fan rotor 9.

[0080] The solidity Cs is equal to the ratio between the chord at the blade tip ci and an inter-blade pitch 23. The inter-blade pitch 23 corresponds to the angular distance between the upstream intersection points P of two adjacent blades 14; the inter-blade pitch 23 is therefore equal to the external radius R eof the fan rotor 9 (half-diameter) multiplied by the angle between a first straight line Di, included in a plane normal to the X axis which comes from the upstream intersection point P (see figures 2a and 2b) of a first blade 14 and intersects the X axis, and a second straight line D2, included in the plane normal to the X axis which comes from the upstream intersection point P of a second blade 14 immediately adjacent to the first blade 14 and intersects the X axis (see figure 2b). Since the solidity is a ratio of distances, it is measured when the propulsion system 1 is cold (under the aforementioned conditions). For example, the solidity Cs may be greater than or equal to 0.9 and less than or equal to 1.3, preferably greater than or equal to 0.9 and less than or equal to 1.1.

[0081] The pressure ratio FPR of the fan section 2 corresponds to the ratio between the average pressure at the outlet of the fan stator 17 and the average pressure at the inlet of the fan rotor 9. For example, the pressure ratio is greater than or equal to 1.1 and less than or equal to 1.45. The average pressures are measured here on the flow vein (from the surface which radially delimits the flow vein at the inlet of the fan rotor 9 to the fan casing 12.

[0082] The peripheral speed U is greater than or equal to 260 m / s and less than or equal to 400 m / s, for example greater than or equal to 270 m / s and less than or equal to 350 m / s.

[0083] The sizing of the fan section 2 so that its coefficient of performance Cp is between 1.05 and 1.3 makes it possible to maximize the performance of the fan section 2 in cruising mode while maintaining acceptable operability (essentially at takeoff). The coefficient of performance Cp in fact takes into account both the aerodynamic work that the fan rotor 9 must perform (via its pressure ratio FPR) and the physical material of the blades 14 which performs this work (via the solidity Cs). The higher the coefficient of performance, the more the fan rotor 9 is capable of providing significant aerodynamic work with blades 14 having a small footprint. The efficiency of the fan rotor 9 is therefore improved with reduced footprint and mass.The coefficient of performance Cp is preferably less than or equal to 1.3 in order to avoid any risk of pumping and to preserve sufficient mechanical strength in the event of collision with a bird and ingestion.

[0084] For example, the fan section 2 may have a pressure ratio FPR of 1.35, a strength of 0.9, a peripheral speed of 305 m / s, and a coefficient of performance of 1.3. In another example, the fan section 2 may have a pressure ratio FPR of 1.38, a strength of 1.05, a peripheral speed of 315 m / s, and a coefficient of performance of 1.05. In yet another example, the fan section 2 may have a pressure ratio FPR of 1.3, a strength of 0.95, a peripheral speed of 295 m / s, and a coefficient of performance of 1.3.

[0085] Preferably, the blades 14 of the fan rotor 9 also have an aspect ratio greater than or equal to 2.4 and less than or equal to 3.0. The aspect ratio corresponds to the ratio between the height h of the fan blades and their chord at the tip ci. When the fan section has a coefficient of performance of between 1.05 and 1.3 with blades 14 whose aspect ratio is between 2.4 and 3.0, it is possible to reduce the thickness of the fan blades 14 while complying with the requirements of current certifications regarding ingestion (typically the CS-E-800 certification - collision with a bird and ingestion). The fan blades 14 are then more "tapered", which further reduces the mass of the fan section 2 and therefore the specific consumption of the propulsion system 1, while guaranteeing the required aerodynamic work and avoiding the risks of surge.The overall length of the fan nacelle can also be reduced.

[0086] The height h of a blade 14 is measured here in a plane normal to the X axis, between the upstream intersection point P (between the tip 21 and the leading edge 14a of the blades 14 of the fan rotor 9) and a point of intersection between the blade and the platform, in the plane normal to the axis (figure 1).

[0087] The diameter D of the fan rotor can then be between 70 inches (177.8 cm) and 185 inches (469.9 cm) inclusive. The diameter D is for example between 70 inches (177.8 cm) and 120 inches (304.8 cm) inclusive, for example of the order of 90 inches (228.6 cm), which makes it possible to integrate the propulsion system 1 in a conventional manner, in particular under the wing of an aircraft 100. The diameter of the fan rotor 9 is measured here in a plane normal to the axis X of rotation at the upstream point of intersection P (between the tip 21 and the leading edge 14a of the blades 14 of the fan rotor 9). Note that since FIG. 1 is a partial view, the diameter D is only partially visible.

[0088] 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 of the fan rotor 9. In one embodiment, the number of blades 16 in the fan stator 17 is at least equal to 30 and at most equal to 52, for example exactly equal to 48 when the fan rotor comprises twenty-two blades 14.

[0089] The fan rotor 9 also has a hub-to-head ratio of between 0.22 and 0.32, for example. In the case of a fixed-pitch fan rotor 9, the hub-to-head ratio may be between 0.22 and 0.30, for example between 0.235 and 0.27. The hub-to-head ratio corresponds to the ratio between the internal radius Ri and the external radius R eof 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 path 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 the platform of the fan rotor 9). The external radius R e is equal 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.

[0090] Examples:

[0091] Example 1

[0092] A first example of an engine whose fan section 2 has a coefficient of performance between 1.05 and 1.3 comprises 22 fan blades 14, a peripheral speed U in cruising mode equal to 315 m / s, a pressure rate FPR in cruising mode equal to 1.4, a solidity equal to 1.0. The coefficient of performance of this fan section 2 is then equal to 1.1 (s / m) 2. This fan section 2 also comprises 48 stator blades 16. The fan blades 14 have a chord at the tip ci of 32.6 cm, a height of 84.6 cm and an aspect ratio of 2.8. The fan rotor 9 also has a diameter of 229 cm and a hub-to-head ratio equal to 0.26. The fan blades are made of a composite material comprising an embedded fibrous reinforcement (comprising carbon fibers having a Young's modulus greater than 250 GPa) in an epoxy matrix. The fan section is shrouded and the fan blades 14 are variable pitch.

[0093] Propulsion System 1 is configured to provide 164 kN of thrust 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, 3 e edition) and at sea level.

[0094] Such a fan section 2 can be integrated into a propulsion system 1 having a bypass ratio of 14 in cruising mode and comprising a reduction mechanism 19 with a reduction ratio of 3.2. The propulsion system 1 can further comprise two low-pressure compressor stages 4, ten high-pressure compressor stages 5, two high-pressure turbine stages 7 and five low-pressure turbine stages 8.

[0095] Compared to a fan section with a coefficient of performance less than 1.05 (s / m) 2 , for example equal to 0.7, the fan section of the example engine detailed above is capable of achieving a better level of efficiency with less mass and size while maintaining an acceptable level of operability with regard to certification.

[0096] Example 2 Engine 1 is a twin-body propulsion system comprising a shrouded fan section 2 corresponding to the current technical standard (at the date of filing of this application) which is sought to be improved.

[0097] The engine 2 is a double-body propulsion system 1 comprising a shrouded fan section 2 in accordance with the teaching of the present application having a coefficient of performance of between 1.05 and 1.3.

[0098] As in the first example, the fan section of engine 1 has a coefficient of performance less than 1.05 (s / m) 2, while the coefficient of performance of engine 2 is equal to 1.11 and therefore between 1.05 and 1.3, at equivalent thrust. Consequently, engine 2 is capable of achieving at least an equivalent level of efficiency, with less mass, drag and size compared to engine 1, while being easier to install on aircraft and maintaining an acceptable level of operability with regard to certification.

Claims

CLAIMS 1. Fan section (2) of an aeronautical propulsion system, the fan section (2) comprising a fan rotor (9) having a coefficient of performance greater than or equal to 1.05 and less than or equal to 1.3, where the coefficient of performance is defined as follows: i FPR3.5 Cp = — - x 10 Cs 5 where: Cp is the coefficient of performance and is expressed in seconds per meter squared (s / m) 2 ; FPR is the pressure ratio of the fan section (2); Cs is the strength of the fan rotor (9) and ; and U is the peripheral speed at the tip of the fan rotor blade (9) and is expressed in meters per second (m / s); the pressure ratio and the peripheral speed being measured in cruising mode.

2. Blower section (2) according to claim 1, wherein the pressure ratio of the blower is greater than or equal to 1.1 and less than or equal to 1.

45.

3. Blower section (2) according to one of claims 1 and 2, wherein the solidity is greater than or equal to 0.9 and less than or equal to 1.3, preferably greater than or equal to 0.9 and less than or equal to 1.

1.

4. Fan section (2) according to one of claims 1 to 3, in which the peripheral speed is greater than or equal to 260 m / s and less than or equal to 400 m / s, for example greater than or equal to 270 m / s and less than or equal to 350 m / s.

5. Fan section (2) according to one of claims 1 to 4, wherein an elongation of the fan blades (14), which corresponds to a ratio between a height of the fan blades and a top chord (ci) of the fan blades, is greater than or equal to 2.4 and less than or equal to 3.

0.

6. Fan section (2) according to one of claims 1 to 5, wherein a diameter of the fan rotor (9) is greater than or equal to 177.8 cm and less than or equal to 304.8 cm, for example less than or equal to 304.8 cm, for example of the order of 228.6 cm.

7. Fan section (2) according to one of claims 1 to 6 having a hub-head ratio greater than or equal to 0.22 and less than or equal to 0.32, for example greater than or equal to 0.235 and less than or equal to 0.30, for example still less than or equal to 0.

27.

8. 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) according to one of claims 1 to 7, the fan rotor (9) being driven in rotation by the fan shaft (20); and - 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).

9. Aeronautical propulsion system (1) according to claim 8, further comprising a fan casing surrounding the fan blades (14).

10. Propulsion system (1) according to one of claims 8 and 9, in which a dilution ratio of the propulsion system (1) is greater than or equal to 10, for example between 10 and 35 inclusive, for example between 10 and 18 inclusive.

11. Aircraft (100) comprising at least one propulsion system (1) according to one of claims 8 to 10 fixed to the aircraft via a mast.

12. A method of dimensioning a fan section (2) of an aeronautical propulsion system comprising a step of dimensioning a pressure ratio of the fan section, a solidity of a rotor (9) of the fan section and a peripheral speed of the rotor (9) such that a coefficient of performance of the fan section (2) is greater than or equal to 1.05 and less than or equal to 1.3, where the coefficient of performance is defined as follows: i FPR3.5 Cp = — - x 10 Cs 5 where: Cp is the coefficient of performance; FPR is the pressure ratio of the fan section (2); Cs is the strength of the fan rotor (9); and U is the peripheral speed at the tip of the fan rotor blade (9); the pressure ratio and the peripheral speed being measured in cruising mode.

13. A sizing method according to claim 12, further comprising a step of sizing the fan blades (14) such that their aspect ratio is greater than or equal to 2.4 and less than or equal to 3.0, where the aspect ratio corresponds to a ratio between a height of the fan blades (14) and a top chord (ci) of the fan blades.

14. A method of manufacturing a fan section comprising the following steps: dimensioning a fan section in accordance with the dimensioning method according to one of claims 12 and 13; and manufacturing the fan section thus dimensioned.

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