Aeronautical propulsion system having improved propulsion efficiency

Abstract

An aeronautical propulsion system (1) comprises an open fan rotor (9) having an aerodynamic load parameter kaero greater than or equal to 0.07 and less than or equal to 0.12 and a mechanical load parameter kmeca greater than or equal to 11.5 × 106 tr.m2.min-1 and less than or equal to 20.0 × 106 tr.m2.min-1, where the aerodynamic load kaero and the mechanical load kmecaare defined by the following formulae: formula (1), and where: FN is the thrust of the propulsion system (1); BPR is the dilution rate; F = 39.7; Re and Ri are the outer and inner radii of the fan rotor; ω is the limit speed of the fan rotor; and S is the surface of the jet flow at the inlet of the fan rotor (9).

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 BPR (bypass ratio in English, 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] This decoupling allows the rotational speed of the fan rotor to be reduced. Furthermore, in order to reduce the pressure ratio of the fan rotor and improve the bypass ratio of the propulsion system, it has been proposed to increase the diameter of the fan rotor (and thus reduce the peripheral speed of the fan blades). The consequence is a reduction in the centrifugal forces in the fan rotor. However, increasing the fan diameter has a negative impact on the mass and drag of the fan section, and by extension on the propulsion system, and increases the torque to be applied by the fan shaft to the fan rotor. EXPOSE

[0010] An aim of the present application is to optimize the performance of the propulsion system, while maintaining acceptable mechanical and aerodynamic loading for the fan section of the propulsion system.

[0011] For this purpose, according to a first aspect, an aeronautical propulsion system is proposed comprising:

[0012] - a drive shaft movable in rotation around an axis of rotation;

[0013] - a fan shaft;

[0014] - a fan section comprising an unducted fan rotor rotated by the fan shaft, the fan rotor comprising a plurality of blades, each blade;

[0015] - a reduction mechanism coupling the drive shaft and the fan shaft in order to drive the fan shaft at a rotational speed lower than the rotational speed of the drive shaft; the propulsion system being configured so that the fan rotor has an aerodynamic load parameter greater than or equal to 0.07 and less than or equal to 0.12 and a mechanical load parameter greater than or equal to 11.5 x 10® rpm 2 .min' 1and less than or equal to 20.0 x 10® tr.m 2 .min' 1 , where the aerodynamic load and the mechanical load are defined by the following formulas: and where . kaero is the aerodynamic load parameter; kmeca is the mechanical load parameter;

[0016] FN is the thrust generated by the propulsion system when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level and is expressed in Newton;

[0017] BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is stationary in takeoff regime in a standard atmosphere and at sea level; n is the number of blades in the fan rotor;

[0018] F = 39.7 N _1 ' 35 .(rpm)' 2 .rrr 5

[0019] R eis the outer radius of the fan rotor and corresponds to a distance between the axis of rotation and a point of intersection between a tip and a leading edge of the fan rotor blades, and is expressed in meters (m);

[0020] Ri is the inner radius of the fan rotor and corresponds to a distance between the axis of rotation and a point of intersection between the leading edge of the fan rotor blades and a surface of the fan rotor that radially delimits on the inside a flow path in the fan rotor and is expressed in meters (m); a is the limiting rotational speed of the fan rotor and is expressed in revolutions per minute (rpm); and S is the area of ​​the flow path at the inlet of the fan rotor, which is equal to rt x (R e 2 - Ri 2 ) and is expressed in square meters (m 2 ).

[0021] Some preferred but non-limiting features of the propulsion system according to the first aspect are the following, taken individually or in combination:

[0022] - the outer radius of the fan rotor is between 40 inches (101.6 cm) and 92.5 inches (233.7 cm) inclusive, for example between 60 inches (152.4 cm) and 85 inches (215.9 cm) inclusive, for example of the order of 478 inches (198.1 cm);

[0023] - a fan rotor hub-to-head ratio is between 0.22 and 0.32;

[0024] - the fan rotor comprises at least twelve fan blades and at most eighteen fan blades, for example at least fourteen fan blades and at most sixteen fan blades;

[0025] - a reduction rate of the reduction mechanism is greater than or equal to 5.0 and less than or equal to 11.0;

[0026] - a dilution ratio of the propulsion system is greater than or equal to 40, for example between 40 and 80 inclusive;

[0027] - 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;

[0028] - 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, for example less than or equal to 1.30;

[0029] - the propulsion system is sized so that a thrust of the propulsion system, the propulsion system is stationary in take-off mode in a standard atmosphere, is between 18,000 Ibf and 51,000 Ibf;

[0030] - the propulsion system further comprises a drive turbine and a compressor directly connected by the drive shaft, wherein the drive turbine comprises at least three and at most five stages;

[0031] - compressor comprises at least two and at most four stages;

[0032] - 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

[0033] - the high-pressure compressor comprises at least eight and at most eleven stages.

[0034] According to a second aspect, there is provided an aircraft comprising at least one propulsion system according to the first aspect attached to the aircraft by means of a mast.

[0035] According to a third aspect, there is provided a method of dimensioning or manufacturing a propulsion system comprising a reduction mechanism coupling a drive shaft and an unducted fan rotor to drive the fan rotor at a speed lower than a speed of the drive shaft, wherein the fan rotor has an aerodynamic load parameter greater than or equal to 0.07 and less than or equal to 0.12 and a mechanical load parameter greater than or equal to 11.5 x 10 e tr.m 2 .min' 1 and less than or equal to 20.0 x 10 e tr.m 2 .min' 1 , where the aerodynamic load and the mechanical load are defined by the following formulas: and where - kaero is the aerodynamic load parameter; kmeca is the mechanical load parameter;

[0036] FN is the thrust generated by the propulsion system when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level and is expressed in Newton (N);

[0037] BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is stationary in takeoff regime in a standard atmosphere and at sea level; n is the number of blades in the fan rotor;

[0038] F = 39.7 N _1 ' 35 .(rpm)' 2 .m -5

[0039] R e is the outer radius of the fan rotor and corresponds to a distance between the axis of rotation and a point of intersection between a tip and a leading edge of the fan rotor blades, and is expressed in meters (m);

[0040] Ri is the inner radius of the fan rotor and corresponds to a distance between the axis of rotation and a point of intersection between the leading edge of the fan rotor blades and a surface of the fan rotor that radially delimits on the inside a flow path in the fan rotor and is expressed in meters (m); co is the limiting rotational speed of the fan rotor and is expressed in revolutions per minute (rpm); and

[0041] S is the area of ​​the flow path at the inlet of the fan rotor, which is equal to it x (R e 2 - Ri 2 ) and is expressed in square meters (m 2 ).

[0042] According to a fourth aspect, a method of manufacturing an aeronautical propulsion system is proposed comprising the following steps:

[0043] - dimension the aeronautical propulsion system in accordance with the dimensioning method according to the third aspect; and

[0044] - manufacture the aeronautical propulsion system.

[0045] DESCRIPTION OF FIGURES

[0046] Other characteristics, aims and advantages of the invention 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:

[0047] Figure 1 is a schematic, partial and sectional view of an example of a propulsion system according to a first embodiment, in which the fan section is unducted; Figure 2 is a schematic sectional view of an example of a reduction mechanism according to a first variant embodiment;

[0048] Figure 3 is a schematic sectional view of an example of a reduction mechanism of a second variant embodiment;

[0049] Figure 4 is an example of an aircraft that may comprise at least one propulsion system in accordance with the first or second embodiment of the invention;

[0050] Figure 5 is a flowchart illustrating exemplary steps of a sizing or manufacturing method in accordance with one embodiment of the invention.

[0051] Throughout the figures, similar elements have identical references.

[0052] DETAILED DESCRIPTION

[0053] 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).

[0054] 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.

[0055] 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.

[0056] 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.

[0057] 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.

[0058] 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.

[0059] 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 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.

[0060] 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.

[0061] 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 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 have a variable pitch. 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 the pitch change mechanism 15.

[0062] 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 which 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 particular when the fan section 2 is unducted (see FIG. 2). In a similar manner to the rotor blades 14, the root of the variable-pitch stator blades 17 is pivotally mounted along 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.

[0063] 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 40 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 will 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 will be understood here that the measurements are carried out when the propulsion system 1 is in a test bench (and not installed on an aircraft 100), the measurements then being simpler to carry out. The distances (length, radius, diameter, etc.) are, on the other hand, measured at ambient temperature (approximately 20°C) when the propulsion system 1 is cold, that is to say when the propulsion system 1 has been stopped for a sufficient period for the parts of the propulsion system 1 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 take-off mode.

[0064] 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.

[0065] 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, for example 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).

[0066] Fan section 2 may be unducted.

[0067] The fan section 2 (which can also be referred to as the propeller) 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.

[0068] 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 be between 1.05 and 1.20, for example.

[0069] The reduction mechanism 19 may for example comprise a reduction mechanism 19 with an epicyclic gear train, for example of the “epicyclic” type or of the “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 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 may be of the “epicyclic” (“planetary” in English) type (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 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 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. 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).

[0072] The propulsion system 1 further has an overall compression ratio, which corresponds to the pressure ratio between the pressure at the outlet of the high-pressure compressor 5 and the pressure at the inlet of the fan rotor 9 (measured at the level of the foot of the fan rotor 9), greater than or equal to 40 and less than or equal to 70, for example greater than or equal to 44 and less than or equal to 55.

[0073] In order to optimize the performance of the propulsion system 1 in terms of specific consumption, mass and drag, while maintaining acceptable mechanical loading and aerodynamic loading for the fan section 2, the propulsion system 1 is configured so that the fan rotor 9 has an aerodynamic load parameter kaero greater than or equal to 0.07 and less than or equal to 0.12 and a mechanical load parameter kmeca greater than or equal to 11.5 x 10 6 tr 2 .m 2 .min' 2 and less than or equal to 20.0 x 10 6 tr 2 .m 2 .min -2 .

[0074] The aerodynamic load parameter kaero represents the aerodynamic torque that is applied to a fan blade 14 of the fan rotor 9, between the hub 13 and the tip 21 of the fan blade 14, and can be expressed by the following formula: where: FN is the thrust generated by the propulsion system and is expressed in Newton (N);

[0075] BPR is the bypass ratio of the propulsion system 1; n is the number of blades 14 in the fan rotor 9; co is the limit rotational speed (redline speed) of the fan rotor 9, in revolutions per minute (rpm), which corresponds to the absolute maximum speed likely to be encountered by the fan shaft during the entire flight;

[0076] R e is the external radius of the fan rotor 9 and corresponds to the distance between the axis of rotation X and a point of intersection between the apex 21 and a leading edge 22 of the blades 14 of the fan rotor 9, and is expressed in meters (m);

[0077] Ri is the inner radius of the fan rotor 9 and corresponds to a distance between the axis of rotation X and a point of intersection between the leading edge 21 of the blades 14 of the fan rotor 9 and a surface of the fan rotor 9 which radially delimits on the inside a flow vein in the fan rotor 9 and is expressed in meters (m); and

[0078] F = 39.7 N _1 ' 35 .(rpm)' 2 .m -5 .

[0079] As previously stated, FN is determined when the propulsion system 1 is stationary, uninstalled, in takeoff mode in a standard atmosphere (as defined by the International Civil Aviation Organization (ICAO) manual, Doc 7488 / 3, 3 e edition) and at sea level. The radii R9 and Ri, on the other hand, are determined when the propulsion system 1 is cold to simplify the measurements.

[0080] The leading edge 22 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 lower surface flow and an upper surface flow.

[0081] The mechanical load parameter kmeca translates the stress applied to a fan blade at the internal radius Ri of the fan rotor 9 (at the bottom of the part of the fan blade 14 which is configured to extend into the flow path): k meca = c 2 x S where S is the area of ​​the flow path at the inlet of the fan rotor 9, which is equal to ~ x (R e 2 - Ri 2 ) and is expressed in square meters.

[0082] The centrifugal forces applied to the fan blades 14 of a fan rotor 9 whose aerodynamic load parameter kaero is greater than or equal to 0.07 and less than or equal to 0.12 and whose mechanical load parameter kmeca is greater than or equal to 11.5 x 10 6 tr 2 .m 2 .min' 2 and less than or equal to 20.0 x 10 6 tr 2 .m 2 .min -2 are moderate. The aerodynamic forces are, on the other hand, higher than in conventional geared engines, but remain admissible. In particular, the fan shaft 20 (and indirectly the low-pressure shaft 11) of the propulsion system 1 is capable of transmitting the aerodynamic torque to the fan rotor 9, for example via splines or a bolted connection, without risking accidental breakage or reducing the service life of the fan section 2.

[0083] In order to reduce the mechanical load parameter kmeca of the fan section 2, it is possible, for example, to modify the hub-to-head ratio of the fan rotor 9. The hub-to-head ratio corresponds in fact to the ratio between the internal radius Ri and the external radius R e of the fan rotor 9. The hub-to-head ratio can be changed by changing the diameter D of the fan rotor 9 (which is equal to twice the external radius R e - note that since Figure 2 is a partial view, the diameter D is only partially visible) and / or by modifying the internal radius Ri of the fan rotor 9. However, increasing the diameter of the fan rotor 9 has the effect of increasing the centrifugal forces applied to the fan blades 14. Consequently, the hub-head ratio is, for example, modified by reducing the external radius R eof the fan rotor 9 or by increasing the internal radius Ri of the fan rotor 9. The hub-to-head ratio of the fan rotor 9 may for example be between 0.22 and 0.32. In the case of a fixed-pitch fan rotor, the hub-to-head ratio may be between 0.22 and 0.32. In the case of a variable-pitch fan rotor, the hub-to-head ratio is for example between 0.24 and 0.32 in order to allow the integration of the pitch change mechanism 15. If necessary, to achieve such hub-to-head ratios, the configuration of the fan shaft 20 and / or the bearings which support the fan shaft 20 may be adapted in order to facilitate the integration of the fan rotor 9 and the various components of the propulsion system 1 placed upstream of the reduction mechanism 19.

[0084] In order to obtain an aerodynamic load parameter kaero of between 0.07 and 0.12, the fan rotor 9 comprises at least twelve fan blades 14 and at most eighteen fan blades 14, for example at least fourteen fan blades 14 and at most sixteen fan blades 14.

[0085] The outer radius R e of the fan rotor (9) can then be between 40 inches (101.6 cm) and 92.5 inches (233.7 cm) inclusive, for example between 60 inches (152.4 cm) and 85 inches (215.9 cm) inclusive, for example of the order of 478 inches (198.1 cm), which allows the propulsion system 1 to be integrated in a conventional manner, in particular under the wing of an aircraft.

[0086] Such an external radius R e, combined with a hub-to-head ratio of between 0.22 and 0.32, thus makes it possible to minimize drag, without penalizing the mechanical loading, the aerodynamic loading of the fan rotor or the fan pressure ratio.

[0087] A propulsion system 1 having kmeca mechanical loading parameters and kaero aerodynamic loading parameters as described above can then have a high bypass ratio, including when the thrust class of the propulsion system is medium. Typically, the propulsion system 1 can be configured to provide thrust is configured to provide thrust of between 18,000 Ibf (80,068 N) and 51,000 Ibf (22,2411 N), for example between 20,000 Ibf (88,964 N) and 35,000 Ibf (15,5688 N) and have a bypass ratio greater than or equal to 10, in particular greater than or equal to 40, for example between 40 and 80 inclusive. The propulsion system 1 can then have a high dilution ratio while being able to be integrated under the wing of an aircraft 100. The peripheral speed at the tip 21 of the blades 14 of the fan rotor(s) 9 can also be between 210 m / s and 260 m / s in the case of unducted fan rotor(s) 9.

[0088] The reduction ratio of the reduction mechanism 19 may be between 5.0 and 11.0. Such reduction ratios then make it possible to obtain a fan section 2 in which the fan rotor 9 rotates at a rotational speed co such that the aerodynamic load parameter is greater than or equal to 0.07 and less than or equal to 0.12 and the mechanical load parameter is greater than or equal to 11.5 x 10 6 tr 2 .m 2 .min 2 and less than or equal to 20.0 x 10 6 tr 2 .m 2 .min -2 .

[0089] A double-body propulsion system 1 having an aerodynamic load parameter kaero greater than or equal to 0.07 and less than or equal to 0.12 and a mechanical load parameter kmeca greater than or equal to 11.5 x 10 6 tr 2 .m 2 .min -2 and less than or equal to 20.0 x 10 6 tr 2 .m 2 .min -2may 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.

[0090] Comparative example:

[0091] Engine 1 is a twin-body propulsion system comprising an unducted fan section 2 corresponding to the current technical standard (at the date of filing of this application) which we seek to improve.

[0092] The engine 2 is a double-body propulsion system 1 comprising an unducted fan section 2 in accordance with the teaching of the present application having an aerodynamic load parameter kaero equal to 0.075 and a mechanical load parameter kmeca equal to 14.6 x 10® tr 2 .m 2 .min' 2 .

[0093] Engine 1 has an aerodynamic load parameter less than 0.07 and a mechanical load parameter less than 11.5 x 10 6 tr 2 .m 2 .min 2 . Engine 1 has low aerodynamic and mechanical loads to the benefit of fan efficiency at the expense of fan size and mass. The size of the fan has a significant impact on the installation of the engine on aircraft.

[0094] The integration of engine 2 into the aircraft is simplified, compared to engine 1, without degrading its energy performance. The increase in aerodynamic load and the increase in the fan pressure rate make it possible to reduce the fan diameter while maintaining thrust.

[0095] This is facilitated by the increase in mechanical load which allows an increase in the rotation speed of the fan rotor.

[0096] Compared with engine 1, fan section 2 of engine 2 is subjected, particularly at takeoff, to slightly greater mechanical and aerodynamic stresses, while remaining within an acceptable range for the sizing of the engine, so that fan shaft 10 is capable of transmitting the aerodynamic torque to fan rotor 9. On the other hand, engine 2 is more compact and has a lower mass than engine 1, thanks in particular to the reduction in its diameter and its bypass ratio, without reducing its efficiency. In the present comparative example, the mass reduction of 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 approximately 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.

[0097] 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 less than 1.2) 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.

[0098] It is possible to increase the value of the aerodynamic load parameter with limited impact on the fan efficiency via a reduction of the fan blades. Reducing the fan blade aspect ratio is usually considered unfavorable for the fan efficiency and the blade vibration situation with respect to the first torsional mode.

[0099] However, improvements in aerodynamic methods can further minimize the impact on fan efficiency. Reducing the blade height and using a three-dimensionally woven composite with an appropriate weave pattern or incorporating stiffer fibers can restore the vibration situation to the first torsional mode.

[0100] The use of fan blades comprising a three-dimensionally woven composite material with stiffer fibers also makes it possible to withstand increased centrifugal mechanical stresses, particularly in the event of bird ingestion. The use of bearing steels with increased mechanical properties or ceramic rolling elements makes it possible to have the same hub ratio on engine 2 as on engine 1 despite the increase in the mechanical load parameter. The use of a titanium alloy for the hub with improved mechanical properties helps to meet the target hub ratio.

[0101] The overall compression ratio and high-pressure turbine inlet temperature of engine 2 have also been increased to improve the engine's thermal efficiency.

Claims

CLAIMS 1. Aeronautical propulsion system (1) comprising: - a drive shaft (11) movable in rotation around an axis of rotation; - a fan shaft (20); - a fan section (2) comprising an unducted fan rotor (9) driven in rotation 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); the propulsion system (1) being configured so that the fan rotor (9) has an aerodynamic load parameter greater than or equal to 0.07 and less than or equal to 0.12 and a mechanical load parameter greater than or equal to 11.5 x 10® rpm 2 .min 1and less than or equal to 20.0 x 10® tr.m 2 .min 1 , where the aerodynamic load and the mechanical load are defined by the following formulas: and where: kaero is the aerodynamic load parameter; kmeca is the mechanical load parameter; FN is the thrust generated by the propulsion system (1) 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); BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is stationary in take-off regime in a standard atmosphere and at sea level; n is the number of blades (14) in the fan rotor (9); F = 39.7 N _1 ' 35 .(rpm)' 2 .m -5 R eis the external radius of the fan rotor (9) and corresponds to a distance between the axis of rotation (X) and a point of intersection between a vertex (21) and a leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in meters (m); Ri is the inner radius of the fan rotor and corresponds to a distance between the axis of rotation (X) and a point of intersection between the leading edge of the blades (14) of the fan rotor (9) and a surface of the fan rotor (9) which radially delimits on the inside a flow vein in the fan rotor (9) and is expressed in meters (m); co is the limiting rotational speed of the fan rotor (9) and is expressed in revolutions per minute (rpm); and S is the area of ​​the flow path at the inlet of the fan rotor (9), which is equal to jt x (Re 2 - Ri 2 ) and is expressed in square meters (m 2 ).

2. A propulsion system (1) according to claim 1, wherein the outer radius (Re) of the fan rotor (9) is between 40 inches (101.6 cm) and 92.5 inches (233.7 cm) inclusive, for example between 60 inches (152.4 cm) and 85 inches (215.9 cm) inclusive, for example in the order of 478 inches (198.1 cm).

3. Propulsion system (1) according to one of claims 1 and 2, in which a hub-to-head ratio of the fan rotor (9) is between 0.22 and 0.

32.

4. Propulsion system (1) according to one of claims 1 to 3, wherein the fan rotor (9) comprises at least twelve fan blades (14) and at most eighteen fan blades (14), for example at least fourteen fan blades (14) and at most sixteen fan blades (14).

5. Propulsion system (1) according to one of claims 1 to 4, in which a reduction ratio of the reduction mechanism (19) is greater than or equal to 5.0 and less than or equal to 11.

0.

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

7. Propulsion system (1) according to one of claims 1 to 6, 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.

8. Propulsion system (1) according to one of claims 1 to 7, in which the fan section (2) further has a fan compression ratio, corresponding to a pressure ratio between an outlet of the fan rotor (9) and an inlet of the fan rotor (9) less than or equal to 1.45, for example less than or equal to 1.

30.

9. Propulsion system (1) according to one of claims 1 to 8, wherein the propulsion system (1) is dimensioned so that a thrust of the propulsion system, when the propulsion system (1) is stationary in take-off mode in a standard atmosphere, is between 18,000 Ibf (80,068 N) and 51,000 Ibf (22,2411 N).

10. Propulsion system (1) according to one of claims 1 to 9, further comprising a drive turbine (7) and a compressor (4) directly connected by the drive shaft (11), wherein the drive turbine (7) comprises at least three and at most five stages.

11. Propulsion system (1) according to claim 10, wherein the compressor (4) comprises at least two and at most four stages.

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

13. Propulsion system (1) according to claim 12, wherein the high pressure compressor (5) comprises at least eight and at most eleven stages.

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

15. Method for dimensioning a propulsion system (1) comprising a reduction mechanism (19) coupling a drive shaft (11) and an unducted fan rotor (9) to drive the fan rotor (9) at a speed lower than a speed of the drive shaft (11), wherein the fan rotor (9) has an aerodynamic load parameter greater than or equal to 0.07 and less than or equal to 0.12 and a mechanical load parameter greater than or equal to 11.5 x 10 e tr.m2 .min' 1 and less than or equal to 20.0 x 10 e tr.m 2 .min' 1 , where the aerodynamic load and the mechanical load are defined by the following formulas: and where . kaero is the aerodynamic load parameter; kmeca is the mechanical load parameter; FN is the thrust generated by the propulsion system (1) 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); BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is stationary in take-off regime in a standard atmosphere and at sea level; n is the number of blades (14) in the fan rotor (9); F = 39.7 N -1 ' 35 .(rpm)' 2 .irr 5 R eis the external radius of the fan rotor (9) and corresponds to a distance between the axis of rotation (X) and a point of intersection between a vertex (21) and a leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in meters (m); Ri is the inner radius of the fan rotor and corresponds to a distance between the axis of rotation (X) and a point of intersection between the leading edge of the blades (14) of the fan rotor (9) and a surface of the fan rotor (9) which radially delimits on the inside a flow vein in the fan rotor (9) and is expressed in meters (m); co is the rotational speed Imite of the fan rotor (9) and is expressed in revolutions per minute (rpm); and S is the area of ​​the flow path at the inlet of the fan rotor (9), which is equal to (Re 2 - Ri 2 ) and is expressed in square meters (m 2 ).

16. Method of manufacturing an aeronautical propulsion system comprising the following steps: - dimensioning the aeronautical propulsion system in accordance with the method according to claim 15; and - manufacturing the aeronautical propulsion system.

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