Aeronautical propulsion system having a ducted fan and a high bypass ratio

Abstract

The invention relates to an aeronautical propulsion system comprising: - a ducted fan section (2) comprising a fan (22), the fan having a fan rotor (9) comprising blades (14); - a gas generator (3) comprising a drive shaft (10, 11) configured to directly or indirectly rotate the fan rotor (9), wherein: a. 0 ≤< P.D.10-4 ≤ 160 and b. 2.7. (P. D.10-4)0,6≤ 100.D2≤ BPR+ 1≤3.9(P.D.10-4)0.6 where d is a diameter of the fan rotor in metres that is measured in a plane normal to an axis (X) of rotation of the fan rotor (9) at the intersection (24) between a tip edge (21) and a leading edge (23) of the blades of the fan rotor (9); P is a thrust generated by the fan (22) in Newtons (N) when the propulsion system is stationary at takeoff speed in a standard atmosphere and at sea level; and BPR is a bypass ratio of the flow passing through the propulsion system defined as a ratio of a mass flow rate of a secondary air flow (F2) flowing through the propulsion system around the gas generator (3) to a primary air flow (F1) flowing through the gas generator (3) measured when the propulsion system (1) is stationary and inactive at a speed corresponding to a takeoff speed of an aircraft comprising the aeronautical propulsion system in a standard atmosphere and at sea level.

Description

[0001] DESCRIPTION

[0002] TITLE: AERONAUTICAL PROPULSIVE SYSTEM WITH DUCTED FAN AND HIGH DILUTION RATE

[0003] TECHNICAL FIELD

[0004] This disclosure relates generally to the field of propulsion systems, and more particularly to aeronautical propulsion systems having a ducted fan section and a high or even very high bypass ratio.

[0005] STATE OF THE ART

[0006] An aircraft 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 a high pressure turbine and a low pressure turbine.

[0007] When the propulsion system is in operation, the high-pressure compressor is rotated by the high-pressure turbine via a high-pressure shaft. The fan and, if applicable, the low-pressure compressor are rotated by the low-pressure turbine via a low-pressure shaft.

[0008] One of the objectives of technological research is to improve the environmental performance of aircraft. Therefore, in all phases of design and development, relevant factors are taken into account to obtain less energy-consuming, 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.

[0009] For example, in order to improve the propulsive efficiency of an aircraft propulsion system and reduce its specific consumption, it has been found that it is advantageous to increase the rotational speed of the low-pressure turbine and the low-pressure compressor, and to decrease the rotational speed of the fan. Similarly, in order to limit the drag and mass of the aircraft, and thus reduce its fuel consumption, it has been found that it is advantageous to make propulsion systems more compact, that is to say to reduce the size of all or part of their components.

[0010] However, in doing so, it appears that the rotating components of aeronautical propulsion systems, in particular those present within the compressor section and the turbine section, are subjected to increased centrifugal forces, while having smaller dimensions, which is likely to complicate the design of propulsion systems to avoid altering their mechanical strength and / or limiting their service life.

[0011] SUMMARY

[0012] An aim of this presentation is to reduce the fuel consumption of an aeronautical propulsion system without limiting the service life of its rotating components.

[0013] This goal is achieved through an aeronautical propulsion system, comprising:

[0014] - a ducted fan section comprising a fan, the fan having a fan rotor comprising blades,

[0015] - a gas generator comprising a drive shaft configured to drive, directly or indirectly, the fan rotor in rotation, in which the propulsion system is dimensioned so as to verify the following pair of relationships: a. 0 < PD10 -4 < 160 and b. 2.7. ( v 3.9. (PD ÎO -4 ) J06 where D is a fan rotor diameter in meters that is measured in a plane normal to an axis of rotation of the fan rotor, at an intersection between a tip edge and a leading edge of the fan rotor blades;

[0016] P is a maximum thrust generated by the fan in Newton when the propulsion system is stationary in takeoff mode, in a standard atmosphere and at sea level; and BPR is a flow dilution ratio through the propulsion system, defined as a ratio between a mass flow rate of a secondary air flow flowing in the propulsion system around the gas generator, and a primary air flow flowing in the gas generator, measured when the propulsion system is stationary, uninstalled, in a mode corresponding to a takeoff of an aircraft comprising the aeronautical propulsion system, in a standard atmosphere and at sea level.

[0017] According to one embodiment, the dilution rate is between 10 and 20 inclusive.

[0018] According to one embodiment, the gas generator comprises at least one compressor downstream of the fan, an overall compression ratio of the propulsion system being between 40 and 60 inclusive, the overall compression ratio being defined as a pressure ratio between a total pressure measured upstream of a combustion chamber of the propulsion system and a pressure at the inlet of the fan rotor measured upstream of a foot of the fan rotor.

[0019] According to one embodiment, the fan blades are fixed-pitch blades.

[0020] According to one embodiment, the aeronautical propulsion system comprises a reduction mechanism coupling in rotation:

[0021] - the drive shaft and

[0022] - a fan shaft driving the fan rotor, the propulsion system being configured to drive the fan shaft at a rotational speed lower than the rotational speed of the drive shaft, a reduction ratio of the reduction mechanism being preferably greater than or equal to 2.5 and less than or equal to 11, still preferably greater than or equal to 2.7 and less than or equal to 6.0, still preferably greater than or equal to 2.7 and less than or equal to 3.6, still preferably substantially 3.0.

[0023] According to one embodiment, the diameter of the fan rotor is greater than or equal to 1.270 m and less than or equal to 3.048 m, preferably greater than or equal to 2.159 m and less than or equal to 3.048 m, still preferably substantially equal to 2.286 m.

[0024] According to one embodiment, a hub-to-head ratio of the fan rotor is greater than or equal to 0.22 and less than or equal to 0.32, preferably greater than or equal to 0.235 and less than or equal to 0.30, more preferably less than or equal to 0.27, the hub-to-head ratio being defined as a ratio Rt / R e , where: - R, is an internal radius of the fan rotor, measured between the axis of rotation X and a point of intersection between the leading edge of the fan blades and the aerodynamic surface of a fan rotor platform radially delimiting inside the flow vein at the inlet of the fan rotor;

[0025] - R e is an external radius of the fan rotor and is equal to one half of the fan diameter.

[0026] According to one embodiment, the fan rotor comprises at least fourteen blades and at most twenty-four blades, preferably at least sixteen blades and at most twenty-four blades, still preferably twenty-two blades.

[0027] According to one embodiment, the gas generator comprises a high pressure body and a low pressure body, the low pressure body rotating at a lower rotational speed than the high pressure body.

[0028] According to one embodiment, the low pressure body comprises a low pressure turbine, the low pressure turbine having at least three stages and at most eight stages.

[0029] According to one embodiment, the low pressure body comprises a low pressure compressor, the low pressure compressor having at least two stages and at most five stages.

[0030] According to one embodiment, the high pressure body comprises a high pressure turbine, the high pressure turbine being two-stage.

[0031] According to one embodiment, the high pressure body comprises a high pressure compressor, the high pressure compressor comprising at least eight and at most eleven stages.

[0032] Furthermore, an aircraft comprising an aeronautical propulsion system as defined above is proposed.

[0033] Further provided is a method of manufacturing an aeronautical propulsion system, the propulsion system comprising:

[0034] - a ducted fan section comprising a fan, the fan having a fan rotor comprising blades,

[0035] - a gas generator comprising a drive shaft configured to drive, directly or indirectly, the fan rotor in rotation, the manufacturing method comprising a step of dimensioning the fan rotor during which a diameter D of the fan rotor in meters and measured in a plane normal to an axis of rotation of the fan rotor, at an intersection between an end edge and a leading edge of the blades of the fan rotor is chosen so that: a. 0 < PD10 -4 < 160 and b. 2.7. ( v 3.9. (PD ÎO -4 ) J06where P is a maximum thrust generated by the fan in Newtons when the propulsion system is stationary in take-off mode, in a standard atmosphere and at sea level; and BPR is a flow dilution ratio of the flow through the propulsion system, defined as a ratio between a mass flow rate of a secondary air flow flowing in the propulsion system around the gas generator, and a primary air flow flowing in the gas generator, measured when the propulsion system is stationary, uninstalled, in a mode corresponding to a take-off of an aircraft comprising the aeronautical propulsion system, in a standard atmosphere and at sea level.

[0036] PRESENTATION OF THE DRAWINGS

[0037] Other characteristics and advantages will emerge from the following description, which is purely illustrative and not limiting, and must be read in conjunction with the appended figures, among which:

[0038] - figure 1 schematically represents an aircraft comprising propulsion systems,

[0039] - figure 2 schematically represents, in partial view and in section, an example of a propulsion system with a ducted fan section,

[0040] - figure 3 schematically represents a first example of a planetary reduction mechanism,

[0041] - figure 4 schematically represents a first example of an epicyclic reduction mechanism,

[0042] - figure 5 is a graph representing the dimensioning domain of the aeronautical propulsion system according to one embodiment,

[0043] - Figure 6 schematically represents a fan section of an aeronautical propulsion system. DETAILED DESCRIPTION OF AN EMBODIMENT

[0044] In the example illustrated in Figure 1, the aircraft is an airplane 100 comprising a fuselage 101 and two wings 102. In this example, the aircraft comprises two propulsion systems 1, each propulsion system 1 being attached to a respective wing 102 of the airplane 100 via a pylon. In another embodiment, the aircraft could comprise one or more propulsion systems attached to the fuselage 101.

[0045] Figure 2 schematically represents, in partial view and in section, a first example of a propulsion system 1.

[0046] In this example, propulsion system 1 is a twin-spool, ducted-fan gas turbine engine.

[0047] In Figure 2, the propulsion system 1 has a main direction extending along a longitudinal axis X. The propulsion system 1 comprises a fan section 2 and a primary body 3, often called a “gas generator”.

[0048] The blower section 2 comprises a blower 22 and a blower housing 12. The blower 22 comprises a blower rotor 9. The blower housing 12 surrounds the blower rotor 9. The blower rotor 9 is rotatably mounted relative to the blower housing 12.

[0049] The fan rotor 9 comprises a fan hub 13 and fan blades 14 extending radially from the hub 13. The fan blades 14 may be fixed relative to the fan hub 13 or have a variable pitch. In the latter case, not shown, each of the fan blades 14 is pivotally mounted relative to the fan hub 13 along a pitch axis and is connected to a pitch change mechanism mounted in the propulsion system 1. The pitch change mechanism makes it possible to adjust the pitch angle of the fan blades 14 as a function of the flight phases. Such a mechanism is known in principle to those skilled in the art familiar with the architectures of aircraft propulsion systems.

[0050] The fan rotor 9 comprises at least fourteen fan blades 14 and at most twenty-four fan blades 14, preferably at least sixteen fan blades 14, preferably twenty-two fan blades 14.

[0051] In addition, in this example, the fan section 2 also comprises a fan stator 16 fixedly mounted on the fan casing 12. The fan stator 16 comprises fixed vanes 17 generally referred to as “outlet guide vanes” (or “OGV”). This set of fixed vanes has the function of straightening and regulating the airflow flowing downstream of the fan rotor 9 to contribute to the thrust of the engine. This set of fixed vanes also acts as a noise reducer.

[0052] Alternatively, the outlet blades 17 could have a variable pitch. If necessary, and similarly to the fan blades 14 of the fan rotor 9, the root of the outlet blades 17 is pivotally mounted along a pitch axis and is connected to a pitch change mechanism (not shown), the pitch being adjusted according to the flight phases by the pitch change mechanism.

[0053] The number of outlet blades 17 depends on the acoustic criteria defined for the propulsion system 1 and is at least equal to the number of fan blades 1.

[0054] The primary body 3 comprises a compressor section 29, a combustion chamber 6 and a turbine section 30.

[0055] The compressor section 29 comprises a low pressure compressor 4 and a high pressure compressor 5.

[0056] The low pressure compressor 4 comprises a rotor 41 capable of being driven in rotation relative to the casing 31 of the propulsion system 1 and a stator 42 fixedly mounted on the casing 31.

[0057] The rotor 41 of the low pressure compressor 4 comprises moving wheels 4a and the stator 42 of the low pressure compressor 4 comprises fixed wheels 4b. The moving wheels 4a are arranged alternately with the fixed wheels 4b, thus forming a succession of low pressure compressor stages.

[0058] Likewise, the high-pressure compressor 5 comprises a rotor 51 capable of being driven in rotation relative to the casing 31 of the propulsion system 1 and a stator 52 fixedly mounted on the casing 31.

[0059] The rotor 51 of the high-pressure compressor 5 comprises moving wheels 5a and the stator 52 of the high-pressure compressor 5 comprises fixed wheels 5b. The moving wheels 5a are arranged alternately with the fixed wheels 5b, thus forming a succession of high-pressure compressor stages.

[0060] The turbine section 30 comprises a high pressure turbine 7 and a low pressure turbine 8.

[0061] The high-pressure turbine 7 comprises a rotor 71 capable of being driven in rotation relative to the casing 31 of the propulsion system 1 and a stator 72 fixedly mounted on the casing 31. The rotor 71 of the high-pressure turbine 7 comprises moving wheels 7a and the stator 72 of the high-pressure turbine 7 comprises fixed wheels 7b. The moving wheels 7a are arranged alternately with the fixed wheels 7b, thus forming a succession of high-pressure turbine stages.

[0062] Likewise, the low pressure turbine 8 comprises a rotor 81 capable of being driven in rotation relative to the casing 31 of the propulsion system 1 and a stator 82 fixedly mounted on the casing 31.

[0063] The rotor 81 of the low pressure turbine 8 comprises moving wheels 8a and the stator 82 of the low pressure turbine 8 comprises fixed wheels 8b. The moving wheels 8a are arranged alternately with the fixed wheels 8b, thus forming a succession of low pressure turbine stages.

[0064] The propulsion system 1 comprises a low pressure shaft 11 connecting the rotor 81 of the low pressure turbine 8 to the rotor 41 of the low pressure compressor 4, the low pressure shaft 11 being rotatably mounted relative to the casing 31 around the longitudinal axis X.

[0065] When the propulsion system 1 is in operation, the rotor 81 of the low pressure turbine 8 drives the rotor 41 of the low pressure compressor 4 in rotation via the low pressure shaft 11.

[0066] The propulsion system 1 further comprises a fan shaft 20 and a reduction mechanism 19. The fan rotor 9 is fixedly mounted on the fan shaft 20. The reduction mechanism 19 has an inlet and an outlet. The inlet of the reduction mechanism 19 is connected to the low-pressure shaft 11 and the outlet of the reduction mechanism 19 is connected to the fan shaft 20. Thus, when the propulsion system 1 is in operation, the rotor 81 of the low-pressure turbine 8 rotates not only the rotor 41 of the low-pressure compressor 4, but also the fan rotor 9, via the low-pressure shaft 11, the reduction mechanism 19 and the fan shaft 20.

[0067] By means of the reduction mechanism 19, the fan rotor 9 is rotated at a speed lower than the rotational speed of the rotor 81 of the low pressure turbine 8.

[0068] The reduction mechanism 19 thus makes it possible to independently control the rotation speed of the fan 22 and the rotation speed of the low-pressure turbine 8 and the low-pressure compressor 4. The low-pressure turbine 8, the low-pressure shaft 11, the low-pressure compressor A, the fan shaft 20, the reduction mechanism 19 and the fan 22 together form the “low-pressure body” of the propulsion system 1.

[0069] The propulsion system 1 further comprises a high pressure shaft 10 connecting the rotor 71 of the high pressure turbine 7 to the rotor 51 of the high pressure compressor 5, the high pressure shaft 10 being rotatably mounted relative to the casing 31 around the longitudinal axis X. The high pressure shaft 10 is coaxial with the low pressure shaft 11 and extends around the low pressure shaft 11.

[0070] When the propulsion system 1 is in operation, the rotor 71 of the high-pressure turbine 7 drives the rotor 51 of the high-pressure compressor 5 into rotation via the high-pressure shaft 10.

[0071] The high-pressure turbine 7, the high-pressure shaft 10 and the high-pressure compressor 5 together form the “high-pressure body” of the propulsion system 1.

[0072] The low pressure shaft 11 and the high pressure shaft 10 may be co-rotating, i.e. driven in the same direction of rotation about the longitudinal axis X. Alternatively, the low pressure shaft 11 and the high pressure shaft 10 may be counter-rotating, i.e. driven in opposite directions of rotation about the longitudinal axis X.

[0073] The dual-body propulsion system 1 may in particular comprise a single-stage high-pressure turbine 7, i.e. comprising exactly one stage, or a two-stage high-pressure turbine 7, i.e. comprising exactly two stages (as illustrated in the example of figure 2).

[0074] The high-pressure compressor 5 comprises at least eight stages (as illustrated in the example of Figure 2) and at most eleven stages.

[0075] The low pressure turbine 8 comprises at least three stages (as illustrated in the example of Figure 2) and at most eight stages.

[0076] Low pressure compressor A has at least two stages and at most five stages.

[0077] When the propulsion system is in operation, an air flow F entering the propulsion system 1 passes through the fan 22 and is then divided between a primary air flow F1 and a secondary air flow F2, which circulate from upstream to downstream in the propulsion system 1.

[0078] The secondary air flow F2, also called the "bypass air flow", flows in the secondary vein, around the primary body 3. The secondary air flow F2 allows the periphery of the primary body 3 to be cooled and is used to generate the majority of the thrust provided by the propulsion system 1.

[0079] The primary air flow F1 flows in a primary vein 29 inside the primary body 3, passing successively through the compressor section 29 (low pressure compressor 4 and high pressure compressor 5), the combustion chamber 6 where it is mixed with fuel to serve as an oxidizer, and the turbine section 30 (high pressure turbine 7 and low pressure turbine 8). The passage of the primary air flow F1 through the turbine section 30 receiving energy from the combustion chamber 6 causes rotation of the movable wheels 7a, 8a of the turbine section 30, which in turn drive rotation of the movable wheels 4a, 5a of the compressor section 29 as well as the fan rotor 9.

[0080] In a propulsion system including a reduction mechanism 19 such as that illustrated in FIG. 2, the decoupling between the rotational speed of the fan 22 and the rotational speed of the low-pressure turbine 8 makes it possible to reduce the rotational speed and the pressure ratio of the fan rotor 9 while increasing the power extracted by the low-pressure turbine 8. Indeed, the overall efficiency of the propulsion system 1 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 as it passes through the propulsion system 1.In a high bypass ratio propulsion system, the bulk of the flow generating the propulsive force is made up of 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. In order to improve 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 16 (or, in the absence of stator 16, of the fan rotor 9) and the average pressure at the inlet of the fan rotor 9, is less than or equal to 1.70, preferably less than or equal to 1.50, for example between 0.90 and 1.45.The average pressures are measured here over the height of at least one of the fan blades 14, i.e. from the surface which radially delimits the air flow path at the inlet of the fan rotor 9 to the end edge 21 of the fan blade 14. By end edge is meant the edge of the fan blade 14 furthest from the axis of rotation X of the fan rotor 9 (top).

[0081] The peripheral speed at the end edge 21 of the fan blades 14 can also be between 260 meters per second (ms 1 ) and 400 meters per second (ms 1 ) included. The blower pressure ratio can then be between 1.20 and 1.45.

[0082] In a direct-drive propulsion system, the fan rotor 9 can, alternatively, be directly coupled to the low-pressure shaft 11, i.e. without a reduction mechanism. The low-pressure shaft 11 is then combined with the fan shaft 20 so that the fan rotor 9 is driven by the low-pressure shaft 11 at the same rotational speed as the rotor 81 of the low-pressure turbine 8.

[0083] It should be noted that, since Figure 2 is a partial view, the diameter D is only partially visible.

[0084] The reduction mechanism 19 may comprise an epicyclic or planetary, single-stage or two-stage reduction mechanism.

[0085] For example, Figure 3 illustrates a reduction mechanism 19 of the planetary type (or "star" in English). The reduction mechanism 19 comprises a sun gear 19a (input of the reduction mechanism 19), centered on an 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 gear 19b (output of the reduction mechanism 19) coaxial with the sun gear 19a and configured to drive in rotation the fan shaft 20 about its axis X of rotation, and a series of satellites 19c distributed circumferentially about the axis X of rotation of the rotor 9 of the fan section 2, 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.

[0086] In another example, Figure 4 illustrates a reduction mechanism 19 of the epicyclic type (or "planetary" in English), 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. Whatever the configuration of the reduction mechanism 19, the diameter of the crown 19b and of the planet carrier 19d are greater than the diameter of the sun gear 19a, so that the rotational speed of the rotor 9 of the fan section 2 is lower than the rotational speed of the low pressure shaft 11.

[0087] The reduction ratio of the reduction mechanism 19 is less than or equal to 11, for example greater than or equal to 2.5 and less than or equal to 11, preferably greater than or equal to 2.7 and less than or equal to 6.9, typically around 3.0.

[0088] It should be noted that, in this application, and unless otherwise stated, the parameters (pressure, flow rate, thrust, speed, etc.) are systematically determined in stationary conditions, in "non-installed" conditions, i.e. on a test bench and not installed on an aircraft (the measurements then being simpler to carry out), in take-off mode in a standard atmosphere and at sea level. Such conditions are, for example, as defined in the International Civil Aviation Organization (ICAO) manual, Doc 7488 / 3, 3 eedition. Any other equivalent normative condition that is indicated for the certifications of aircraft propulsion systems may be considered, such as the standards of the EÀSÀ (European Union Aviation Safety Agency), such as its "Certification Specification - Engines" (CS-E) standards or its TCDS "Type Certificate Data Sheets" standards. Another normative text may be the American certification regulation 14-CFR. The distances (length, radius, diameter) are, however, measured at ambient temperature (approximately 20°C) when the propulsion system 1 is cold, i.e. when the propulsion system has been stopped for a sufficient period for the parts of the propulsion system to be at ambient temperature.

[0089] The present application proposes an aeronautical propulsion system sized so as to reduce the fuel consumption necessary for the production of a given thrust. According to its most general embodiment, the propulsion system is sized so as to verify the following pair of relationships:

[0090] 1) 0 < PD10 -4 < 160 and

[0091] 2) 2.7. ( 3.9. (PD 1O -4 ) 0 6 where D represents the diameter of the fan rotor, measured in a plane normal to the axis X of rotation of the fan rotor 9, at the intersection 24 between an end edge 21 and a leading edge 23 of the blades 14 of the fan rotor 9; and P represents a maximum thrust provided by the fan of the propulsion system, expressed in Newtons. For the sake of brevity, the expressions X = PD 10 will be defined below 4

[0092] 100 Z) 2 > i and Y = — BPR :+ , so that inequations 1 and 2 can be written

[0093] 1. 0 < X ​​< 160

[0094] 2. 2.7. 06 < Y < 3.9. X 06 .

[0095] The coefficients 2.7 and 3.9 of inequality 2 have the appropriate dimensions so that the three terms of the inequality have the same dimension.

[0096] With reference to Figure 5, we thus define a domain in which the dimensioning of the propulsion system is carried out. When the propulsion system is a double-body system, comprising a high-pressure body and a low-pressure body, the expression X, which corresponds to the abscissa of Figure 5, makes it possible to represent a torque provided by the low-pressure body. The expression Y, for its part, corresponds to the ordinate of Figure 5 and is an indirect representation of the size of the high-pressure body seen by the air passing through it (called "thermodynamic size"). The larger the size of the high-pressure body, the higher the mass flow rate of air passing through this body.

[0097] Thus, equation 1 defines a maximum value of the torque provided by the low-pressure body, while equation 2 defines lower and upper bounds of the thermodynamic size of the high-pressure body for a desired low-pressure shaft torque. The domain delimited by these equations corresponds to the area in Figure 5 between the two curves and the limit X = 160, and has the following advantages:

[0098] - a significant improvement in fuel consumption for a given thrust P. Indeed, inequality 2.7. X 06 < Y defines an upper limit of dimension for the fan rotor, beyond which the fuel consumption becomes too high,

[0099] - a possibility of industrial exploitation. In other words, the dimensions of the propulsion system and / or the materials which make it possible to verify the inequality Y < 3.9. X 06 are realistic.

[0100] Below this limit, the high-pressure body becomes too small, so that marginal losses become too large compared to the thrust provided by the high-pressure body. For example, and purely illustratively, gaps between the tips of the high-pressure turbine blades and the casing 31 of the propulsion system cannot be reduced beyond a certain limit. When the high-pressure body - in particular the high-pressure turbine blades - reaches a critical lower dimension, the size of these gaps and the associated energy losses become too large compared to the contribution of the high-pressure body to the thrust.

[0101] Furthermore, downsizing the high-pressure body requires an increase in temperature in the combustion chamber 6, and therefore materials capable of withstanding high temperatures. Below a lower limit of high-pressure body size, no known material is capable of withstanding the corresponding increase in temperature.

[0102] Compared with the aeronautical propulsion systems known from the state of the art, the propulsion system according to the present application comprises a small high-pressure body (radially and axially) and a large low-pressure body (in particular radially). An advantage of such a configuration is to improve the efficiency of the high-pressure body, that is to say to allow a lower specific fuel consumption (SFC) for the same power supplied by the high-pressure body.

[0103] The propulsion system 1 is configured to provide a maximum thrust of 51,000 Ibf (226,859 N), in particular between 18,000 Ibf (80,068 N) and 51,000 Ibf (226,859 N), possibly more precisely between 19,000 Ibf (84,516 N) and 38,000 Ibf (169,032 N), preferably between 20,000 Ibf (88,964 N) and 35,000 Ibf (155,688 N).

[0104] The diameter D of the shrouded fan rotor 9 is preferably less than 3.048 m. The diameter D is preferably between 1.270 m and 3.048 m, more particularly between 2.159 m and 3.048 m, still preferably substantially equal to 2.286 m. This allows the integration of the propulsion system 1 in a conventional manner, in particular under a wing of the aircraft 1. The diameter of the fan rotor 9 is measured here in a plane P1 normal to the longitudinal axis X, which corresponds to the axis of rotation of the fan rotor 9, at an intersection 24 between an end edge 21 and a leading edge 23 of the fan blades 14.

[0105] 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, it is meant a bypass ratio greater than or equal to 10, for example between 10 and 80 inclusive, preferably between 10 and 35 inclusive, preferably between 10 and 18 inclusive. The bypass ratio is defined as a ratio between the mass flow rate of the secondary air flow F2 and the mass flow rate of the primary air flow F1. A propulsion system 1 as proposed, having these values ​​of diameter, maximum thrust and bypass ratio, complies with the two inequalities 1 and 2 defined above.

[0106] The hub-to-head ratio corresponds to the ratio between the internal radius Ri and the external radius Re 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 23 and the surface which radially delimits on the inside the flow vein 25 at the inlet of the fan rotor 9 (and corresponds to the point of connection of the leading edge 23 with the aerodynamic surface of a platform of the fan rotor 9). The external radius Re corresponds to the distance between the axis of rotation X and the point of intersection 24 between the leading edge 23 and the end edge 21 of the fan blades (and corresponds to half the fan diameter D). The lower the hub-to-head ratio, the more efficient the fan rotor 9 is. However, the reduction in the hub-to-head ratio of the fan rotor 9 implies an increase in the mechanical load of the hub 13 of the fan rotor 9.The sizing of the fan rotor 9 is such that its hub-to-head ratio is less than or equal to 0.32. The hub-to-head ratio is preferably between 0.22 and 0.32 inclusive. This makes it possible in particular to obtain advantageous thrust via the secondary air flow F2 compared to the flow entering the primary air flow F1.

[0107] The overall compression ratio of the propulsion system (in English, Overall Pressure Ratio or OPR) is defined as the pressure ratio between the pressure at the outlet of the compressor - or in the case where the propulsion system comprises several compressors, of the compressor located furthest downstream of the flow - and the pressure at the inlet of the fan rotor 9, the latter being measured upstream of the root of the fan rotor 9. The pressure at the outlet of the compressor corresponds to the total pressure upstream of the combustion chamber 6.

[0108] Thus, for a twin-spool propulsion system such as that shown in Figure 2, the overall compression ratio is the ratio between the pressure at the outlet of the high-pressure compressor 5, which corresponds to the total pressure upstream of the combustion chamber 6, and the pressure at the inlet of the fan rotor 9 measured upstream of the root of the fan rotor 9. Advantageously, the overall compression ratio is less than or equal to 60, preferably between 40 and 60. A high total compression ratio makes it possible to reduce the specific fuel consumption of the engine, but increases the temperature downstream of the compressors, which can compromise the structural integrity of the materials forming the last compressor stages. In addition, a high total compression ratio requires several compressor stages and therefore a high total mass of the compressors.Thus, a total compression ratio between 40 and 60 allows relatively low specific consumption while allowing a moderate temperature downstream of the compressors.

[0109] Table 2 below presents a comparative example between a reference propulsion system corresponding to the current technical standard (at the filing date of this application), which is sought to be improved, and a propulsion system 1 included in the design domain defined by inequalities 1 and 2. The standard propulsion system and the propulsion system 1 conforming to this application are both of the double-body type and comprise a ducted fan section. The temperatures at the inlet of the low-pressure turbine rotor (point 6.) are Red Line (or R / L) temperatures, i.e. corresponding to limit values ​​beyond which the propulsion system concerned is not intended to operate in an ordinary manner.

[0110] [Table 1]

[0111] To move from the reference propulsion system to propulsion system 1 in accordance with the present application, the BPR was increased, allowing a gain in specific consumption. The fan diameter D was increased to allow a maximum thrust close to the maximum thrust of the reference engine to be maintained. The overall compression ratio (OPR) was also increased, as was the temperature at the inlet of the high-pressure turbine rotor, thus improving the thermal efficiency of propulsion system 1. This increase in temperature obtained for propulsion system 1 in accordance with the present application, in points 5. and 6. of table 2, is made possible by the use of turbine disks whose manufacture involves powder metallurgy, and by a first low-pressure turbine stage comprising blades made of ceramic matrix composite material.

[0112] An improvement in thermopropulsive efficiency of approximately 10% is observed for the propulsion system 1 in accordance with the present application in comparison with the reference propulsion system.

Claims

CLAIMS 1. Aeronautical propulsion system, comprising: - a shrouded fan section (2) comprising a fan (22), the fan having a fan rotor (9) comprising blades (14), - a gas generator (3) comprising a drive shaft (10, 11) configured to drive, directly or indirectly, the fan rotor (9) in rotation, in which the propulsion system (1) is dimensioned so as to verify the following pair of relationships: a. 0 < P. .10 -4 < 160 and b. 2.

7. ( v 3.

9. (PD 1O -4 ) J06 where D is a diameter of the fan rotor in meters which is measured in a plane normal to an axis (X) of rotation of the fan rotor (9), at an intersection (24) between a tip edge (21) and a leading edge (23) of the blades (14) of the fan rotor (9); P is a maximum thrust generated by the fan (22) in Newton (N) when the propulsion system is stationary in takeoff mode, in a standard atmosphere and at sea level; and BPR is a dilution ratio of the flow passing through the propulsion system (1), defined as a ratio between a mass flow rate of a secondary air flow (F2) flowing in the propulsion system around the gas generator (3), and of a primary air flow (F1) flowing in the gas generator (3), measured when the propulsion system (1) is stationary, uninstalled, in a mode corresponding to a takeoff of an aircraft comprising the aeronautical propulsion system, in a standard atmosphere and at sea level.

2. Aeronautical propulsion system according to claim 1, in which the dilution ratio (BPR) is between 10 and 20 inclusive.

3. Aeronautical propulsion system according to claim 1 or 2, in which the gas generator comprises at least one compressor downstream of the fan, an overall compression ratio (OPR) of the propulsion system being between 40 and 60 inclusive, the overall compression ratio being defined as a pressure ratio between a total pressure measured upstream of a combustion chamber (6) of the propulsion system and a pressure at the inlet of the fan rotor (9) measured upstream of a foot of the fan rotor (9).

4. Aeronautical propulsion system according to any one of claims 1 to 3, the blades (1) of the fan being fixed-pitch blades.

5. Aeronautical propulsion system according to any one of claims 1 to 4 comprising a reduction mechanism (19) coupling in rotation: - the drive shaft (10, 11) and - a fan shaft (20) driving the fan rotor (9), the propulsion system being configured to drive the fan shaft (20) at a rotational speed lower than the rotational speed of the drive shaft (10, 11), a reduction ratio of the reduction mechanism (19) being preferably greater than or equal to 2.5 and less than or equal to 11, still preferably greater than or equal to 2.7 and less than or equal to 6.0, still preferably greater than or equal to 2.7 and less than or equal to 3.6, still preferably substantially 3.

0.

6. Aeronautical propulsion system according to any one of claims 1 to 5, in which the diameter (D) of the fan rotor (9) is greater than or equal to 1.270 m and less than or equal to 3.048 m, preferably greater than or equal to 2.159 m and less than or equal to 3.048 m, still preferably substantially equal to 2.286 m.

7. Aeronautical propulsion system according to any one of claims 1 to 6, in which a hub-head ratio of the fan rotor (9) is greater than or equal to 0.22 and less than or equal to 0.32, preferably greater than or equal to 0.235 and less than or equal to 0.30, more preferably less than or equal to 0.27, the hub-head ratio being defined as by a ratio Rt / R e , Or : - R, is an internal radius of the fan rotor (9), measured between the rotation axis X and a point of intersection between the leading edge (23) of the fan blades (14) and the aerodynamic surface of a fan rotor platform (9) radially delimiting inside the flow vein at the inlet of the fan rotor (9); - R e is an external radius of the fan rotor (9) and is equal to half of the fan diameter (D).

8. Aeronautical propulsion system according to any one of claims 1 to 7, in which the fan rotor (9) comprises at least fourteen blades (14) and at most twenty-four blades, preferably at least sixteen blades and at most twenty-four blades, still preferably twenty-two blades.

9. Aeronautical propulsion system according to any one of claims 1 to 8, in which the gas generator (3) comprises a high pressure body and a low pressure body, the low pressure body rotating at a lower rotational speed than the high pressure body.

10. Aeronautical propulsion system according to claim 9, in which the low pressure body comprises a low pressure turbine (8), the low pressure turbine (8) having at least three stages and at most eight stages.

11. Aeronautical propulsion system according to one of claims 9 or 10, in which the low pressure body comprises a low pressure compressor (4), the low pressure compressor (4) having at least two stages and at most five stages.

12. Aeronautical propulsion system according to any one of claims 9 to 11, in which the high pressure body comprises a high pressure turbine (7), the high pressure turbine (7) being two-stage.

13. Aeronautical propulsion system according to any one of claims 9 to 12, in which the high pressure body comprises a high pressure compressor (5), the high pressure compressor (5) comprising at least eight and at most eleven stages.

14. Aircraft comprising an aeronautical propulsion system according to any one of claims 1 to 13.

15. Method of manufacturing an aeronautical propulsion system, the propulsion system comprising: - a shrouded fan section (2) comprising a fan (22), the fan having a fan rotor (9) comprising blades (14), - a gas generator (3) comprising a drive shaft (10, 11) configured to drive, directly or indirectly, the fan rotor (9) in rotation, the manufacturing method comprising a step of dimensioning the fan rotor (9) during which a diameter D of the fan rotor (9) in meters and measured in a plane normal to an axis (X) of rotation of the fan rotor (9), at an intersection (24) between an end edge (21) and a leading edge (23) of the blades (14) of the fan rotor (9) is chosen so that: a. 0 < PD10 -4 < 160 and b. 2.

7. ( v 3.

9. (PD ÎO -4 ) J06where P is a maximum thrust generated by the fan (22) in Newton (N) when the propulsion system is stationary in takeoff mode, in a standard atmosphere and at sea level; and BPR is a dilution ratio of the flow passing through the propulsion system, defined as a ratio between a mass flow rate of a secondary air flow (F2) flowing in the propulsion system around the gas generator (3), and of a primary air flow (F1) flowing in the gas generator (3), measured when the propulsion system (1) is stationary, uninstalled, in a mode corresponding to a takeoff of an aircraft comprising the aeronautical propulsion system, in a standard atmosphere and at sea level.

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