Aeronautical propulsion system with improved propulsion efficiency

Abstract

The invention relates to an aeronautical propulsion system (1) comprising: - a fan shaft (20); - a fan section (2) comprising a fan rotor (22) rotated by the fan shaft; - a primary spool (3) comprising a compressor section (29) comprising a low-pressure compressor (4) and a high-pressure compressor (5), a combustion chamber (6) and a turbine section (30) comprising a high-pressure turbine (7) rotating the high-pressure compressor (5) by means of a high-pressure shaft (10), and a low-pressure turbine (8) rotating the low-pressure compressor (4) by means of a low-pressure shaft (11); - a reduction mechanism (19) coupling the low-pressure shaft (8) and the fan shaft (20) in order to drive the fan shaft (20) at a speed lower than the speed of rotation of the low-pressure shaft (8); the low-pressure compressor (4) comprising a first stage (44) and at least one other stage (46, 48), each stage comprising, successively in the direction of flow of the gas, a rotor (44r, 46r, 48r) and a stator (44s, 46s, 48s), the low-pressure compressor (4) having a maximum speed having a value ranging from 10700 revolutions per minute (rpm) to 13000 revolutions per minute (rpm), and a value of the ratio between a mean radius (Rmoy) of the low-pressure compressor and an internal radius of the rotor (Ric) of the first stage ranging from 1.1 to 1.4.

Description

[0001] TITLE: AERONAUTICAL PROPULSIVE SYSTEM WITH IMPROVED PROPULSIVE EFFICIENCY

[0002] FIELD OF THE INVENTION

[0003] The present invention relates generally to the field of propulsion systems, and more particularly to aeronautical propulsion systems comprising a ducted or non-ducted fan and having a high, or even very high, dilution ratio.

[0004] TECHNOLOGICAL BACKGROUND

[0005] 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.

[0006] 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.

[0007] 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.

[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, 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 rotational 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 rotational speed lower than that of the low-pressure shaft.However, in such a system, the overall performance of the low-pressure compressor may be limited due to dimensional constraints that may result in particular from the mounting space of bearings located in an enclosure between the low-pressure shaft and the moving wheels of the low-pressure compressor. The volumes occupied in this enclosure may depend in particular on the choice of the rotation speed of said compressor and the power that is chosen to be obtained via the compressors, the combustion chamber and the turbines. The size of the reduction mechanism may also influence the dimensional constraints acting on the efficiency of the low-pressure compressor inside the vein and on the upstream side of the enclosure.In addition, the dimensional constraints will also depend on the space available to integrate the compressor into its environment, then considering the volume of the air stream in the low pressure compressor, this stream volume being between an inner stream wall and an outer stream wall. The performance of the low pressure compressor thus depends on both the aerodynamics and the space around the compressor.

[0009] STATEMENT OF THE INVENTION

[0010] An aim of the present invention is to optimize the performance of the propulsion system and in particular to improve the efficiency of the low pressure compressor while maintaining a high dilution ratio.

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

[0012] - a fan shaft;

[0013] - a fan section comprising a fan rotor rotated by the fan shaft;

[0014] - a primary body comprising a compressor section comprising a low pressure compressor and a high pressure compressor, a combustion chamber and a turbine section comprising a high pressure turbine rotating the high pressure compressor by means of a high pressure shaft and a low pressure turbine rotating the low pressure compressor by means of a low pressure shaft;

[0015] - a reduction mechanism coupling the low pressure shaft and the fan shaft in order to drive the fan shaft at a speed lower than the rotational speed of the low pressure shaft; the low pressure compressor comprising a first stage and at least one other stage, each stage successively comprising in the direction of gas flow a rotor and a stator, the low pressure compressor having a maximum speed having a value ranging from 10700 revolutions per minute (rpm) to 13000 revolutions per minute (rpm), and a value of the ratio between an average radius Rmoy of the low pressure compressor and an internal radius of the rotor of the first stage ranging from 1.1 to 1.A.

[0016] The low-pressure compressor thus has optimized performance, and in particular efficiency. The range mentioned above concerning the value of the ratio between a mean radius Rmoy of the low-pressure compressor and an internal radius of the first-stage rotor includes high-speed low-pressure compressors. The possibility of using high-speed low-pressure compressors makes it possible to obtain a higher low-pressure compression ratio than that of prior-art propulsion systems without increasing the size of the compressor. The overall compression ratio on the compressor section and therefore the thermal efficiency of the propulsion system are thus improved.

[0017] The mean radius (Rmoy) of the compressor is defined as the average of the radii of each rotor, the radius corresponding to the distance between the axis of rotation of the propulsion system and a point taken at mid-height and mid-distance between a leading edge and a trailing edge of the rotor. The internal radius Rie corresponds to the radius of the first rotor, i.e. the most upstream rotor, the internal radius being the distance between the axis of rotation of the propulsion system and a point taken at the blade root at the leading edge of the first-stage rotor.

[0018] In particular, in order for the low pressure compressor (or LP compressor) to have an optimized efficiency, its load or aerodynamic load may preferably be less than or equal to 0.33, the compressor load being defined by the following formula: in which

[0019] AH represents the difference in air enthalpy between the outlet and the inlet of the LP compressor and is expressed in Joules per kilogram (J / Kg), N1 represents the number of stages of the LP compressor and Umoy is the average tangential velocity of the air flow at 50% of the height of the LP compressor rotors and is expressed in meters per second (m / s), Umoy can be calculated according to the formula below

[0020] Tl

[0021] Umoy = Rmoy x Rcomp x (— ) where Rmoy is the average radius of the rotors at 50% of the height and Rcomp is the compressor speed in revolutions per minute (rpm). The aerodynamic load (or charge) of the compressor corresponds to the aerodynamic load of the compressor rotor stages and corresponds to a ratio between a requested work (air compression) and the characteristics / operating conditions of the compressor. With equivalent geometry and identical operating conditions, the higher the value of the load, the higher the compression will be but accompanied by a lower efficiency and more difficult compressor operability. This is why it is preferable for the load of a compressor to have a value not exceeding 0.33.

[0022] The difference in air enthalpy between the outlet and the inlet of the low pressure compressor corresponds to the difference in enthalpy between the leading edge of the first rotor and the trailing edge of the rectifier (or stator) positioned downstream of the last rotor.

[0023] The tangential airflow velocity (Uavg) is defined as the average of the tangential velocities at 50% height for each stage of the low pressure compressor.

[0024] Furthermore, the Mach number at the foot of the leading edge of the rotor of the first stage of the LP compressor can preferably have a value of 1 ± 10% so as to avoid transonic shock waves.

[0025] Some preferred but non-limiting features of the propulsion system according to the first aspect are the following, taken individually or in combination:

[0026] - the fan rotor has a diameter (D) having a value ranging from 1.270 meters (m) to 4.300 m; the diameter is measured at the leading edge of each blade of the fan rotor; this diameter allows the propulsion system to be integrated into a conventional aircraft without having to modify said aircraft, in particular without having to modify the dimensions of the landing gear or the wings and without having to reinforce the wings;

[0027] - the fan section is not shrouded and the fan rotor has a diameter (D) ranging from 1.980 m to 4.300 m; the diameter is measured at the leading edge of each propeller blade;

[0028] - the fan section is shrouded and the fan rotor has a diameter having a value ranging from 1.270 m to 3.048 m; the diameter is measured at the leading edge of each propeller blade; the diameter may for example have a value of 2.286 m;

[0029] - the low pressure compressor is a three-stage compressor and has a compression ratio defined as being a ratio between a gas pressure at the outlet of the low pressure compressor and a gas pressure at the inlet of the low pressure compressor, said compression ratio having a value ranging from 2.8 to 4.8; such a compression ratio of the LP compressor makes it possible to obtain a compression ratio per stage ranging from 1.4 to 1.7 so that the compressor operates with optimized efficiency; the compression ratio per stage being defined as being the compression ratio of the compressor at the power (1 / n); the compression ratio is in particular measured by combs in test engines, in particular by measuring the flow speed and the pressure in climb conditions at 35,000 feet, at Mach 0.8 and in ISÀ+15;

[0030] - the low pressure compressor is a two-stage compressor and has a compression ratio defined as a ratio between a gas pressure at the outlet of the low pressure compressor and a gas pressure at the inlet of the low pressure compressor, having a value ranging from 1.7 to 2.5; such a compression ratio of the LP compressor makes it possible to obtain a compression ratio per stage ranging from 1.4 to 1.7 so that the compressor operates with optimized efficiency;

[0031] - the propulsion system has a bypass ratio defined as a ratio between a mass flow rate of secondary airflow (F2) flowing through the fan section around the primary body and a mass flow rate of primary airflow (F1) flowing through the primary body, having a value ranging from 12 to 16; this value preferably applies to a ducted propeller;

[0032] - the internal radius of the first stage rotor has a value ranging from 250 millimeters (mm) to 350 mm; this radius value range is particularly valid for engines with a SMR (Short Medium Range) thrust range, ranging in particular from 100,000 to 175,000 Newton, these values ​​being measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level;

[0033] - the reduction mechanism has a reduction ratio having a value ranging from 2.5 to 11, preferably ranging from 2.7 to 6.0 and even more preferably ranging from 2.7 to 3.6; the reduction ratio may for example have a value of 3.0;

[0034] - the 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, has a value ranging from 270 m / s to 380 m / s; the fan section is preferably ducted;

[0035] - a hub-to-head ratio of the fan rotor, defined as a ratio between an inner radius (Ri) of the fan rotor and an outer radius (Re) of the fan rotor, has a value ranging from 0.22 to 0.32, preferably ranging from 0.235 to 0.300, and for example approximately 0.27;

[0036] - a ratio between the value of a first chord measured at the top of the fan rotor blade and the value of a second chord measured at the root of the fan rotor blade has a value ranging from 0.56 to 0.80;

[0037] - the fan rotor has a solidity at the tip of the fan rotor blade having a value ranging from 1.0 to 1.3, the solidity being defined as a ratio between the value of a first chord measured at the tip of the blade and the value of an inter-blade pitch;

[0038] - the fan rotor comprises a number of blades ranging from 16 to 24, and for example equal to 22; this number of fan blades is in particular optimized by taking into account the values ​​of fan diameter, solidity at the tip of the fan rotor blade and peripheral speed at the tip of the fan rotor blades, in particular the values ​​previously described;

[0039] - the fan section comprises a fan stator comprising a number of fixed blades ranging from 36 to 52, preferably ranging from 38 to 44, and for example equal to 40; this number of fan stator blades is in particular optimized by taking into account the values ​​of fan diameter, solidity at the tip of the blade of the fan rotor and peripheral speed at the tip of the blades of the fan rotor, in particular the values ​​previously described;

[0040] - the low pressure turbine comprises a number of stages ranging from 3 to 5; this number of stages ensures an aerodynamic load on the low pressure turbine which is optimized and which avoids excessive mass and / or bulk of the turbine;

[0041] - the high pressure turbine comprises two stages;

[0042] - the high pressure compressor includes a number of stages ranging from 8 to 11 stages;

[0043] - the low pressure compressor has a power density with a value ranging from 0.64 to 0.97, the power density of the compressor being defined by the following formula: and where: BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level; n et is the number of compressor stages; Um is the mean peripheral speed of the compressor rotor measured when the propulsion system is stationary in takeoff regime in a standard atmosphere and at sea level and is expressed in pounds of thrust and is expressed in meters per second (m / s); and Rext_moy is the mean external radius of the compressor, which is equal to an arithmetic mean of the external radii of the rotors of the low pressure compressor, the external radius corresponding to a distance between a rotor apex and the axis of rotation, midway at the apex between a leading edge and a trailing edge of the rotor, and is expressed in meters (m).

[0044] Generally speaking, the number of stages described for high and low pressure turbines as well as for high and low pressure compressors allows the turbines and compressors to have enough stages to fulfill their function and that this number of stages is not too large so that the turbines or compressors are not too heavy or too bulky. The dimensions of the blades and the number of blades are also important parameters for the same reasons.

[0045] Also provided, according to a second aspect, is an aircraft comprising a propulsion system according to the first aspect.

[0046] There is also provided, according to a third aspect, a method of dimensioning or a method of manufacturing a propulsion system according to the first aspect, said method comprising the following steps:

[0047] - choose, for the low pressure compressor, a maximum speed ranging from 10700 revolutions per minute (rpm) to 13000 revolutions per minute (rpm), and

[0048] - choose an average radius (Rmoy) of the low pressure compressor and an internal radius of the rotor (Rie) so that a value of the ratio between the average radius (Rmoy) of the low pressure compressor and the internal radius of the rotor (Rie) of the first stage is within a range from 1.1 to 1.4.

[0049] The sizing method also applies to the propulsion system comprising the preferred features previously described.

[0050] The sizing method may include in particular a step according to which a power density having a value ranging from 0.64 to 0.97 is chosen for the low pressure compressor (4), the power density of the compressor being defined by the following formula: and where: BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level; n e t is the number of compressor stages; Um is the mean peripheral speed of the compressor rotor measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in meters per second (m / s); and Rext_moy is the mean external radius of the compressor, which is equal to an arithmetic mean of the external radii of the rotors of the low pressure compressor, the external radius corresponding to a distance between a rotor apex and the axis of rotation, midway at the apex between a leading edge and a trailing edge of the rotor, and is expressed in meters (m).

[0051] BRIEF DESCRIPTION OF THE FIGURES Other characteristics and advantages of the invention will appear on reading the description which follows, given solely by way of example and made with reference to the appended drawings, in which:

[0052] Figure 1 schematically represents an aircraft comprising propulsion systems;

[0053] Figure 2 schematically represents, in partial view and in section, an example of a propulsion system in which the fan section is faired;

[0054] Figure 3 schematically represents, in partial view and in section, an example of a propulsion system in which the fan section is unducted;

[0055] Figure 4 schematically represents a first example of a planetary reduction mechanism with a fixed satellite door;

[0056] Figure 5 schematically represents a first example of an epicyclic reduction mechanism with rotating satellite door;

[0057] Figure 6 schematically represents a partial perspective and top view of a fan rotor; and

[0058] Figure 7 schematically represents, in partial view and in section, an example of a low pressure compressor.

[0059] DETAILED DESCRIPTION OF AN EXAMPLE OF IMPLEMENTATION

[0060] 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.

[0061] Figure 2 schematically represents, in partial view and in section, a first example of a propulsion system 1.

[0062] In this example, propulsion system 1 is a twin-spool, ducted-fan gas turbine engine.

[0063] 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”.

[0064] 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.

[0065] 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, 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 according to the flight phases.

[0066] 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 and at most twenty-four fan blades 14, and for example twenty-two fan blades.

[0067] 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.

[0068] 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.

[0069] 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 14. The number of outlet blades 17 can range from 36 to 52, preferably from 38 to 44, and is for example equal to 40.

[0070] Figure 3 schematically represents, in partial view and in section, a second example of propulsion system 1.

[0071] In Figure 3, components identical or similar to those of the propulsion system of Figure 2 are designated by identical references. In the example illustrated in Figure 3, the propulsion system 1 is a twin-spool, unducted fan gas turbine engine. It may be an “Open Rotor” or “Unducted Single Fan” type gas turbine engine.

[0072] Unlike the first example in Figure 2, the fan rotor 9, which can also be referred to as a "propeller", is not surrounded by a fan casing.

[0073] Since the fan section 2 is not shrouded, the fan blades 14 have variable pitch.

[0074] Alternatively, the propulsion system 1 could comprise two unducted and counter-rotating fan rotors 9. Such a propulsion system 1 is known, in English terminology, by the acronym “CROR” for “Contra-Rotating Open Rotor” or “UDF” for “Unducted Double Fan”. The fan rotors 9 can 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 (“puller” in English).

[0075] The absence of a fairing around the fan rotor 9 makes it possible to increase the bypass ratio significantly without the propulsion system 1 being penalized by the mass of the casings 12 or nacelles intended to surround the fan section 2. The bypass ratio of the propulsion system 1 comprising an unfairly fairinged fan section 2 is thus greater than or equal to 40, for example between 40 and 80 inclusive.

[0076] The following description relates to the first example propulsion system illustrated in Figure 2 but also applies to the second example propulsion system.

[0077] The primary body 3 comprises a compressor section 29, a combustion chamber 6 and a turbine section 30.

[0078] The compressor section 29 comprises a low pressure compressor 4 and a high pressure compressor 5.

[0079] 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.

[0080] 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.

[0081] Similarly, 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. The rotor 51 of the high-pressure compressor 5 comprises movable wheels 5a and the stator 52 of the high-pressure compressor 5 comprises fixed wheels 5b. The movable wheels 5a are arranged alternately with the fixed wheels 5b, thus forming a succession of high-pressure compressor stages.

[0082] The turbine section 30 comprises a high pressure turbine 7 and a low pressure turbine 8.

[0083] 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.

[0084] 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.

[0085] 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.

[0086] 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.

[0087] The propulsion system 1 comprises a low pressure shaft 11 connecting the rotor 41 of the low pressure turbine 4 to the rotor 81 of the low pressure compressor 8, the low pressure shaft 11 being rotatably mounted relative to the casing 31 around the longitudinal axis X.

[0088] 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.

[0089] 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. The reduction mechanism 19 may comprise an epicyclic or planetary, single-stage or two-stage reduction mechanism.

[0090] For example, Figure 4 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.

[0091] In another example, Figure 5 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.

[0092] 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 rotational speed of the rotor 9 of the fan section 2 is lower than the rotational speed of the low pressure shaft 11.

[0093] The reduction ratio of the reduction mechanism 19 has a value ranging from 2.5 to 11, preferably ranging from 2.7 to 6.0, even more preferably ranging from 2.7 to 3.6, and for example a value of approximately 3.0.

[0094] 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.

[0095] The low pressure turbine 8, the low pressure shaft 11, the low pressure compressor 4, the fan shaft 20, the reduction mechanism 19 and the fan 22 together form the “low pressure body” of the propulsion system 1.

[0096] The propulsion system 1 further comprises a high pressure shaft 10 connecting the rotor 51 of the high pressure turbine 5 to the rotor 71 of the high pressure compressor 7, 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.

[0097] When the propulsion system 1 is in operation, the rotor 71 of the high pressure turbine 7 drives the rotor 51 of the low pressure compressor 5 in rotation via the low pressure shaft 11.

[0098] The high-pressure turbine 7, the high-pressure shaft 10 and the high-pressure compressor 4 together form the “high-pressure body” of the propulsion system 1.

[0099] 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.

[0100] The dual-body propulsion system 1 may in particular comprise a single-stage high-pressure turbine 7, i.e. comprising exactly one stage (as illustrated in the example of FIG. 3), or a two-stage high-pressure turbine 7, i.e. comprising exactly two stages (as illustrated in the example of FIG. 2). Preferably, the high-pressure turbine 7 is two-stage.

[0101] The high-pressure compressor 5 comprises at least eight stages (as illustrated in the example of Figure 2) and at most eleven stages.

[0102] The low pressure turbine 8 comprises at least three stages (as illustrated in the example of FIG. 2) and at most seven stages. Preferably, the low pressure turbine 8 comprises at least three stages and at most 5 stages.

[0103] The low pressure compressor 4 comprises at least two stages and at most four stages. Preferably, the low pressure compressor 4 comprises two or three stages.

[0104] 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.

[0105] 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.

[0106] 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.

[0107] 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 ranging from 10 to 80, preferably ranging from 10 to 35, preferably ranging from 12 to 16. 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, these mass flow rates being 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, 3rd edition) and at sea level.By "not installed" it is meant that the measurements are carried out when the propulsion system 1 is on a test bench (and not installed on an aircraft), the measurements then being simpler to carry out.

[0108] 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 (English acronym for International Standard Atmosphere) defined by standard 1502533 / edition 1975 / addendum 1985. In addition, the distances (length, radius, diameter, 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.

[0109] 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 1.05 and 1.30.The average pressures are measured here over the height of at least one of the fan blades 14, i.e. the surface which radially delimits the air flow path at the inlet of the fan rotor 9 at the tip 21 of the fan blade 14.

[0110] 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, 3rd edition) and at sea level.

[0111] The 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, has a value ranging from 270 m / s to 380 m / s.

[0112] The diameter D of the fan rotor may have a value ranging from 1.270 m to 4.300 m. Such a diameter allows the propulsion system to be integrated into an aircraft in a conventional manner, in particular under the wing of the aircraft 1. In the case where the fan section is not ducted, the diameter D may preferably have a value ranging from 1.980 m to 4.300 m. In the case where the fan section is ducted, the diameter D may preferably have a value ranging from 1.270 m to 3.048 m.

[0113] The fan rotor 22 also has a head-to-hub ratio ranging from 0.22 to 0.32, preferably ranging from 0.235 to 0.300, and for example approximately 0.27. The head-to-hub ratio corresponds to the ratio between the internal radius Ri and the external radius Re of the fan rotor 22. The internal radius Ri corresponds to the distance between the axis of rotation X and the point of intersection between the leading edge 14' and the surface which radially delimits the flow path at the inlet of the fan rotor 22 (and corresponds to the point of connection of the leading edge 14' with the aerodynamic surface of a platform of the fan rotor 22). The external radius Re corresponds to the distance between the axis of rotation X and the point of intersection between the leading edge 40 and the tip 42 of the fan blades 14 (and corresponding to half the fan diameter D). The lower the head-hub ratio, the more efficient the fan rotor 22 is.However, the reduction in the hub-to-head ratio of the fan rotor 22 implies an increase in the mechanical load of the hub 13 of the fan rotor 22.

[0114] A ratio between a first chord C1 measured at the fan rotor blade tip and a second chord C2 measured at the fan rotor blade root has a value ranging from 0.56 to 0.80. As illustrated in Figure 6, a chord corresponds to a straight line from the leading edge to the trailing edge in a direction perpendicular to the radial axis of the blade.

[0115] Furthermore, the fan rotor 22 has a blade tip solidity 42 of the fan rotor 22 having a value ranging from 0.9 to 1.3, the solidity being defined as a ratio between the value of the first chord C1 measured at the tip of the blade and the value of an inter-blade pitch P. The inter-blade pitch P corresponds to the distance between two blades measured in blade roots, for example between the leading edge of two successive blades.

[0116] In order to optimize the performance of the propulsion system and in particular the efficiency of the low pressure compressor, the latter has a maximum speed having a value ranging from 10700 rpm to 13000 rpm, and a value of the ratio between an average radius of the low pressure compressor and an internal radius of the rotor of the first stage ranging from 1.1 to 1.4, the first stage of the low pressure compressor being the most upstream stage of the compressor.

[0117] In one embodiment shown in Figure 7, the low pressure compressor comprises a first stage 44 comprising a rotor 44r and a stator 44s, a second stage 46 comprising a rotor 46r and a stator 46s, and a third stage 48 comprising a rotor 48r and a stator 48s.

[0118] The average radius Rmoy of the compressor is defined as the average of the radii of each rotor R44, R46 and R48, the radius corresponding to the distance between, on the one hand, the axis of rotation X of the propulsion system and, on the other hand, a point taken at mid-height and mid-distance between a leading edge and a trailing edge of the rotor blades. The internal radius Rie corresponds to the radius of the first rotor 44r, which is the distance between the axis of rotation X and a point taken at the blade root 44r at mid-distance between a leading edge and a trailing edge of the rotor. This internal radius Rie has a value ranging from 250 mm to 350 mm.

[0119] In particular, in order for the low pressure compressor 4 to have an optimized efficiency, its thermodynamic load or load may preferably be less than or equal to 0.33, the compressor load being defined by the following formula:

[0120] AH

[0121] Compressor load = — — - - in which

[0122] AH represents the difference in air enthalpy between the outlet and the inlet of the LP compressor and is expressed in Joules per kilogram (J / Kg), N1 represents the number of stages of the LP compressor and Umoy is the average tangential speed at 50% of the height of the rotors of the LP compressor and is expressed in meters per second (m / s), Umoy can be calculated according to the formula below where Rmoy corresponds to the previously defined average radius and Rcomp is the compressor speed in revolutions per minute (rpm).

[0123] Furthermore, the Mach number at the foot of the leading edge of the rotor of the first stage of the LP compressor can preferably have a value of 1 ± 10% so as to avoid transonic shock waves.

[0124] When the low pressure compressor 4 has three stages as in Figure 6, the compression ratio per stage has a value ranging from 2.8 to 4.8. Such a compression ratio of the LP compressor makes it possible to obtain a compression ratio per stage ranging from 1.4 to 1.7. According to another embodiment not shown in which the low pressure compressor 4 has two stages, the compression ratio per stage has a value ranging from 1.7 to 2.5. Such a compression ratio of the LP compressor makes it possible to obtain a compression ratio per stage ranging from 1.4 to 1.7.

[0125] The compressor thus operates with optimized efficiency regardless of its number of stages.

[0126] Furthermore, the low pressure compressor 4 has a power density having a value ranging from 0.64 to 0.97, the power density of the compressor being defined by the following formula: and where: BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level and is expressed in pounds of thrust; n et is the number of compressor stages; Um is the mean peripheral speed of the compressor rotor measured when the propulsion system is stationary in takeoff regime in a standard atmosphere and at sea level and is expressed in pounds of thrust and is expressed in meters per second (m / s); and Rext_moy is the mean external radius of the compressor, which is equal to an arithmetic mean of the external radii of the rotors of the low pressure compressor, the external radius corresponding to a distance between a rotor apex and the axis of rotation, midway at the apex between a leading edge and a trailing edge of the rotor, and is expressed in meters (m).

[0127] The invention also includes a method for dimensioning the propulsion system previously, comprising in particular:

[0128] - a fan shaft;

[0129] - a fan section comprising a fan rotor rotated by the fan shaft;

[0130] - a primary body comprising a compressor section comprising a low pressure compressor and a high pressure compressor, a combustion chamber and a turbine section comprising a high pressure turbine rotating the high pressure compressor by means of a high pressure shaft and a low pressure turbine rotating the low pressure compressor by means of a low pressure shaft;

[0131] - a reduction mechanism coupling the low pressure shaft and the fan shaft in order to drive the fan shaft at a speed lower than the rotational speed of the low pressure shaft; the low pressure compressor comprising a first stage and at least one other stage, each stage successively comprising in the direction of gas flow a rotor and a stator, said sizing method comprising the following steps:

[0132] - choose, for the low pressure compressor, a maximum speed ranging from 10700 revolutions per minute (rpm) to 13000 revolutions per minute (rpm), and

[0133] - choose an average radius (Rmoy) of the low pressure compressor and an internal radius of the rotor (Rie) so that a value of the ratio between the average radius (Rmoy) of the low pressure compressor and the internal radius of the rotor (Rie) of the first stage is within a range from 1.1 to 1.4.

[0134] The sizing method applies to the propulsion system including the additional characteristics previously described which can be taken individually or in combination.

[0135] The sizing method may in particular comprise a step according to which a power density having a value ranging from 0.64 to 0.97 is chosen for the low pressure compressor (4), the power density of the compressor being defined by the following formula: and where: BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level; n et is the number of compressor stages; Um is the mean peripheral speed of the compressor rotor measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in meters per second (m / s); and Rext_moy is the mean external radius of the compressor, which is equal to an arithmetic mean of the external radii of the rotors of the low pressure compressor, the external radius corresponding to a distance between a rotor apex and the axis of rotation, midway at the apex between a leading edge and a trailing edge of the rotor, and is expressed in meters (m).

[0136] EXAMPLE

[0137] The advantages of the invention are shown in the comparative example of two propulsion systems whose characteristics are presented in the table below. A first engine (engine 1) is a reference engine not having all the characteristics of the propulsion system according to the invention. A second engine (engine 2) has the characteristics of the propulsion system according to the invention. Engines 1 and 2 are double-spool gas turbine engines comprising a reduction gear.

[0138] In the table below:

[0139] - T / O comes from the Anglicism “take-off” and the T / O thrust thus corresponds to the thrust during the take-off phase;

[0140] BPR comes from the Anglicism “by-pass ratio” and corresponds to the dilution rate of the propulsion system;

[0141] OPR comes from the Anglicism “Overall pressure ratio” and corresponds to the overall compression ratio of the compressor;

[0142] GR comes from the Anglicism “Gear Ratio” and corresponds to the reduction ratio;

[0143] BP stands for “low pressure”;

[0144] HP means “high pressure”;

[0145] CBP stands for “low pressure compressor”;

[0146] CH P stands for “high pressure compressor”;

[0147] - TBP means “low pressure turbine”;

[0148] - THP means “high pressure turbine”;

[0149] N2S corresponds to the following product: (body regime of the stage considered) 2 x surface area of ​​the floor considered;

[0150] RHP1 corresponds to the first stage of the high pressure turbine;

[0151] RHP2 corresponds to the second stage of the high pressure turbine;

[0152] RBP1 corresponds to the first stage of the low pressure turbine;

[0153] RBP corresponds to a stage of the low pressure turbine;

[0154] EGT comes from the anglicism “Exhaust Gas Temperature”; and

[0155] - T41 is the HP turbine wheel inlet temperature.

[0156] Engine 1 has a low-pressure compressor with a relatively low compression ratio of 2.01 and therefore low peripheral speeds. Thus, in order to obtain a good overall compression ratio on the compressor section, the compression ratio of the high-pressure compressor must be of the order of 20. The disadvantage of such an engine is that if one wishes to increase the overall compression ratio, it is necessary either to add a stage to the high-pressure compressor or to increase its average radius and therefore to increase the overall size of the high-pressure compressor. Advantageously, engine 2 has a low-pressure compressor with a compression ratio of 3.01 and therefore higher peripheral speeds than that of engine 1, in order to respect the aerodynamic load of the compressor while maintaining a size close to that of engine 1.It is thus possible to reduce the compression ratio of the high pressure compressor and therefore facilitate its aerodynamic design while increasing the overall compression ratio on the compressor section and thus improving the thermal efficiency of the engine.

Claims

CLAIMS 1. Aeronautical propulsion system (1) comprising: - a fan shaft (20); - a fan section (2) comprising a fan rotor (22) driven in rotation by the fan shaft; - a primary body (3) comprising a compressor section (29) comprising a low pressure compressor (4) and a high pressure compressor (5), a combustion chamber (6) and a turbine section (30) comprising a high pressure turbine (7) rotating the high pressure compressor (5) by means of a high pressure shaft (10) and a low pressure turbine (8) rotating the low pressure compressor (4) by means of a low pressure shaft (11); - a reduction mechanism (19) coupling the low pressure shaft (8) and the fan shaft (20) in order to drive the fan shaft (20) at a speed lower than the rotational speed of the low pressure shaft (8); the low pressure compressor (4) comprising a first stage (44) and at least one other stage (46, 48), each stage successively comprising in the direction of gas flow a rotor (44r, 46r, 48r) and a stator (44s, 46s, 48s), the low pressure compressor (4) having a maximum speed having a value ranging from 10700 revolutions per minute (rpm) to 13000 revolutions per minute (rpm), and a value of the ratio between an average radius (Rmoy) of the low pressure compressor and an internal radius of the rotor (Rie) of the first stage ranging from 1.1 to 1.

4.

2. Propulsion system according to claim 1, in which the fan rotor (22) has a diameter (D) having a value ranging from 1.270 meters to 4.300 meters.

3. A propulsion system according to claim 2, wherein the fan section (20) is not ducted and the fan rotor (22) has a diameter having a value ranging from 1.980 meters to 4.300 meters.

4. A propulsion system according to claim 2, wherein the fan section (20) is ducted and the fan rotor (22) has a diameter having a value ranging from 1.270 meters to 3.048 meters.

5. Propulsion system according to any one of claims 1 to 4, in which the low pressure compressor (4) is a three-stage compressor (44, 46, 48) and has a compression ratio defined as a ratio between a pressure of gas at the outlet of the low pressure compressor and a gas pressure at the inlet of the low pressure compressor, having a value ranging from 2.8 to 4.

8.

6. Propulsion system according to any one of claims 1 to 4, in which the low pressure compressor (4) is a two-stage compressor (44, 46) and has a compression ratio defined as a ratio between a gas pressure at the outlet of the low pressure compressor and a gas pressure at the inlet of the low pressure compressor, having a value ranging from 1.7 to 2.

5.

7. A propulsion system according to any preceding claim, having a bypass ratio defined as a ratio between a mass flow rate of secondary airflow (F2) flowing through the fan section around the primary body and a mass flow rate of primary airflow (F1) flowing through the primary body, having a value ranging from 12 to 16.

8. Propulsion system according to any one of the preceding claims, in which the internal radius (Rie) of the rotor (44r) of the first stage (44) has a value ranging from 250 to 350 millimeters.

9. Propulsion system according to any one of the preceding claims, in which the reduction mechanism (19) has a reduction ratio having a value ranging from 2.5 to 11.

10. Propulsion system according to any one of the preceding claims, in which the peripheral speed at the tip of the blades (14) of the fan rotor (22) has a value ranging from 270 meters / second to 380 meters / second.

11. A propulsion system according to any preceding claim, wherein a hub-to-head ratio of the fan rotor (22), defined as a ratio between an inner radius (Ri) of the fan rotor and an outer radius (Re) of the fan rotor, has a value ranging from 0.22 to 0.

32.

12. Propulsion system according to any one of the preceding claims, in which a ratio between a first chord (C1) measured at the top of the blade (14) of the fan rotor (22) and a second chord (C2) measured at the root of the blade (14) of the fan rotor (22) has a value ranging from 0.56 to 0.

80.

13. Propulsion system according to any one of the preceding claims, in which the fan rotor (22) has a blade tip solidity (14) of the fan rotor (22) having a value ranging from 1.0 to 1.3, the solidity being defined as a ratio between the value of a first chord (C1) measured at the tip of the blade (14) and the value of an inter-blade pitch (P).

14. A propulsion system according to any preceding claim, wherein the fan rotor (22) comprises a number of blades ranging from 16 to 24.

15. A propulsion system according to any preceding claim, wherein the fan section (20) comprises a fan stator (16) having a number of fixed blades (17) ranging from 36 to 52.

16. Propulsion system according to any one of the preceding claims, in which the low pressure turbine (8) comprises a number of stages ranging from 3 to 5.

17. Propulsion system according to any one of the preceding claims, in which the high pressure turbine (7) comprises two stages.

18. Propulsion system according to any one of the preceding claims, in which the high pressure compressor (5) comprises a number of stages ranging from 8 to 11 stages.

19. A propulsion system according to any preceding claim, wherein the low pressure compressor (4) has a power density having a value ranging from 0.64 to 0.97, the power density of the compressor being defined by the following formula: and where: BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level; n et is the number of compressor stages; Um is the average peripheral speed of the compressor rotor measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in meters per second (m / s); and Rext_moy is the average external radius of the compressor, which is equal to an arithmetic mean of the external radii of the rotors of the low pressure compressor, the external radius corresponding to a distance between a top of the rotor and the axis of rotation, midway at the apex between a leading edge and a trailing edge of the rotor, and is expressed in meters (m).

20. Aircraft (100) comprising a propulsion system (1) according to any one of the preceding claims.

21. Method for dimensioning a propulsion system (1), said propulsion system comprising: - a fan shaft; - a fan section comprising a fan rotor rotated by the fan shaft; - a primary body comprising a compressor section comprising a low pressure compressor and a high pressure compressor, a combustion chamber and a turbine section comprising a high pressure turbine rotating the high pressure compressor by means of a high pressure shaft and a low pressure turbine rotating the low pressure compressor by means of a low pressure shaft; - a reduction mechanism coupling the low pressure shaft and the fan shaft in order to drive the fan shaft at a speed lower than the rotational speed of the low pressure shaft; the low pressure compressor comprising a first stage and at least one other stage, each stage successively comprising in the direction of gas flow a rotor and a stator, the method comprising the following steps - choose, for the low pressure compressor, a maximum speed ranging from 10700 revolutions per minute (rpm) to 13000 revolutions per minute (rpm), and - choose an average radius (Rmoy) of the low pressure compressor and an internal radius of the rotor (Rie) so that a value of the ratio between the average radius (Rmoy) of the low pressure compressor and the internal radius of the rotor (Rie) of the first stage is within a range from 1.1 to 1.

4.

22. Method for dimensioning a propulsion system according to any one of claims 2 to 18.

23. Method for dimensioning a propulsion system according to claim 21 or 22, further comprising a step according to which a power density having a value ranging from 0.64 to 0.97 is chosen for the low pressure compressor (4), the power density of the compressor being defined by the following formula: and where: BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level; n e t is the number of compressor stages; Um is the mean peripheral speed of the compressor rotor measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in meters per second (m / s); and Rext_moy is the mean external radius of the compressor, which is equal to an arithmetic mean of the external radii of the rotors of the low pressure compressor, the external radius corresponding to a distance between a rotor tip and the axis of rotation, midway at the tip between a leading edge and a trailing edge of the rotor, and is expressed in meters (m).

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