Aeronautical propulsion system having improved propulsion efficiency

Abstract

An aeronautical propulsion system (1) comprises a high-pressure compressor (5), the mean external radius of which satisfies formula (I), where: Rext_5 is the mean outer radius, in millimeters (mm), of the high-pressure compressor (5); D9 is the diameter, in millimeters (mm), of the fan rotor (9), BPR is the dilution rate of the propulsion system (1); Te is the maximum temperature, in degrees Celsius (°C), at the inlet of the first turbine (8) when the propulsion system (1) is stationary in the takeoff mode; N is the speed, in revolutions per minute (rpm), of the second shaft (10) and is measured when the propulsion system (1) is stationary in the takeoff mode; S7 is a mean surface, in square millimeters (mm2), of the second turbine (7); and A = 1 (tr / min)2.mm / °C and B = 8 977 (rpm)2*mm2.

Description

[0001] DESCRIPTION

[0002] TITLE: Aeronautical propulsion system with improved propulsive efficiency

[0003] TECHNICAL FIELD

[0004] The present application generally relates to the field of propulsion systems, and more particularly to aeronautical propulsion systems comprising a ducted or unducted fan and having a high, or even very high, bypass ratio.

[0005] STATE OF THE ART

[0006] A propulsion system generally comprises, from upstream to downstream in the direction of gas flow, a fan section, a compressor section which may include a low-pressure compressor and a high-pressure compressor, a combustion chamber and a turbine section which may include in particular a high-pressure turbine and a low-pressure turbine. The high-pressure compressor is rotated by the high-pressure turbine via a high-pressure shaft. The fan and, where appropriate, the low-pressure compressor are rotated by the low-pressure turbine via a low-pressure shaft.

[0007] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into consideration the impact factors in all phases of design and development to obtain less energy-intensive, more environmentally friendly aeronautical components and products whose integration and use in civil aviation have moderate environmental consequences with the aim of improving the energy efficiency of aircraft.

[0008] Thus, in order to improve the propulsive efficiency of the propulsion system and to reduce its specific consumption as well as the noise emitted by the fan section, propulsion systems have been proposed having a high BPR (bypass ratio in English, corresponding to the ratio between the flow rate of the secondary air flow and the flow rate of the primary air flow). To achieve such bypass ratios, the fan section can be decoupled from the low-pressure turbine, thus making it possible to independently optimize their respective rotation speeds. Generally, the decoupling is achieved using a reduction mechanism placed between the upstream end of the low-pressure shaft and a rotor of the fan section. The rotor of the fan section is then driven by the low-pressure shaft via the reduction mechanism at a rotation speed lower than that of the low-pressure shaft.

[0009] EXPOSED

[0010] One aim of this application is to optimize the performance of the propulsion system, in particular its size and overall dynamics, without penalizing the efficiency or mechanical performance of the propulsion system.

[0011] For this purpose, according to a first aspect, an aeronautical propulsion system is proposed comprising: - a fan rotor connected to a fan shaft;

[0012] - a first turbine configured to drive the fan rotor (9) via a first shaft;

[0013] - a second turbine configured to drive a second compressor via a second shaft, the second shaft being configured to rotate at a higher speed than the first shaft;

[0014] - a reduction mechanism coupling the first shaft and the fan shaft to drive the fan shaft at a rotational speed lower than the rotational speed of the first shaft; wherein an average external radius of the second compressor complies with the following formula: where: Rext_5 is the average external radius of the second compressor in millimeters;

[0015] Dg is the fan rotor diameter in millimeters, measured in a plane normal to an axis of rotation of the fan rotor at an intersection between a tip and a leading edge of the fan rotor blades;

[0016] BPR is the propulsion system bypass ratio and is measured when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level;

[0017] Te is the maximum temperature at the inlet of the first turbine when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in degrees Celsius (°C);

[0018] N is the rotational speed of the second shaft and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level, in revolutions per minute (rpm);

[0019] S? is an average surface area of ​​the second turbine expressed in square millimeters (mm 2 ) ; and A = 1 (rpm) 2 . mm / °C and B = 8,977 (rpm) 2 *mm 2 .

[0020] Some preferred but non-limiting features of the propulsion system according to the first aspect are the following, taken individually or in combination:

[0021] - the average external radius of the second compressor also respects the following formula: where: C = 7,668 2 *mm 2 ;

[0022] - the average external radius of the second compressor also respects the following formula: ext_s > (E * n5+ F) * R ext _7 where: ns is the number of stages in the second compressor; Rext_7 is the average external radius of the second turbine, in millimeters (mm); and E = - 23.9 * 10 -3 square millimeters (mm 2 ) and F = 0.96;

[0023] - the average external radius of the second compressor also respects the following formula: ext_s < (.E * n5+ G) * R ext _7 where: ns is the number of stages in the second compressor; R ex t_7 is the average external radius of the second turbine, in millimeters (mm); and E = - 23.9 * 10 -3 square millimeters (mm 2 ) and G = 1.06;

[0024] - the second turbine is two-stage and the second compressor is axial; - the second compressor comprises at least eight stages and at most eleven stages;

[0025] - the diameter of the fan rotor is between 2,032 mm and 4,699 mm inclusive, for example between 2,159 mm and 3,048 mm inclusive, for example of the order of 2,286 mm;

[0026] - the fan section is faired and a dilution ratio of the propulsion system is greater than or equal to 10, for example between 10 and 35 inclusive, for example between 10 and 18 inclusive;

[0027] - the fan section is unducted and a bypass ratio of the propulsion system is greater than or equal to 40, for example between 40 and 80 inclusive;

[0028] - a hub-to-head ratio at the inlet of the second compressor is between 0.41 and 0.60;

[0029] - the rotation speed of the second shaft is greater than or equal to 15,000 revolutions per minute and less than or equal to 27,000 revolutions per minute;

[0030] - an overall compression ratio of the propulsion system, corresponding to the ratio between a pressure at the outlet of the second compressor and a pressure at the inlet of the fan rotor, is greater than or equal to 40 and less than or equal to 70;

[0031] - the first turbine comprises at least three and at most five stages;

[0032] - the first turbine further drives a first compressor via the first shaft, the first compressor comprising at least two and at most four stages; and / or

[0033] - the fan section further has a fan pressure ratio, corresponding to a pressure ratio between an outlet of the fan rotor and an inlet of the fan rotor less than or equal to 1.45, for example less than or equal to 1.30.

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

[0035] According to a third aspect, there is provided a method of dimensioning and manufacturing a propulsion system comprising a reduction mechanism coupling a first turbine and a fan rotor to drive the fan rotor at a speed lower than a speed of the first turbine, and a second turbine configured to rotate at a speed higher than the first turbine, wherein an average external radius of the second compressor complies with the following formula: where: Rext_5 is the average external radius of the second compressor in millimeters;

[0036] Dg is the fan rotor diameter in millimeters, measured in a plane normal to an axis of rotation of the fan rotor at an intersection between a tip and a leading edge of the fan rotor blades. Note that since Figures 1 and 2 are partial views, the diameter Dg is only partially visible;

[0037] BPR is the propulsion system bypass ratio and is measured when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level;

[0038] Te is the maximum inlet temperature of the first turbine when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in degrees Celsius (°C); N is the rotational speed of the second shaft and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level, in revolutions per minute (rpm);

[0039] S? is an average surface area of ​​the second turbine expressed in square millimeters (mm 2 ) ; And

[0040] A = 1 (rpm) 2 . mm / °C and B = 8,977 (rpm) 2 *mm 2 .

[0041] Some preferred but non-limiting features of the sizing or manufacturing method according to the first aspect are the following, taken individually or in combination:

[0042] - the average external radius of the second compressor respects the following formula: where: C = 7668 (rpm) 2 *mm 2 ;

[0043] - the average external radius of the second compressor also respects the following formula: ext_s > (E * n5+ F) * R ext _7 where: ns is the number of stages in the second compressor; Rext_7 is the average external radius of the second turbine, in millimeters (mm); and E = - 23.9 * 10 -3 square millimeters (mm 2 ) and F = 0.96; and / or

[0044] - the average external radius of the second compressor also respects the following formula: ext_s < E * n5+ G) * R ext _7 where: ns is the number of stages in the second compressor; Rext_7 is the average external radius of the second turbine, in millimeters (mm); and E = - 23.9 * 10 -3 square millimeters (mm 2 ) and G = 1.06.

[0045] According to a fourth aspect, a method of manufacturing an aeronautical propulsion system is proposed comprising the following steps:

[0046] - dimension the aeronautical propulsion system in accordance with the dimensioning method according to the third aspect; and

[0047] - manufacture the aeronautical propulsion system.

[0048] DESCRIPTION OF FIGURES

[0049] Other characteristics, aims and advantages will emerge from the following description, which is purely illustrative and not limiting, and which must be read in conjunction with the attached drawings in which:

[0050] Figure 1 is a schematic, partial and sectional view of an example of a propulsion system according to a first embodiment, in which the fan section is ducted;

[0051] Figure 2 is a schematic, partial, sectional view of an example of a propulsion system according to a first embodiment, in which the fan section is unducted;

[0052] Figure 3 is a schematic sectional view of an example of a reduction mechanism according to a first variant;

[0053] Figure 4 is a schematic sectional view of an example of a reduction mechanism according to a second variant;

[0054] Figure 5 is an example of an aircraft that may include at least one propulsion system according to the first or second embodiment; Figure 6 is a flowchart illustrating example steps of a sizing or manufacturing method according to one embodiment.

[0055] Throughout the figures, similar elements have identical references.

[0056] DETAILED DESCRIPTION

[0057] A propulsion system 1 has a main direction extending along a longitudinal axis X and comprises, from upstream to downstream in the direction of flow of the gases in the propulsion system 1 when it is in operation, a fan section 2 and a primary body 3, often called a “gas generator”, comprising a compressor section 4, 5, a combustion chamber 6 and a turbine section 7, 8. The propulsion system 1 is here an aeronautical propulsion system 1 configured to be fixed to an aircraft 100 via a pylon (or mast).

[0058] The compressor section 4, 5 comprises a succession of stages each comprising a moving blade wheel (rotor) 4a, 5a rotating in front of a fixed blade wheel (stator) 4b, 5b. The turbine section 7, 8 also comprises a succession of stages each comprising a fixed blade wheel (stator) 7b, 8b behind which a moving blade wheel (rotor) 7a, 8a rotates.

[0059] In the present application, the axial direction corresponds to the direction of the longitudinal axis X, in correspondence with the rotation of the shafts of the gas generator, and a radial direction is a direction perpendicular to this axis X and passing through it. Furthermore, the circumferential (or lateral, or even tangential) direction corresponds to a direction perpendicular to the longitudinal axis X and not passing through it. Unless otherwise specified, internal (respectively, interior) and external (respectively, exterior), respectively, are used in reference to a radial direction so that the internal part or face of an element is closer to the axis X than the external part or face of the same element.

[0060] In operation, an air flow F entering the propulsion system 1 is divided between a primary air flow F1 and a secondary air flow F2, which circulate from upstream to downstream in the propulsion system 1.

[0061] The secondary airflow F2 (also called the "bypass airflow") flows around the primary body 3. The secondary airflow F2 cools the periphery of the primary body 3 and is used to generate most of the thrust provided by the propulsion system 1.

[0062] The primary air flow F1 flows in a primary vein inside the primary body 3, passing successively through the compressor section 4, 5, the combustion chamber 6 where it is mixed with fuel to serve as an oxidant, and the turbine section 7, 8. The passage of the primary air flow F1 through the turbine section 7, 8 receiving energy from the combustion chamber 6 causes rotation of the rotor of the turbine section 7, 8, which in turn drives rotation of the rotor of the compressor section 4, 5 as well as a rotor part 9 of the fan section 2.

[0063] In a twin-body propulsion system 1, the compressor section 4, 5 may comprise a low-pressure compressor 4 and a high-pressure compressor 5. The turbine section 7, 8 may comprise a high-pressure turbine 7 and a low-pressure turbine 8. The rotor of the high-pressure compressor 5 is rotated by the rotor of the high-pressure turbine 7 via a high-pressure shaft 10. The rotor of the low-pressure compressor 4 and the rotor portion 9 of the fan section 2 are rotated by the rotor of the high-pressure turbine 8 via a low-pressure shaft 11. Thus, the primary body 3 comprises a high-pressure body comprising the high-pressure compressor 5, the high-pressure turbine 7 and the high-pressure shaft 10, and a low-pressure body comprising the fan section 2, the low-pressure compressor 4, the high-pressure turbine 8 and the low-pressure shaft 11.The rotational speed of the high-pressure spool is greater than the rotational speed of the low-pressure spool. In a triple-spool propulsion system 1, the turbine section 7, 8 further comprises an intermediate turbine, positioned between the high-pressure turbine 7 and the high-pressure turbine 8 and configured to drive the rotor of the low-pressure compressor 4 via an intermediate shaft. The fan rotor 9 and the rotor of the high-pressure compressor 5 remain driven by the low-pressure shaft 11 and the high-pressure shaft 10, respectively.

[0064] The low pressure shaft 11 is generally housed, over a section of its length, in the high pressure shaft 10 and is coaxial with the high pressure shaft 10. The low pressure shaft 11 and the high pressure shaft 10 may be co-rotating, i.e. driven in the same direction around the longitudinal axis X. Alternatively, the low pressure shaft 11 and the high pressure shaft are counter-rotating, i.e. driven in opposite directions around the longitudinal axis X. Where appropriate, the intermediate shaft is housed between the high pressure shaft 10 and the low pressure shaft 11. The intermediate shaft and the low pressure shaft 11 may be co-rotating or counter-rotating.

[0065] The fan section 2 comprises at least the fan rotor 9 adapted to be driven in rotation relative to a stator part of the propulsion system 1 by the turbine section 7, 8. Each fan rotor 9 comprises a hub 13 and blades 14 extending radially from the hub 13. The blades 14 of each rotor 9 may be fixed relative to the hub 12 or have a variable pitch. In this case, the root of the blades 14 of each rotor 9 is pivotally mounted along a pitch axis and is connected to a pitch change mechanism 15 mounted in the propulsion system 1, the pitch being adjusted according to the flight phases by a pitch change mechanism 15. The pitch change mechanism 15 is illustrated in broken lines in Figure 1 to show that this feature is optional.

[0066] The fan section 2 may further comprise a fan stator 16, or rectifier, which comprises blades 17 mounted on a hub 18 of the fan stator 16 which have the function of straightening the secondary air flow F2 which flows out of the fan rotor 9. The blades 17 of the fan stator 18 may be fixed relative to the hub 18 or have a variable pitch. In a manner similar to the rotor blades 14, the root of the stator blades 17 is pivotally mounted along a pitch axis X and is connected to a pitch change mechanism 15a, which is generally separate from that of the fan rotor 9, the pitch being adjusted according to the flight phases by the pitch change mechanism.

[0067] The fan rotor 9 further comprises at least twelve blades 14 and at most twenty-four blades 14, for example at least sixteen blades 14 and at most twenty-two blades 14. The number of blades 16 in the fan stator 17 depends on the acoustic criteria defined for the propulsion system 1 and is at least equal to the number of blades 14. In order to improve the propulsive efficiency of the propulsion system 1 and to reduce its specific consumption as well as the noise emitted by the fan section 2, the propulsion system 1 has a high bypass ratio. By high bypass ratio, we mean here a bypass ratio greater than or equal to 10, for example between 10 and 80 inclusive.To calculate the bypass ratio, the secondary airflow mass flow rate F2 and the primary airflow mass flow rate F1 are measured when the propulsion system 1 is stationary, uninstalled, in take-off mode in a standard atmosphere (as defined by the International Civil Aviation Organization (ICAO) manual, Doc 7488 / 3, 3. eedition) and at sea level (conditions known as SLS, for Seal Level Standard). It will be noted that, in the present application, the parameters (pressure, flow rate, thrust, speed, etc.) are systematically determined under these conditions. By "not installed", it will be understood here that the measurements are carried out when the propulsion system 1 is in a test bench (and not installed on an aircraft 100), the measurements then being simpler to carry out. The distances (length, radius, diameter, etc.) are, on the other hand, measured at ambient temperature (approximately 20°C) when the propulsion system 1 is cold, that is to say when the propulsion system 1 has been stopped for a sufficient period for the parts of the propulsion system to be at ambient temperature, it being understood that these dimensions vary little compared to the conditions in which the propulsion system 1 is in take-off mode.

[0068] The fan rotor 9 is decoupled from the low-pressure shaft 11 by means of a reduction mechanism 19, placed between an upstream end of the low-pressure shaft 11 and the fan rotor 9, in order to independently optimize their respective rotational speed. In this case, the propulsion system 1 further comprises an additional shaft, called the fan shaft 20. The low-pressure shaft 11 connects the high-pressure turbine 8 to an inlet of the reduction mechanism 19 while the fan shaft 20 connects the outlet of the reduction mechanism 19 to the fan rotor 9. The fan rotor 9 is therefore driven by the low-pressure shaft 11 via the reduction mechanism 19 and the fan shaft 20 at a rotational speed lower than the rotational speed of the high-pressure turbine 8.

[0069] This decoupling makes it possible to reduce the rotational speed and the pressure ratio of the fan rotor 9 and to increase the power extracted by the high-pressure turbine 8. Indeed, the overall efficiency of the propulsion systems is conditioned to the first order by the propulsive efficiency, which is favorably influenced by a minimization of the variation in kinetic energy of the air when passing through the propulsion system 1. In a propulsion system 1 with a high bypass ratio, the majority of the flow generating the propulsive force is constituted by the secondary air flow F2 of the propulsion system 1, the kinetic energy of the secondary air flow F2 being mainly affected by the compression that the secondary air flow F2 undergoes when passing through the fan section 2. The propulsive efficiency and the pressure ratio of the fan section 2 are therefore linked: the lower the pressure ratio of the fan section 2, the better the propulsive efficiency will be.In order to optimize the propulsive efficiency of the propulsion system 1, the pressure ratio of the fan, which corresponds to the ratio between the average pressure at the outlet of the fan stator 17 (or, in the absence of a stator, of the fan rotor 9) and the average pressure at the inlet of the fan rotor 9, is less than or equal to 1.70, for example less than or equal to 1.50, for example between 1.05 and 1.45. The average pressures are measured here over the height of the blade 14 (from the surface which radially delimits the flow path at the inlet of the fan rotor 9 to the tip 21 of the fan blade 14).

[0070] The propulsion system 1 is configured to provide thrust of between 18,000 Ibf (80,068 N) and 51,000 Ibf (22,2411 N), for example between 20,000 Ibf (88,964 N) and 35,000 Ibf (15,5688 N).

[0071] The fan section 2 may be ducted or unducted. In the case of a ducted fan section 2, the fan section 2 comprises a fan casing 12 and the fan rotor 9 is housed in the fan casing 12.

[0072] A shrouded fan section 2 comprises a fan rotor 9 extending upstream of a fan stator. The blades of the fan stator are then generally called outlet vanes (“Outlet Guide Vane” or “OGV” in English) and have a fixed pitch relative to the hub of the fan stator. Furthermore, the peripheral speed at the tip 21 of the blades of the fan rotor 9 may also be between 260 m / s and 400 m / s. The blades 14 of the fan rotor 9 may be fixed or have a variable pitch.

[0073] In an unducted fan section 2, the fan section 2 is not surrounded by a fan casing. Since the fan section 2 is unducted, the blades 14 of the fan rotor 9 have a variable pitch. Propulsion systems comprising at least one unducted fan rotor 9 are known by the terms “open rotor” or “unducted fan”. The propulsion system 1 may comprise two unducted and counter-rotating fan rotors 9. Such a propulsion system 1 is known by the acronym CROR for “Contra-Rotating Open Rotor” or UDF for “Unducted Double Fan”. The fan rotor(s) 9 may be placed at the rear of the primary body 3 so as to be of the pusher type or at the front of the primary body 3 so as to be of the tractor type.Alternatively, the propulsion system 1 may comprise a single unducted fan rotor 9 and an unducted fan stator 16 (straightener). Such a propulsion system 1 is known by the English acronym USF for “Unducted Single Fan”. In the case of a propulsion system 1 of the USF type, the blades 17 of the straightener 16 are fixed in rotation relative to the axis X of rotation of the upstream fan rotor 9 and consequently do not undergo centrifugal force. The blades 17 of the straightener 16 are also variable-pitch.

[0074] The removal of the fairing around the fan section 2 makes it possible to increase the bypass ratio very significantly without the propulsion system 1 being penalized by the mass of the casings or nacelles intended to surround the fan section 2. The peripheral speed at the tip 21 of the blades 14 of the fan rotor(s) 9 can also be between 210 m / s and 260 m / s.

[0075] The reduction mechanism 19 may comprise, for example, a reduction mechanism 19 with an epicyclic gear train, for example of the “epicyclic” type or of the “planetary” type according to the terminology sometimes encountered by those skilled in the art, single-stage or two-stage.According to a first variant, the reduction mechanism 19 may be of the planetary type (“star” in English) (Figure 3) and comprise a sun gear 19a (input of the reduction mechanism 19), centered on an axis X of rotation of the reduction mechanism 19 (generally confused with the longitudinal axis X) and configured to be driven in rotation by the low pressure shaft 11, a ring gear 19b (output of the reduction mechanism 19) coaxial with the sun gear 19a and configured to drive the fan shaft 20 in rotation around the axis X of rotation, and a series of satellites 19c distributed circumferentially around the axis X of rotation between the sun gear 19a and the ring gear 19b, each satellite 19c being meshed internally with the sun gear 19a and externally with the ring gear 19b.The series of satellites 19c is mounted on a planet carrier 19d which is fixed relative to a stator part 19e of the propulsion system 1, for example relative to a casing of the compressor section 4, 5. According to a second variant, the reduction mechanism 19 may be of the epicyclic type (“planetary” in English) (Figure 4), in which case the crown 19b is fixedly mounted on the stator part 19e of the propulsion system 1 and the fan shaft 20 is driven in rotation by the planet carrier 19d (which is therefore movable in rotation relative to a stator part 19e of the propulsion system 1, for example relative to a casing of the compressor section 4, 5).

[0076] Whatever the configuration of the reduction mechanism 19, the diameter of the crown 19b and the planet carrier 19d are greater than the diameter of the sun gear 19a, so that the rotation speed of the fan rotor 9 is lower than the rotation speed of the low pressure shaft 11.

[0077] The reduction ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11. In the case of a propulsion system 1 comprising a shrouded fan rotor 9, the reduction ratio may be greater than or equal to 2.7 and less than or equal to 6.0, for example around 3.0. In the case of a propulsion system 1 comprising an unshrouded fan rotor, the reduction ratio may be between 9.0 and 11.0.

[0078] The redline speed of the low pressure shaft 11, which corresponds to the absolute maximum speed likely to be encountered by the low pressure shaft 11 during the entire flight (according to the European certification regulation EASA CS-E 740 (or according to the American certification regulation 14-CFR Part 33)), is between 8500 rpm and 12000 rpm, for example between 9000 rpm and 11000 rpm. The redline speed corresponds to the maximum rotation speed when the propulsion system is healthy (and potentially at the end of its life). It is therefore likely to be reached by the low pressure shaft 11 in flight conditions. This redline speed is part of the data declared in the engine certification (type certification data sheet).Indeed, this rotation speed is usually used as a reference speed for the dimensioning of propulsion systems 1 and in certain certification tests (such as blade loss or rotor integrity tests).

[0079] In order to optimize the performance of the propulsion system 1, in particular its size and overall dynamics, the average external radius of the high-pressure compressor complies with the following formula: where: Rext_5 is the mean external radius of the high pressure compressor, in millimeters (mm);

[0080] Dg is the diameter of the fan rotor 9 in millimeters (mm), measured in a plane normal to the axis of rotation X at an intersection between a tip 21 and a leading edge 22 of the blades 14 of the fan rotor 9;

[0081] BPR is the bypass ratio of the propulsion system 1;

[0082] Te is the maximum inlet temperature of the low pressure turbine 8, in degrees Celsius (°C)

[0083] N is the rotational speed of the high pressure shaft 10, in revolutions per minute (rpm);

[0084] S? is an average surface area of ​​the high pressure turbine 7, in square millimeters (mm 2 ) ; and A = 1 (rpm) 2 . mm / °C and B = 8,977 (rpm) 2 *mm 2 .

[0085] The average external radius Rext_5 of the high pressure compressor 5 is equal to the arithmetic mean of the external radii R1 of all the rotors (moving blade wheels) of the high pressure compressor 5. In a given stage, the external radius R1 of the rotor corresponds to the distance between the tip 5e of the moving blades of the rotor 5a and the X axis of rotation, halfway (at the tip 5e) between the leading edge 5c and the trailing edge 5d of the moving blades of the rotor 5a (50% of the chord at the tip 5e).

[0086] The mean surface area S7 of the high-pressure turbine 7 is equal to the arithmetic mean of the flow path surfaces at 50% of the chord at the apex 7e of all the rotors 7b (moving blade wheel) of the high-pressure turbine 7. The flow path surface area of ​​a given rotor 7b of the high-pressure turbine 7 is equal to o (R ex t_7 2 - Rintj 2 ), where R ex t_7 and RH_7 correspond respectively to the mean external radius and the mean internal radius of the high pressure turbine 7. The mean external radius R ext_7 (respectively, the average internal radius Rint_7) of the high-pressure turbine 7 is equal to the arithmetic mean of the external radii R3 (respectively, the internal radii R4) of the rotors 7b (moving blade wheels) of the high-pressure turbine 7. In a given stage, the external radius R3 of the rotor corresponds to the distance between the tip 7e of the moving blades of the rotor 7b and the axis X of rotation, halfway (at the tip 7e) between the leading edge 7c and the trailing edge 7d of the moving blades of the rotor 7b (50% of the chord at the tip 7e); the internal radius R4 of the rotor corresponds to the distance between the axis of rotation X and the surface which radially delimits the flow vein in the rotor 7b, in a plane passing through 50% of the chord at the vertex 7e, so that the external radii R3 and internal radii R4 are measured in the same plane.

[0087] The average surface area S7 of the high-pressure turbine can, for example, be between 0.035 m 2 and 0.1 m2 .

[0088] As previously indicated, the BPR dilution ratio, the temperature Te and the rotation speed N are determined in take-off conditions. The average external radius R ex t_5 and diameter D9 are, on the other hand, determined when the propulsion system 1 is cold.

[0089] The inlet temperature of the low pressure turbine Te can be between 950°C and 1230°C. The inlet section of the low pressure turbine can be between 0.08 m 2 and 0.35 m 2 .

[0090] When the average external radius R ex t_5 respects the formula (1), the radial and longitudinal dimensions of the high pressure compressor 5 as well as the mass of the high pressure compressor 5 are reduced, with similar performances. Indeed, the power density of the high pressure compressor 5 is proportional to the ratio between the inlet flow of the low pressure compressor 4 and the temperature T eat the inlet of the low pressure turbine 8, knowing that the inlet flow rate of the low pressure compressor 4 is a function of the bypass ratio BPR of the propulsion system 1 and the diameter Dg of the fan rotor 9 (at iso ratio D / S, where S corresponds to the inlet section of the fan rotor 9). Thus, by sizing the average external radius Rext_5 according to the fan diameter D and the bypass ratio BPR, it is possible to reduce the average external radius of the high pressure compressor 5 while ensuring that the power density of the high pressure compressor 5 is sufficient to obtain an efficient propulsion system 1 (and therefore has good energy performance).

[0091] The average external radius Rext_5 is also dimensioned taking into account the mechanical load of the high pressure turbine 7, which is proportional to the product of the rotational speed of the high pressure shaft 10 squared (N 2) and to the average surface S7 of the high pressure turbine 7.

[0092] The dimensioning of the average external radius R ex t_5 in accordance with formula (1) therefore makes it possible to reduce both the size and the mass of the high-pressure body, without mechanically penalizing the high-pressure turbine 7 or affecting the efficiency of the propulsion system 1. The reduction in the radial size of the high-pressure compressor 5 also makes it easier to integrate the propulsion system 1 into an aircraft 100 and to reduce its drag.

[0093] This reduction in the average external radius (R ext_5) of the high-pressure compressor 5 can in particular be obtained by increasing the bypass ratio BPR of the propulsion system 1 and by reducing the pressure ratio of the fan rotor 9, and therefore by improving the efficiency of the propulsion system 1. For a shrouded fan section 2, the bypass ratio BPR of the propulsion system 1 can then be greater than or equal to 10, for example between 10 and 35 inclusive, for example between 10 and 18 inclusive. The fan pressure ratio can furthermore be between 1.20 and 1.45. For an unshrouded fan section 2, the bypass ratio BPR of the propulsion system 1 can be greater than or equal to 40, for example between 40 and 80 inclusive. The fan pressure ratio can then be between, for example, 1.05 and 1.20.

[0094] Alternatively or in addition, this reduction can be obtained by improving the aerodynamic performance of the high-pressure body, by increasing the admissible Mach at iso-efficiency at iso-aerodynamic load and / or by increasing the aerodynamic load at iso-efficiency (with compensation possibly by reducing the Mach in the vein of the high-pressure compressor 5), in order to increase the power of the high-pressure turbine 7 and its mechanical resistance. For example, reducing the leaks in the high-pressure turbine 7 (by maintaining the clearances) makes it possible to increase the efficiency of the compression carried out in the high-pressure compressor 5 and of the expansion in the high-pressure turbine 7.

[0095] When the average external radius (R ext_5) complies with formula (1), the high-pressure turbine 7 can also rotate at rotational speeds (in revolutions per minute) that are sufficiently high (and at acceptable mechanical loadings) to allow the high-pressure compressor 5 to achieve a high power density and a compression ratio greater than 21 (in take-off mode) without requiring an increase in the number of stages in the high-pressure compressor 5 (and thus limiting its longitudinal size). The rotational speed of the high-pressure shaft 10 can then be between 15,000 revolutions per minute and 27,000 revolutions per minute. The high-pressure turbine 7 is, for example, two-stage and the high-pressure compressor 5 is axial, the high-pressure compressor 5 being able to comprise at least eight stages and at most eleven stages, for example nine stages.

[0096] We also obtain an overall compression ratio of the propulsion system 1, which corresponds to the pressure ratio between the pressure at the outlet of the high pressure compressor 5 and the pressure at the inlet of the fan rotor 9 (measured at the level of the foot of the fan rotor 9), which can be greater than or equal to 40 and less than or equal to 70, for example greater than or equal to 44 and less than or equal to 55.

[0097] For example, the average external radius Rext_5 also respects the following formula:

[0098] QÙ . Q, = 1668 (rpm) 2 *mm 2 .

[0099] When the average external radius Rext_5 also respects formula (2), the high pressure compressor 5 remains large enough to achieve the compression of the air at the inlet of the combustion chamber 6 and thus participate in the overall compression ratio of the propulsion system 1. In addition, the aerodynamic load and the Mach in the flow path remain compatible with the efficiency objectives and the operability of the high pressure compressor 5.

[0100] For a high pressure compressor 5 whose average external radius R ex t_5 respects formula (1) (and where applicable formula (2)), the diameter Dg of the fan rotor can then be between

[0101] 2,032 mm (80 inches) and 4,699 mm (185 inches) inclusive. When the fan rotor 9 is shrouded, the diameter Dg is for example between 2,159 mm (85 inches) and 3,048 mm (120 inches) inclusive, for example of the order of 2,286 mm (90 inches), which allows the propulsion system 1 to be integrated in a conventional manner, in particular under the wing of an aircraft. When the fan rotor 9 is unshrouded, the diameter Dg is for example greater than or equal to 2,540 mm (100 inches), for example between

[0102] 3,048 cm (120 in) and 3,962 mm (156 in).

[0103] Where appropriate, in order to achieve a compromise between the radial compactness of the high pressure compressor 5 and the aerodynamic load of the high pressure compressor 5, the average external radius R ex t_5 also respects the following formula:

[0104] (E * n5+ F) * R ext 7 < R ext 5 < (E * n5+ G) * R ext-7(3) where: ns is the number of stages in the high pressure compressor 5; and

[0105] E = - 23.9 * 10 -3 square millimeters (mm 2 ), F = 0.96 and G = 1.06, F and G being dimensionless.

[0106] At the same number of stages, compliance with formula (3) makes it possible to obtain a high-pressure compressor 5 that is more radially compact without penalizing its efficiency. Indeed, the larger the average external radius R ex t_5 of the high pressure compressor 5 is lower, the more compact the high pressure compressor 5 is. The decrease in the average external radius R ext_5 certainly has the effect of increasing the aerodynamic load of the high pressure compressor 5 (and therefore of reducing its efficiency): however, formula (1) makes it possible to obtain an acceptable compromise between the radial compactness and the efficiency of the high pressure compressor 5, and to gain in mass and specific consumption of the propulsion system 1. In addition, formula (1) taking into account the average external radius Rext_7 of the high pressure turbine 7, the mechanical loading of the high pressure turbine 7 (which is proportional to the product of the rotation speed of the high pressure shaft 10 squared and the average surface area S7 of the high pressure turbine 7) remains acceptable.

[0107] High pressure compressors 5 complying with formula (1) and where appropriate formulas (2) and (3) may have a hub-to-head ratio at the inlet of the high pressure compressor 5, which corresponds to the ratio between the external radius R1 of the high pressure compressor 5 and the internal radius R2 of the rotor 5b of the high pressure compressor 5, between 0.41 and 0.60.

[0108] The outer radius R1 and the inner radius R2 are measured here in a plane normal to the axis of rotation X which intersects the leading edge and the blade root of the rotor 5b furthest upstream of the high-pressure compressor 5 (i.e., of the first stage of the high-pressure compressor 5). The inner radius R2 of the high-pressure compressor 5 corresponds to the distance, in this plane, between the outer radial surface of the hub of the rotor 5b (which radially delimits the flow path in the rotor 5b on the inside) and the axis of rotation X.

[0109] When the high-pressure compressor 5 includes an impeller (a compressor that is entirely or partly centrifugal), the impeller then counts as two stages in the calculation of the number of stages ns of the high-pressure compressor 5.

[0110] Such high-pressure compressors 5 then have an optimized outlet section. Indeed, the smaller the hub-to-head ratio, the smaller the external diameter of the high-pressure compressor 5 (at iso-section). A hub-to-head ratio of between 0.77 and 0.90, in combination with an average external radius Rext_5 and a rotation speed optimized as described above, thus makes it possible to obtain not only a more efficient high-pressure compressor 5 (due to the adapted rotation speed of the high-pressure turbine 7), in a suitable size, while optimizing the surface which compresses the gases at the inlet of the combustion chamber 6.

[0111] A double-body propulsion system 1 with an average external radius Rex t_5 in the intervals defined above may in particular comprise a high pressure turbine 8 comprising at least three stages and at most five stages and a low pressure compressor 4 comprising at least two stages and at most four stages.

[0112] Comparative example 1: streamlined propulsion system

[0113] Engine 1 is a double-body propulsion system comprising a shrouded fan section 2 corresponding to the current technical standard (at the date of filing of this application) which we seek to improve.

[0114] The engine 2 is a double-body propulsion system 1 comprising a shrouded fan section 2 in accordance with the teaching of the present application having an average radius of the high-pressure compressor between the maximum and minimum values ​​defined by the relationships (1), (2) and (3) defined above.

[0115]

[0116] The high pressure compressor of engine 1 has a high external mean radius which does not comply with the formulas (1) and (3) defined above.

[0117] In comparison, the high-pressure compressor 5 of the engine 2 has a moderate external mean radius which complies with formulas (1) to (3). The high-pressure body of the engine 2 is therefore more compact than that of the engine 1, without mechanically loading the high-pressure turbine 7, while ensuring that the aerodynamic load of the high-pressure compressor 5 remains moderate to obtain an efficient propulsion system 1 (and therefore has good energy performance).

[0118] To move from engine 1 (reference) to engine 2 (compliant with the disclosure), the fan diameter and the bypass ratio BPR were increased, which made it possible to improve the propulsive efficiency and to maintain a comparable fan thrust taking into account the reduction in the pressure ratio of the fan section 2. Furthermore, the overall compression ratio was increased, as well as the inlet temperature of the high pressure turbine 7, which made it possible to increase the thermal efficiency of the propulsion system 1.

[0119] Comparative example 2: unducted propulsion system

[0120] Engine 3 is a twin-body propulsion system comprising an unducted fan section 2 corresponding to the current technical standard (at the date of filing of this application) which we seek to improve.

[0121] The engine 4 is a twin-body propulsion system 1 comprising an unducted fan section 2 in accordance with the teaching of the present application having an average radius of the high-pressure compressor between the maximum and minimum values ​​defined by the relationships (1), (2) and (3) defined above.

[0122] The high pressure compressor of engine 3 has a high external mean radius which does not comply with the formulas (1) and (3) defined above.

[0123] In comparison, the high-pressure compressor 5 of the engine 4 has a moderate external mean radius that complies with formulas (1) to (3). The high-pressure body of the engine 4 is therefore more compact than that of the engine 3, while the mechanical loading of the high-pressure turbine 7 is lower, while ensuring that the power density of the high-pressure compressor 5 is sufficient to obtain an efficient propulsion system 1 (and therefore has good energy performance). To move from the engine 3 (reference) to the engine 4 (compliant with the disclosure), the fan diameter Dg and the bypass ratio BPR were reduced, which made it possible to improve the integration of the engine 2. The pressure ratio of the fan section of the engine 2 was, on the other hand, slightly increased (while remaining less than 1.45) to maintain equivalent thrust.In addition, the overall compression ratio was increased without increasing the inlet temperature of the high pressure turbine 7, which improves the efficiency of the primary body without increasing the thermal load of the low pressure turbine 8. Finally, since the temperature of the low pressure turbine 8 was kept stable, it was possible to increase its mechanical loading in order to reduce its number of stages.

[0124] The increase in the power density of the high pressure compressor of the engines 2 and 4 can be obtained as follows: increase in the compression ratio of the high pressure compressor; reduction in the diameter of the low pressure shaft with a supercritical dynamic situation maintained acceptable (first deformation mode in the operating range, in an operating zone in which the dynamic situation of the low pressure shaft 11 can be controlled - typically, for a propulsion system comprising a reduction mechanism 19, the first deformation mode of the low pressure shaft 11 occurs at a transient speed corresponding for example to a speed lower than the idle speed and is therefore not likely to damage the propulsion system 1.Thus, a transient speed is chosen which only intervenes for a very short time in the operation of the propulsion system 1 - the second mode of deformation outside the operating range) via a modification of the internal geometry of the low pressure shaft and an increase in the distance between the bearings downstream of the low pressure turbine; better integration of the upstream bearing of the high pressure body combined with the reduction of the low pressure shaft diameter allows a lowering of the internal vein. This better integration can be obtained for example by using rolling elements made of ceramic material, which allow an increase in the contact pressure and a reduction in thermal rejections.

[0125] Alternatively, instead of reducing the diameter of the low pressure shaft, it is possible to modify the shape of the low pressure shaft to give it a bottle shape: in particular, the radius of the low pressure shaft under the high pressure turbine is modified so as to maintain the dynamic situation of the low pressure shaft and its radius upstream of the combustion chamber is reduced so as to facilitate the integration of the internal vein of the high pressure compressor. The reduction of the aerodynamic load of the high pressure turbine also increases its diameter and facilitates the design of the high pressure turbine disk.

Claims

CLAIMS 1. Aeronautical propulsion system (1) comprising: - a fan rotor (9) connected to a fan shaft (20); - a first turbine (8) configured to drive the fan rotor (9) via a first shaft (11); - a second turbine (7) configured to drive a second compressor (5) via a second shaft (10), the second shaft (10) being configured to rotate at a higher speed than the first shaft (11); - a reduction mechanism (19) coupling the first shaft (11) and the fan shaft (20) in order to drive the fan shaft (20) at a rotational speed lower than the rotational speed of the first shaft (11); wherein an average external radius of the second compressor (5) complies with the following formula: where: Rext_5 is the average external radius of the second compressor (5) in millimeters (mm); Dg is the diameter of the fan rotor (9) in millimeters (mm), measured in a plane normal to an axis of rotation (X) of the fan rotor (9) at an intersection between a tip (21) and a leading edge (22) of the blades (14) of the fan rotor (9); BPR is the bypass ratio of the propulsion system (1) and is measured when the propulsion system (1) is stationary in takeoff mode in a standard atmosphere and at sea level; Te is the maximum temperature at the inlet of the first turbine (8) when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in degrees Celsius (°C); N is the rotational speed of the second shaft (10) and is measured when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level, in revolutions per minute (rpm); S? is an average surface area of ​​the second turbine (7) expressed in square millimeters (mm 2 ) ; And A = 1 (rpm) 2 . mm / °C and B = 8,977 (rpm) 2 *mm 2 .

2. Propulsion system (1) according to claim 1, wherein the average external radius of the second compressor (5) further complies with the following formula: where: C = 7668 (rpm) 2 *mm 2 .

3. Propulsion system (1) according to one of claims 1 and 2, in which the average external radius of the second compressor (5) further complies with the following formula: ext_s > (E * n5+ F) * R ext _7 where: ns is the number of stages in the second compressor (5); Rext_7 is the average external radius of the second turbine (7), in millimeters (mm); and E = - 23.9 * 10 -3 square millimeters (mm 2 ) and F = 0.

96.

4. Propulsion system (1) according to one of claims 1 to 3, in which the average external radius of the second compressor (5) further complies with the following formula: ext_s < E * n5+ G) * R ext _7 where: ns is the number of stages in the second compressor (5); Rext_7 is the average external radius of the second turbine (7), in millimeters (mm); and E = - 23.9 * 10 -3 square millimeters (mm 2 ) and G = 1.

06.

5. Propulsion system (1) according to one of claims 1 to 4, in which the second turbine (7) is two-stage and the second compressor (5) is axial.

6. Propulsion system (1) according to one of claims 1 to 5, in which the second compressor (5) comprises at least eight stages and at most eleven stages.

7. Propulsion system (1) according to one of claims 1 to 6, in which the diameter of the fan rotor (9) is between 2,032 mm (80 inches) and 4,699 mm (185 inches) inclusive, for example between 2,159 mm (85 inches) and 3,048 mm (120 inches) inclusive, for example of the order of 2,286 mm (90 inches).

8. Propulsion system (1) according to one of claims 1 to 7, in which the fan section (2) is faired and a dilution ratio of the propulsion system (1) is greater than or equal to 10, for example between 10 and 35 inclusive, for example between 10 and 18 inclusive.

9. Propulsion system (1) according to one of claims 1 to 7, in which the fan section (2) is unducted and a bypass ratio of the propulsion system (1) is greater than or equal to 40, for example between 40 and 80 inclusive.

10. Propulsion system (1) according to one of claims 1 to 9, in which a hub-head ratio at the inlet of the second compressor (5) is between 0.41 and 0.

60.

11. Propulsion system (1) according to one of claims 1 to 10, wherein the rotational speed of the second shaft (10) is greater than or equal to 15,000 revolutions per minute (rpm) and less than or equal to 27,000 revolutions per minute (rpm).

12. Propulsion system (1) according to one of claims 1 to 11, in which an overall compression ratio of the propulsion system (1), corresponding to the ratio between a pressure at the outlet of the second compressor (5) and a pressure at the inlet of the fan rotor (9), is greater than or equal to 40 and less than or equal to 70.

13. Propulsion system (1) according to one of claims 1 to 12, in which the first turbine (8) comprises at least three and at most five stages.

14. Propulsion system (1) according to one of claims 1 to 13, wherein the first turbine further drives a first compressor (4) via the first shaft (11), the first compressor (4) comprising at least two and at most four stages.

15. Propulsion system (1) according to one of claims 1 to 14, wherein the fan section (2) further has a fan pressure ratio, corresponding to a pressure ratio between an outlet of the fan rotor (9) and an inlet of the fan rotor (9) less than or equal to 1.45, for example less than or equal to 1.

30.

16. Aircraft (100) comprising at least one propulsion system (1) according to one of claims 1 to 15 fixed to the aircraft via a mast.

17. Method for dimensioning a propulsion system (1) comprising a reduction mechanism (19) coupling a first turbine (8) and a fan rotor (9) to drive the fan rotor (9) at a speed lower than a speed of the first turbine (8), and a second turbine (7) configured to rotate at a speed higher than the first turbine (8), wherein an average external radius of the second compressor (5) complies with the following formula: where: Rext_5 is the average external radius of the second compressor (5) in millimeters (mm); Dg is the diameter of the fan rotor (9) in millimeters (m), measured in a plane normal to an axis of rotation of the fan rotor (9) at an intersection between a tip (21) and a leading edge (22) of the blades (14) of the fan rotor (9); BPR is the bypass ratio of the propulsion system (1) and is measured when the propulsion system (1) is stationary in takeoff mode in a standard atmosphere and at sea level; Te is the maximum temperature at the inlet of the first turbine (8) when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and is expressed in degrees Celsius (°C); N is the rotational speed of the second shaft and is measured when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level, in revolutions per minute (rpm); S? is an average surface area of ​​the second turbine (7) expressed in square millimeters (mm 2 ) ; And A = 1 (rpm) 2 . mm / °C and B = 8,977 (rpm) 2 *mm 2 .

18. A sizing method according to claim 17, wherein the average external radius of the second compressor (5) complies with the following formula: where: C = 7668 (rpm) 2 *mm 2 .

19. Sizing method according to one of claims 17 and 18, wherein the average external radius of the second compressor (5) further complies with the following formula: ext_s > (E * n5+ F) * R ext _7 where: ns is the number of stages in the second compressor (5); Rext_7 is the average external radius of the second turbine (7), in millimeters (mm); and E = - 23.9 * 10 -3 square millimeters (mm 2 ) and F = 0.

96.

20. Sizing method according to one of claims 17 to 19, wherein the average external radius of the second compressor (5) further complies with the following formula: ext_s < E * n5+ G) * R ext _7 where: ns is the number of stages in the second compressor (5); Rext_7 is the average external radius of the second turbine (7), in millimeters (mm); and E = - 23.9 * 10 -3 square millimeters (mm 2 ) and G = 1.

06.

21. Method of manufacturing an aeronautical propulsion system comprising the following steps: - dimensioning the aeronautical propulsion system in accordance with the method according to one of claims 17 to 20; and - manufacture the aeronautical propulsion system.

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