Aeronautical propulsion system with improved propulsion efficiency

Abstract

An aeronautical propulsion system (1) has a mean internal radius of the high-pressure turbine (7) at least equal to formula (I), where: Rmoyen'int is the mean internal radius of the high-pressure turbine (7) in millimeters (mm); D is the diameter of the fan rotor (9) in millimeters (mm); BPR is the bypass ratio of the propulsion system (1); Te is the inlet temperature of the first turbine (8) when the propulsion system (1) is stationary at takeoff speed in a standard atmosphere and at sea level and is expressed in degrees Celsius (°C); Tref = 273 K; GAMMA is the adiabatic coefficient of air; and A = 593 (°C)<sp / > - 1 / 2 and B = 6.5 millimeters (mm).

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 bypass ratio (BPR, corresponding to the ratio between the flow rate of the secondary air flow and the flow rate of the primary air flow). To achieve such bypass ratios, the fan section can be decoupled from the low-pressure turbine, thus making it possible to independently optimize their respective rotation speeds. Generally, the decoupling is achieved using a reduction mechanism placed between the upstream end of the low-pressure shaft and a rotor of the fan section. The rotor of the fan section is then driven by the low-pressure shaft via the reduction mechanism at a rotation speed lower than that of the low-pressure shaft.

[0009] The current trend is to increase the overall compression ratio of the propulsion system, which corresponds to the ratio between the pressure at the outlet of the high-pressure compressor and the pressure at the inlet of the fan. This in fact makes it possible to increase the pressure at the inlet of the combustion chamber gases, and therefore to further improve the overall efficiency of the propulsion system. Increasing the overall compression ratio therefore requires increasing the compression ratio of the high-pressure compressor and / or the low-pressure compressor, especially since at the same time we are trying to reduce the compression ratio of the fan for the reasons explained above. One of the consequences is that the high-pressure spool is smaller and the high-pressure turbine is more mechanically loaded, particularly at the blade root.

[0010] EXPOSED

[0011] One aim of this application is to optimize the propulsion system in order to increase its efficiency without mechanically and / or thermally overloading the high-pressure turbine.

[0012] For this purpose, according to a first aspect, an aeronautical propulsion system is proposed comprising:

[0013] - a first turbine configured to drive a first compressor via a first shaft;

[0014] - 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;

[0015] - a fan rotor (9) connected to a fan shaft;

[0016] - a reduction mechanism coupling the first shaft and the fan shaft in order to drive the fan shaft at a rotational speed lower than the rotational speed of the first shaft; wherein an internal mean radius of the second turbine is at least equal to:

[0017] Radius moyen int > A * ' CAMMA ~ * 10 3 - B where: Rmoyenjnt is the internal mean radius of the second turbine in millimeters;

[0018] D is the diameter of the fan rotor in millimeters, measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of the fan rotor blades;

[0019] 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;

[0020] Te is the 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);

[0021] Tref = 273 K;

[0022] GAMMA is the adiabatic coefficient of air; and

[0023] A = 593 (°C) -1 / 2 and B = 6.5 millimeters (mm).

[0024] Some preferred but non-limiting features of the propulsion system according to the first aspect are the following, taken individually or in combination:

[0025] - the internal average radius of the second turbine is at most equal to 300 mm;

[0026] - a hub-to-head ratio of the second turbine is greater than 0.77 and less than 0.90;

[0027] - the propulsion system further comprises an inter-turbine casing mounted on a bearing assembly, the inter-turbine casing extending between the first turbine and the second turbine;

[0028] - the inter-turbine casing comprises a series of guide vanes configured to straighten an air flow at the inlet of the first turbine; the propulsion system comprises between twenty and thirty guide vanes;

[0029] - the second turbine is two-stage;

[0030] - the second compressor comprises at least eight stages and at most eleven stages;

[0031] - the diameter of the fan rotor is between 80 inches and 185 inches inclusive, preferably between 85 inches and 120 inches inclusive, for example of the order of 90 inches;

[0032] - the fan rotor is shrouded and a bypass ratio of the propulsion system is greater than or equal to 10, for example between 10 and 35 inclusive, preferably between 10 and 18 inclusive;

[0033] - the fan rotor is unducted and a propulsion system bypass ratio is greater than or equal to 40, for example between 40 and 80 inclusive;

[0034] - the reduction mechanism has a reduction ratio greater than or equal to 2.5, preferably greater than or equal to 3.0 and less than or equal to 11.0;

[0035] - the first turbine comprises at least three stages and at most five stages; and / or

[0036] - the first compressor comprises at least two stages and at most four stages.

[0037] According to a second aspect, the present application proposes an aircraft comprising at least one propulsion system according to the first aspect fixed to the aircraft by means of a mast.

[0038] According to a third aspect, the present application provides a method for dimensioning or 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 higher speed than the first turbine, the second turbine being dimensioned such that an internal mean radius of the second turbine is at least equal to: where: Rmoyenjnt is the internal mean radius of the second turbine in millimeters;

[0039] D is the diameter of the fan rotor in millimeters, measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of the fan rotor blades;

[0040] 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;

[0041] Te is the 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);

[0042] Tref = 273 K;

[0043] GAMMA is the adiabatic coefficient of air; and

[0044] A = 593 (°C) -1 / 2 and B = 6.5 millimeters (mm). DESCRIPTION OF FIGURES

[0045] Other characteristics, aims and advantages of the invention will emerge from the following description, which is purely illustrative and non-limiting, and which must be read in conjunction with the appended drawings in which:

[0046] 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;

[0047] 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;

[0048] Figure 3 is a schematic sectional view of an exemplary planetary reduction mechanism;

[0049] Figure 4 is a schematic sectional view of an exemplary epicyclic reduction mechanism;

[0050] Figure 5 is an example of an aircraft that may include at least one propulsion system according to the first or second embodiment; and

[0051] Figure 6 is a flowchart illustrating exemplary steps of a sizing or manufacturing method according to one embodiment.

[0052] Throughout the figures, similar elements have identical references.

[0053] DETAILED DESCRIPTION

[0054] 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).

[0055] 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.

[0056] 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 with 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. In operation, an air flow F entering the propulsion system 1 is divided between a primary air flow F1 and a secondary air flow F2, which circulate from upstream to downstream in the propulsion system 1.

[0057] The secondary airflow F2 (also called the "bypass airflow") flows around the primary body 3. The secondary airflow F2 cools the periphery of the primary body 3 and is used to generate most of the thrust provided by the propulsion system 1.

[0058] The primary air flow F1 flows in a primary vein inside the primary body 3, passing successively through the compressor section 4, 5, the combustion chamber 6 where it is mixed with fuel to serve as an oxidant, and the turbine section 7, 8. The passage of the primary air flow F1 through the turbine section 7, 8 receiving energy from the combustion chamber 6 causes rotation of the rotor of the turbine section 7, 8, which in turn drives rotation of the rotor of the compressor section 4, 5 as well as a rotor part 9 of the fan section 2.

[0059] In a twin-spool propulsion system 1, the compressor section 4, 5 may comprise a low-pressure compressor 4 and a high-pressure compressor 5. The turbine section 7, 8 may comprise a high-pressure turbine 7 and a low-pressure turbine 8. The rotor of the high-pressure compressor 5 is rotated by the rotor of the high-pressure turbine 7 via a high-pressure shaft 10. The rotor of the low-pressure compressor 4 and the rotor portion 9 of the fan section 2 are rotated by the rotor of the low-pressure turbine 8 via a low-pressure shaft 11. Thus, the primary body 3 comprises a high-pressure body comprising the high-pressure compressor 5, the high-pressure turbine 7 and the high-pressure shaft 10, and a low-pressure body comprising the fan section 2, the low-pressure compressor 4, the low-pressure turbine 8 and the low-pressure shaft 11.The rotational speed of the high-pressure spool is greater than the rotational speed of the low-pressure spool. In a triple-spool propulsion system 1, the turbine section 7, 8 further comprises an intermediate turbine, positioned between the high-pressure turbine 7 and the low-pressure turbine 8 and configured to drive the rotor of the low-pressure compressor 4 via an intermediate shaft. The fan rotor 9 and the rotor of the high-pressure compressor 5 remain driven by the low-pressure shaft 11 and the high-pressure shaft 10, respectively.

[0060] The low pressure shaft 11 is generally housed, over a section of its length, in the high pressure shaft 10 and is coaxial with the high pressure shaft 10. The low pressure shaft 11 and the high pressure shaft 10 may be co-rotating, i.e. driven in the same direction around the longitudinal axis X. Alternatively, the low pressure shaft 11 and the high pressure shaft are counter-rotating, i.e. driven in opposite directions around the longitudinal axis X. Where appropriate, the intermediate shaft is housed between the high pressure shaft 10 and the low pressure shaft 11. The intermediate shaft and the low pressure shaft 11 may be co-rotating or counter-rotating.

[0061] The fan section 2 comprises at least the fan rotor 9 adapted to be driven in rotation relative to a stator part of the propulsion system 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.

[0062] The fan section 2 may further comprise a fan stator 16, or rectifier, which comprises blades 17 mounted on a hub 18 of the fan stator 16 and 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 similar manner 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.

[0063] In order to improve the propulsive efficiency of the propulsion system 1 and to reduce its specific consumption as well as the noise emitted by the fan section 2, the propulsion system 1 has a high bypass ratio. By high bypass ratio, we mean here a bypass ratio greater than or equal to 10, for example between 10 and 80 inclusive. To calculate the bypass ratio, the mass flow rate of the secondary air flow F2 and the mass flow rate of the primary air flow F1 are measured when the propulsion system 1 is stationary, uninstalled, in takeoff mode in a standard atmosphere (as defined by the International Civil Aviation Organization (ICAO) manual, Doc 7488 / 3, 3 eedition) and at sea level (conditions known as SLS, for Seal Level Standard). It should 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 is 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) 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 has been stopped for a sufficient period for the parts of the propulsion system to be at ambient temperature.

[0064] The fan rotor 9 is decoupled from the low-pressure shaft 11 by means of a reduction mechanism 19, placed between an upstream end of the low-pressure shaft 11 and the fan rotor 9, in order to independently optimize their respective rotational speed. In this case, the propulsion system 1 further comprises an additional shaft, called the fan shaft 20. The low-pressure shaft 11 connects the low-pressure turbine 8 to an inlet of the reduction mechanism 19 while the fan shaft 20 connects the outlet of the reduction mechanism 19 to the fan rotor 9. The fan rotor 9 is therefore driven by the low-pressure shaft 11 via the reduction mechanism 19 and the fan shaft 20 at a rotational speed lower than the rotational speed of the low-pressure turbine 8.

[0065] This decoupling makes it possible to reduce the rotational speed and the pressure ratio of the fan rotor 9 and to increase the power extracted by the low-pressure turbine 8. Indeed, the overall efficiency of the propulsion systems is conditioned to the first order by the propulsive efficiency, which is favorably influenced by a minimization of the variation in kinetic energy of the air when passing through the propulsion system 1. In a propulsion system 1 with a high bypass ratio, the majority of the flow generating the propulsive force is constituted by the secondary air flow F2 of the propulsion system 1, the kinetic energy of the secondary air flow F2 being mainly affected by the compression that the secondary air flow F2 undergoes when passing through the fan section 2. The propulsive efficiency and the pressure ratio of the fan section 2 are therefore linked: the lower the pressure ratio of the fan section 2, the better the propulsive efficiency will be.In order to optimize the propulsive efficiency of the propulsion system 1, the pressure ratio of the fan, which corresponds to the ratio between the average pressure at the outlet of the fan stator 17 (or, in the absence of a stator, of the fan rotor 9) and the average pressure at the inlet of the fan rotor 9, is less than or equal to 1.70, preferably less than or equal to 1.50, for example between 1.05 and 1.45. The average pressures are measured here over the height of the blade 14 (from the surface which radially delimits the flow path at the inlet of the fan rotor 9 to the tip 21 of the fan blade 14).

[0066] The propulsion system 1 is configured to provide a thrust of between 18,000 Ibf (80,068 N) and 51,000 Ibf (22,2411 N), preferably between 20,000 Ibf (88,964 N) and 35,000 Ibf (15,5688 N).

[0067] 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.

[0068] A ducted fan section 2 comprises a fan rotor 9 extending upstream of a fan stator. The vanes of the fan stator are then generally called outlet vanes (“Outlet Guide Vane” or “OGV” in English) and have a fixed pitch relative to the hub of the fan stator. Furthermore, the bypass ratio of the propulsion system 1 is preferably greater than or equal to 10, for example between 10 and 35 inclusive, preferably between 10 and 18 inclusive. Note that, when the bypass ratio is greater than or equal to 25, the fan rotor 9 is preferably of variable pitch. The peripheral speed at the tip 21 of the vanes of the fan rotor 9 may also be between 260 m / s and 400 m / s. The vanes 14 of the fan rotor 9 may be fixed or have a variable pitch. The blower pressure ratio can then be between 1.20 and 1.45.

[0069] 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.

[0070] The removal of the fairing around the fan section 2 makes it possible to increase the bypass ratio very significantly without the propulsion system 1 being penalized by the mass of the casings or nacelles intended to surround the fan section 2. The bypass ratio of the propulsion system 1 comprising an unfaired fan section 2 is thus greater than or equal to 40, for example between 40 and 80 inclusive. The peripheral speed at the tip 21 of the blades 14 of the fan rotor(s) 9 may also be between 210 m / s and 260 m / s. The fan pressure ratio may then preferably be between 1.05 and 1.20.

[0071] The reduction mechanism 19 may comprise, for example, a reduction mechanism 19 with an epicyclic gear train, for example of the “epicyclic” or “planetary” type, 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 can be epicyclic (“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).

[0072] 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.

[0073] 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, typically 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.

[0074] 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, preferably between 9000 rpm and 11000 rpm. The redline speed corresponds to the maximum rotation speed when the propulsion system is healthy (and potentially at the end of its life). It is therefore likely to be reached by the low pressure shaft 11 in flight conditions. This redline speed is part of the data declared in the engine certification (type certificate data sheet in English).Indeed, this rotation speed is usually used as a reference speed for the dimensioning of propulsion systems 1 and in certain certification tests (such as blade loss or rotor integrity tests).

[0075] In order to improve the efficiency of the propulsion system 1, the high pressure turbine 7 is sized so that its internal mean radius Rmoyenjnt is at least equal to:

[0076] Radius moyenJnt > A * * (T e + T ref ) ' ' GAMMA - * 10 3 - B (1) where: Rmoyenjnt is expressed in millimeters (mm);

[0077] D is the diameter of the fan rotor 9, in millimeters (mm);

[0078] BPR is the bypass ratio of the propulsion system 1;

[0079] Te is the inlet temperature of the low pressure turbine 8, in degrees Celsius (°C);

[0080] Tref = 273 K;

[0081] GAMMA is the adiabatic coefficient of air; and

[0082] A = 593 (°C) -1 / 2 and B = 6.5 millimeters (mm).

[0083] As previously stated, all of these parameters are determined when the propulsion system 1 is stationary, uninstalled, in takeoff mode in a standard atmosphere (as defined by the International Civil Aviation Organization (ICAO) manual, Doc 7488 / 3, 3 e edition) and at sea level, except for the internal mean radius and diameter which are determined when the propulsion system 1 is cold.

[0084] The internal mean radius of the high pressure turbine Rmoyenjnt is equal to the arithmetic mean of the internal radii R1 of the rotors 7b (moving blade wheels) of the high pressure turbine 7. In a given stage, the internal radius R1 of a rotor 7b corresponds to the distance, in a plane normal to the axis of rotation X of the high pressure turbine 7, between the external radial surface of the hub of the rotor 7b (which radially delimits the flow path in the rotor 7b) and the axis of rotation X, midway between the leading edge 7c and the trailing edge 7d of the moving blades of the rotor 7b at the hub (at 50% of the chord at the blade root), when the propulsion system 1 is at rest.

[0085] The sizing of the high pressure turbine 7 so as to comply with the formula (1) defined above makes it possible to obtain a compromise between an acceptable aerodynamic loading at the blade root and sufficient expansion work to increase the compression ratio of the high pressure compressor 5 (at the same number of stages) and therefore increase the overall compression ratio of the propulsion system 1. Alternatively, such sizing of the high pressure turbine 7 makes it possible to reduce the compression ratio of the low pressure compressor 4 (at the same overall compression ratio), and therefore to reduce its average radius at the same load, which makes it possible to improve the BPR by-pass ratio of the propulsion system 1.The high-pressure turbine 7 can in fact rotate at rotational speeds (in revolutions per minute) high enough to allow the high-pressure compressor 5 to achieve 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. Thus, for a two-stage high-pressure turbine 7, the high-pressure compressor 5 can comprise at least eight stages and at most eleven stages, for example nine stages.

[0086] The diameter of the high pressure turbine 7 may furthermore be large enough to allow the passage and, where appropriate, the radial movements of a supercritical low pressure shaft 11.

[0087] In one embodiment, the internal mean radius of the high pressure turbine 11 Rmoyenjnt is at most equal to 300 mm. In a propulsion system 1 with a high dilution ratio, the formula (1) also gives an internal mean radius Rmoyenjnt at least equal to 200 mm.

[0088] The inlet temperature of the low pressure turbine Te can be between 950°C and 1230°C.

[0089] In order to obtain a high-pressure turbine 7 whose internal mean radius Rmoyenjnt complies with formula (1), it is in particular possible to configure the distributor at the outlet of the combustion chamber 6 and the cross-section of the flow stream at the inlet of the high-pressure turbine 7 so that, in take-off mode (as defined above), the Mach number in the stream cross-section is substantially equal to 1. For this, the mean radius of the high-pressure turbine 7 is fixed by the acceptable aerodynamic load in the high-pressure turbine 7, which is itself defined by the energy which must be supplied to the high-pressure compressor 5 and the rotational speed of the high-pressure shaft 10 in order to obtain the desired compression ratio. The cross-section of the flow stream in the high-pressure turbine 7 depends on the reduced flow rate in the flow stream at a given Mach.Here, the flow vein section can therefore be dimensioned so that the Mach number in the vein section is equal to 1. The internal mean radius Rmoyenjnt is then deduced from the flow vein section thus dimensioned and the mean radius determined from the aerodynamic load.

[0090] The rotation speed of the high pressure turbine 7 can then be between 15,000 revolutions per minute and 27,000 revolutions per minute.

[0091] High-pressure turbines 7 complying with formula (1) may then have a hub-to-head ratio, which corresponds to the ratio between the external radius R2 of the high-pressure turbine 7 and the internal radius R1 of the rotor 7b (moving blade wheels) of the high-pressure turbine 7, greater than 0.77 and less than 0.90. The external radius R2 of the high-pressure turbine 7 and the internal radius R1 are measured here in a plane normal to the axis of rotation X at 50% of the chord at the blade root of the rotor 7b furthest downstream of the high-pressure turbine 7 (i.e., of the last stage of the high-pressure turbine 7). The external radius R2 of the high-pressure turbine 7 corresponds to the distance, in this plane, between the tip 7e of the blades of the rotor 7b of the high-pressure turbine 7 and the axis of rotation X of the high-pressure turbine 7, when the propulsion system 1 is at rest. The internal radius R1 was defined above.

[0092] Such high-pressure turbines 7 then have an optimized outlet section Ss. Indeed, the smaller the hub-head ratio, the smaller the external diameter of the high-pressure turbine 7 (at iso-section). A hub-head ratio of between 0.77 and 0.90, in combination with an optimized inlet section Se and a rotation speed as described above, thus makes it possible to obtain not only a more efficient high-pressure turbine 7 (due to its adapted rotation speed), in a suitable size, while optimizing the outlet surface which expands the gases at the outlet of the combustion chamber 5. The sizing of the high-pressure turbine 7 so as to obtain a hub-head ratio of between 0.77 and 0.90 therefore makes it possible to make the high-pressure turbine 7 more efficient, and therefore to reduce the specific consumption of the propulsion system 1, without penalizing its mechanical or thermal load.

[0093] A propulsion system 1 comprising a high-pressure turbine 7 whose internal mean radius Rmoyenjnt complies with formula (1) can then have an overall compression ratio, which corresponds to the pressure ratio between the pressure at the outlet of the high-pressure compressor 5 and the pressure at the inlet of the fan rotor 9 (measured at the level of the foot of the fan rotor 9), greater than or equal to 40 and less than or equal to 70, preferably greater than or equal to 44 and less than or equal to 55.

[0094] In addition, the propulsion system 1 further comprises an inter-turbine casing 24 extending between the high-pressure turbine 7 and the low-pressure turbine 8. The inter-turbine casing delimits the flow path between the high-pressure turbine 7 and the low-pressure turbine 8 and comprises guide vanes 25 configured to straighten an air flow exiting the high-pressure turbine 7 and thus improve the supply of the low-pressure turbine 8 and therefore the efficiency of the low-pressure body. The guide vanes 25 therefore have an aerodynamic surface configured to redirect the air flow entering the low-pressure turbine 8. In one embodiment, the propulsion system 1 comprises between twenty and thirty guide vanes 25.

[0095] If necessary, the inter-turbine casing 24 forms a structural casing of the propulsion system 1 making it possible to improve the overall dynamics of the propulsion system 1. For this purpose, the inter-turbine casing comprises an inner shroud mounted on a set of bearings of the propulsion system 1, typically a rear bearing 26 of the high-pressure shaft, an outer shroud which can be configured to absorb the mechanical forces in the propulsion system 1, as well as a series of arms extending radially between the inner shroud and the outer shroud and configured to allow the passage of services and the absorption of mechanical forces between the inner shroud and the outer shroud. The guide vanes 25 can be separate from the arms and extend between the arms and the low-pressure turbine 8.

[0096] (a) Propulsion systems with ducted fan:

[0097] Engine 1 is a twin-body propulsion system corresponding to the current technical standard (at the date of filing of this application) which is sought to be improved and which includes a ducted fan section.

[0098] Engine 2 is a twin-body propulsion system 1 in accordance with the teaching of the present application which includes a ducted fan.

[0099] The internal mean radius of the high-pressure turbine of engine 1 does not comply with the claimed formula since it is less than 236.9 mm. On the other hand, the internal mean radius of the high-pressure turbine 7 of engine 2 does comply with the claimed formula. It appears that engine 2 has an aerodynamic loading N2 2 S lower at the blade root than engine 1 while maintaining sufficient expansion work to maintain the compression ratio of the high pressure compressor, at the same number of stages.

[0100] To move from engine 1 (reference) to engine 2 (compliant with the disclosure), 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. In addition, the diameter D of the fan and the bypass ratio BPR were increased, which made it possible to improve the propulsive efficiency of engine 2.

[0101] (b) Propulsion systems with unducted fan: Engine 3 is a double-body propulsion system corresponding to the current technical standard

[0102] (as of the filing date of this application) that is sought to be improved which includes an unducted fan section.

[0103] Engine 4 is a twin-body propulsion system 1 in accordance with the teaching of the present application which includes an unducted fan section.

[0104] The internal mean radius of the high-pressure turbine of engine 3 does not comply with the claimed formula since it is less than 255.9 mm. On the other hand, the internal mean radius of the high-pressure turbine 7 of engine 4 does comply with the claimed formula. In a similar manner to the ducted engine 2, engine 4 has an aerodynamic loading N2 2 S lower at the blade root than engine 1 while maintaining sufficient expansion work to maintain the compression ratio of the high pressure compressor, at the same number of stages.

[0105] To move from engine 3 (reference) to engine 4 (compliant with the disclosure), the diameter of the fan D was reduced as well as the bypass ratio in order to facilitate the integration of engine 4 under the wing. The fan pressure ratio was also slightly increased (while remaining well below 1.45) to maintain the fan thrust. The overall compression ratio was also increased, as was the inlet temperature of the high-pressure turbine, in order to improve the thermal efficiency of the high-pressure spool and to compensate for the loss of specific consumption linked to the increase in the fan pressure ratio. The number of stages in the low-pressure compressor was also increased to increase the overall compression ratio.The reduction in the fan diameter allows for a more compact propulsion system, facilitating its installation on aircraft, and a reduction in the mass of the propulsion system, which helps reduce aircraft fuel consumption.

Claims

CLAIMS 1. Aeronautical propulsion system (1) comprising: - a first turbine (8) configured to drive a first compressor (4) 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 fan rotor (9) connected to a fan shaft (20); - 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 internal mean radius of the second turbine (7) is at least equal to: Radius moyenJnt > A * * (T e + T ref ) ' ' GAMMA - * 10 3 - B where: Rmoyenjnt is the internal mean radius of the second turbine (7) in millimeters (mm); D is the diameter of the fan rotor (9) in millimeters (mm), measured in a plane normal to the axis of rotation 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 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); Tref = 273 K; GAMMA is the adiabatic coefficient of air; and A = 593 (°C) -1 / 2 and B = 6.5 millimeters (mm).

2. Propulsion system (1) according to claim 1, in which the internal mean radius of the second turbine (7) is at most equal to 300 mm.

3. Propulsion system (1) according to one of claims 1 and 2, in which a hub-head ratio of the second turbine (7) is greater than 0.77 and less than 0.

90.

4. Propulsion system (1) according to one of claims 1 to 3, further comprising an inter-turbine casing (24) mounted on a bearing assembly (26), the inter-turbine casing extending between the first turbine (8) and the second turbine (7).

5. Propulsion system (1) according to claim 4, wherein the inter-turbine casing (24) comprises a series of guide vanes (25) configured to straighten an air flow at the inlet of the first turbine (8).

6. Propulsion system (1) according to claim 5, comprising between twenty and thirty guide vanes (25).

7. Propulsion system (1) according to one of claims 1 to 6, in which the second turbine (7) is two-stage.

8. Propulsion system (1) according to one of claims 1 to 7, in which the second compressor (5) comprises at least eight stages and at most eleven stages.

9. Propulsion system (1) according to one of claims 1 to 8, in which the diameter of the fan rotor (9) is between 80 inches (2,032 mm) and 185 inches (4,699 mm) inclusive, preferably between 85 inches (2,159 mm) and 120 inches (3,048 mm) inclusive, for example of the order of 90 inches (2,286 mm).

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

11. Propulsion system (1) according to one of claims 1 to 10, in which the fan rotor (9) 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.

12. Propulsion system (1) according to one of claims 1 to 11, in which the reduction mechanism (19) has a reduction ratio greater than or equal to 2.5, preferably greater than or equal to 3.0 and less than or equal to 11.

0.

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

14. Propulsion system (1) according to one of claims 1 to 13, in which the first compressor (4) comprises at least two stages and at most four stages.

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

16. 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 higher speed than the first turbine (8), the second turbine (7) being dimensioned so that an internal mean radius of the second turbine (7) is at least equal to: Radius moyenJnt > A * 10 3 - B where: Rmoyenjnt is the internal mean radius of the second turbine (7) in millimeters (mm); D is the diameter of the fan rotor (9) in millimeters (mm), measured in a plane normal to the axis of rotation 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 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); Tref = 273 K; GAMMA is the adiabatic coefficient of air; and A = 593 (°C) -1 / 2 and B = 6.5 millimeters (mm).

17. A method of manufacturing a propulsion system (1) comprising the following steps: dimensioning the propulsion system in accordance with claim 16; and manufacturing the propulsion system.

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