Aeronautical propulsion system with improved propulsion efficiency

Abstract

The present invention relates to an aeronautical propulsion system (1), characterized in that (Eq1), (Eq2), (Eq3), (Eq4), where L is a distance between the inlet of the compressor (5) and the outlet of the turbine (8), D is a diameter of the fan rotor (9), n is the total number of stages between inlet and outlet, XN is the actual maximum rotational speed of the shaft (11); BPR is a bypass ratio of the system, when it is in the steady-state under takeoff conditions in a standard atmosphere at sea level, Te is the temperature at the inlet of the turbine (8); Tref = 273 K; GAMMA is equal to 1.4, co is an actual maximum rotational speed of the fan rotor (9), GR is a reduction gear ratio of the mechanism (19) for stepping down from the fan shaft (20) to the shaft (11).

Description

[0001] DESCRIPTION

[0002] Aeronautical propulsion system with improved propulsive efficiency

[0003] TECHNICAL FIELD

[0004] The present application generally concerns the field of propulsion systems, and more particularly aeronautical propulsion systems comprising a ducted or non-ducted 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 rotational speeds. Generally, the decoupling is achieved using a reduction mechanism placed between the upstream end of the low-pressure shaft and a rotor of the fan section. The rotor of the fan section is then driven by the low-pressure shaft via the reduction mechanism at a rotational speed lower than that of the low-pressure shaft.Because the reduction mechanism increases the rotational speed of the low pressure shaft, the position of the deformation modes of this low pressure shaft relative to the maximum speed is modified.

[0009] However, a supercritical deformation mode of the low pressure shaft must not appear, which generates resonance frequencies in a stabilized operating range, in order to avoid the risk of damaging the propulsion system.

[0010] STATEMENT OF THE INVENTION

[0011] An aim of the present application is to optimize the propulsion system in order to increase its efficiency while controlling the dynamic behavior of the low pressure shaft integrated into the high pressure shaft.

[0012] To this end, according to a first aspect, an aeronautical propulsion system is proposed comprising: a first turbine configured to drive a first compressor via a first shaft around an axis of rotation; 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 around the axis of rotation; a fan rotor connected to a fan shaft; a reduction mechanism coupling the first shaft and the fan shaft in order to drive the fan shaft at an actual maximum rotation speed XN of the first shaft, which has been measured in an endurance test;wherein an inlet of the second compressor corresponds to the most upstream portion of a most upstream moving blade wheel of the second compressor and to a leading edge at the blade root of a most upstream first stage of the second compressor, and an outlet of the second turbine corresponds to the most downstream portion of a most downstream moving blade wheel of the second turbine and to a trailing edge at the blade root of a most downstream last stage of the second turbine, characterized in that;

[0013] XN = a) * GR (Eq1 ), with , GAMMA

[0014] I 2 z x 0 5 -

[0015] L = 153.6J^^ * (r e + T ref ) ' GAMMA— i * looo * n + 421 (Eq4), where L is a distance between the inlet of the second compressor and the outlet of the second turbine in millimeters,

[0016] D is a diameter of the fan rotor, expressed in millimeters and measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of fan rotor blades, n is the total number of stages between the inlet of the second compressor and the outlet of the second turbine, the actual maximum rotational speed XN of the first shaft is expressed in revolutions per minute;

[0017] BPR is a propulsion system bypass ratio, which is measured when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level,

[0018] T e 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;

[0019] T ref = 273 K ;

[0020] GAMMA is the adiabatic coefficient of air and is equal to 1.4, co is a real maximum rotation speed of the fan rotor,

[0021] GR is a prescribed reduction ratio of the fan shaft reduction mechanism to the first shaft.

[0022] According to one embodiment of the invention, the diameter D of the fan rotor is chosen such that where FN is a thrust generated by the fan rotor, which is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level and which is expressed in Newton.

[0023] According to one embodiment of the invention, the fan rotor is shrouded and the BPR dilution 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. According to one embodiment of the invention, the fan rotor is shrouded and a peripheral speed at the tip of the blades of the fan rotor, when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level, is between 260 m / s and 400 m / s.

[0024] According to another embodiment of the invention, the fan rotor is unducted and the BPR dilution ratio of the propulsion system is greater than or equal to 40, for example between 40 and 80 inclusive.

[0025] According to one embodiment of the invention, the fan rotor is unducted and a peripheral speed at the tip of the fan rotor blades, when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level, is between 210 m / s and 260 m / s.

[0026] According to one embodiment of the invention, the reduction ratio GR of the reduction mechanism from the fan shaft to the first shaft is greater than or equal to 2.5 and less than or equal to 11.

[0027] According to one embodiment of the invention, the fan rotor is shrouded and the reduction ratio GR of the reduction mechanism from the fan shaft to the first shaft is greater than or equal to 2.7 and less than or equal to 3.5, typically around 3.0.

[0028] According to one embodiment of the invention, the fan rotor is unducted and the reduction ratio GR of the reduction mechanism from the fan shaft to the first shaft is between 9.0 and 11.0.

[0029] According to one embodiment of the invention, the actual maximum limit speed XN of rotation of the first shaft is between 8500 revolutions per minute and 12000 revolutions per minute, preferably between 9000 revolutions per minute and 11000 revolutions per minute.

[0030] According to one embodiment of the invention, the propulsion system is configured to provide a thrust FN of between 80068 N and 222411 N, preferably between 88964 N and 155688 N.

[0031] According to one embodiment of the invention, the second turbine is two-stage.

[0032] According to one embodiment of the invention, the second compressor comprises at least 8 stages and at most 11 stages.

[0033] According to one embodiment of the invention, the first turbine comprises at least 3 stages and at most 5 stages.

[0034] According to one embodiment of the invention, the first compressor comprises at least 2 stages and at most 4 stages.

[0035] A second subject of the invention is an aircraft comprising at least one propulsion system as described above, attached to the aircraft by means of a mast.

[0036] A third subject of the invention is a method for manufacturing a propulsion system comprising: a first turbine configured to drive a first compressor via a first shaft about an axis of rotation; 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 second shaft about the axis of rotation; a fan rotor connected to a fan shaft; a reduction mechanism coupling the first shaft and the fan shaft in order to drive the fan shaft at an actual maximum rotational speed of the first shaft, which has been measured in an endurance test;wherein an inlet of the second compressor corresponds to the most upstream portion of a most upstream moving blade wheel of the second compressor and to a leading edge at the blade root of a most upstream first stage of the second compressor, and an outlet of the second turbine corresponds to the most downstream portion of a most downstream moving blade wheel of the second turbine and to a trailing edge at the blade root of a most downstream last stage of the second turbine, characterized in that;

[0037] XN = a) * GR (Eq1 ), q4), where L is a distance between the inlet of the second compressor and the outlet of the second turbine in millimeters,

[0038] D is a diameter of the fan rotor, expressed in millimeters and measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of fan rotor blades, n is the total number of stages between the inlet of the second compressor and the outlet of the second turbine, the actual maximum rotational speed XN of the first shaft is expressed in revolutions per minute;

[0039] BPR is a BPR ratio of the propulsion system, which is measured when the propulsion system is stationary in takeoff regime in a standard atmosphere and at sea level, T e 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;

[0040] T ref = 273 K ;

[0041] GAMMA is the adiabatic coefficient of air and is equal to 1.4, co is a real maximum rotation of the fan rotor,

[0042] GR is a prescribed reduction ratio of the fan shaft reduction mechanism to the first shaft.

[0043] According to one embodiment of the invention of the manufacturing method, the diameter D of the fan rotor is chosen such that where FN is a thrust generated by the propulsion system, which is measured when the propulsion system is stationary in takeoff mode in a standard atmosphere and at sea level and which is expressed in Newton.

[0044] 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] [Fig. 1] is a schematic, partial and sectional view of an example of a propulsion system according to a first embodiment of the invention, in which the fan section is ducted;

[0047] [Fig. 2] is a schematic, partial and sectional view of an example of a propulsion system according to a first embodiment of the invention, in which the fan section is unducted;

[0048] [Fig. 3] is a schematic sectional view of an example of a planetary reduction mechanism;

[0049] [Fig. 4] is a schematic sectional view of an example of an epicyclic reduction mechanism;

[0050] [Fig. 5] is an example of an aircraft that may include at least one propulsion system in accordance with the first or second embodiment of the invention; [Fig. 6] is a flowchart illustrating example steps of a dimensioning method in accordance with one embodiment of the invention.

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

[0052] DETAILED DESCRIPTION OF THE INVENTION

[0053] A propulsion system 1 has a main direction extending along a longitudinal axis X and comprises, from upstream to downstream in the direction of flow of the gases in the propulsion system 1 when it is in operation, a fan section 2 and a primary body 3, often called a “gas generator”, comprising a compressor section 4, 5, a combustion chamber 6 and a turbine section 7, 8. The propulsion system 1 is here an aeronautical propulsion system 1 configured to be fixed to an aircraft 100 via a pylon (or mast).

[0054] The compressor section 4, 5 comprises a succession of stages each comprising a moving blade wheel (rotor) 4a, 5a rotating in front of a fixed blade wheel (stator) 4b, 5b. The turbine section 7, 8 also comprises a succession of stages each comprising a fixed blade wheel (stator) 7a, 8a behind which a moving blade wheel (rotor) 7b, 8b rotates.

[0055] In the present application, the axial direction corresponds to the direction of the longitudinal axis X, in correspondence with the rotation of the shafts of the gas generator, and a radial direction is a direction perpendicular to this axis X and passing through it. Furthermore, the circumferential (or lateral, or even tangential) direction corresponds to a direction perpendicular to the longitudinal axis X and not passing through it. Unless otherwise specified, internal (respectively, interior) and external (respectively, exterior), respectively, are used in reference to a radial direction so that the internal part or face of an element is closer to the axis X than the external part or face of the same element.

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

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

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

[0059] In a twin-spool propulsion system 1, the compressor section 4, 5 may comprise a low-pressure compressor 4 (first compressor 4) and a high-pressure compressor 5 (second compressor 5), located downstream of the low-pressure compressor 4. The turbine section 7, 8 may comprise a high-pressure turbine 8 (second turbine 8), located downstream of the combustion chamber 6, and a low-pressure turbine 7 (first turbine 7), located downstream of the high-pressure turbine 8. The rotor of the high-pressure compressor 5 is rotated by the rotor of the high-pressure turbine 8 via a high-pressure shaft 10 (or second 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 7 via a low-pressure shaft 11 (or first shaft 11).Thus, the primary body 3 comprises a high-pressure body comprising the high-pressure compressor 5, the high-pressure turbine 8 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 7 and the low-pressure shaft 11. The rotational speed of the high-pressure body is greater than the rotational speed of the low-pressure body. In a triple-body propulsion system 1, the turbine section 7, 8 further comprises an intermediate turbine, positioned between the high-pressure turbine 8 and the low-pressure turbine 7 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 rotated relative to a stator portion 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 13 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 the 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 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 16 may be fixed relative to the hub or have a variable pitch. In a similar manner to the rotor blades 14, the root of the stator blades 17 may be pivotally mounted along a pitch axis X and be 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] The fan rotor 9 further comprises at least 14 blades 14 and at most 24 blades 14, preferably at least 16 blades 14 and at most 22 blades 14. The number of blades 16 in the fan stator 16 depends on the acoustic criteria defined for the propulsion system 1 and is at least equal to the number of blades 14.

[0064] 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 (BPR). By high bypass ratio, we mean here a bypass ratio BPR greater than or equal to 10, for example between 10 and 80 inclusive. To calculate the bypass ratio BPR, 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. It will be noted that, in the present application, the parameters (pressure, flow rate, FN thrust mentioned below, 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.

[0065] The fan rotor 9 is decoupled from the low-pressure shaft 11 using 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 7 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 shaft

[0066] 20 of the fan at a rotational speed lower than the rotational speed of the low pressure turbine 7. The reduction ratio GR of the reduction mechanism 19 is greater than 1 and is equal to the rotational speed of the low pressure turbine 7 (or of the low pressure shaft 11), divided by the rotational speed of the fan rotor 9. The geometry of the constituent elements of the reduction mechanism 19 determines the reduction ratio GR thereof.

[0067] 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 7. 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 BPR 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 16 (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 0.90 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 top.

[0068] 21 of the fan blade 14).

[0069] 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.

[0070] 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 BPR 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. 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 fan pressure ratio may then be between 1.20 and 1.45.

[0071] In an unducted fan section 2, the fan section 2 (which may also be referred to as the propeller) is not surrounded by a fan casing. Since the fan section 2 is unducted, the blades 14 of the fan rotor 9 have a variable pitch. Propulsion systems comprising at least one unducted fan rotor 9 are known as “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.

[0072] The removal of the fairing around the fan section 2 makes it possible to increase the BPR by-pass 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 BPR by-pass 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 0.90 and 1.20.

[0073] The reduction mechanism 19 may comprise an epicyclic or planetary, single-stage or two-stage reduction mechanism 19. For example, 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 satellite 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. Alternatively, the reduction mechanism 19 may be epicyclic (“planetary” in English) (Figure 4), in which case the ring gear 19b is fixedly mounted on the stator part 19e of the propulsion system 1 and the fan shaft 20 is rotated by the planet carrier 19d.

[0074] Whatever the configuration of the reduction mechanism 19, the diameter of the crown 19b and the planet carrier 19d are greater than the diameter of the sun gear 19a, so that the rotational speed of the fan rotor 9 is lower than the rotational speed of the low pressure shaft 11.

[0075] The reduction ratio GR 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 GR may be greater than or equal to 2.7 and less than or equal to 3.5, typically around 3.0. In the case of a propulsion system 1 comprising an unshrouded fan rotor, the reduction ratio GR may be between 9.0 and 11.0.

[0076] The dual-body propulsion system 1 may in particular comprise a two-stage high-pressure turbine 8, a high-pressure compressor 5 comprising at least eight stages and at most eleven stages, a low-pressure turbine 7 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.

[0077] According to a first configuration of the invention, in order to optimize the performance of the propulsion system 1, the reduction mechanism 19 couples the first shaft 11 and the fan shaft 20 in order to drive the fan shaft 20 at a rotation speed lower than an actual maximum rotation speed XN of the first shaft 11.

[0078] According to the first configuration of the invention,

[0079] XN = a) * GR (Eq1 ), with , GAMMA

[0080] I 2 z x 0 5 -

[0081] L = 153.6J^^ * (r e + T ref ) ' GAMMA— i * looo * n + 421 (Eq4), where L is a distance between the inlet of the second compressor 5 and the outlet of the second turbine 8 in millimeters,

[0082] D is a diameter of the fan rotor 9, expressed in millimeters and measured in a plane normal to the axis of rotation X at the level of an intersection between the tip 21 and a leading edge 22 of blades 14 of the fan rotor 9, the actual maximum speed XN of rotation of the first shaft 11 is expressed in revolutions per minute;

[0083] BPR is the above-mentioned BPR bypass ratio of propulsion system 1, which is measured when propulsion system 1 is stationary in takeoff mode in a standard atmosphere and at sea level,

[0084] T e is the temperature at the inlet of the first turbine 7 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;

[0085] T ref = 273 K ;

[0086] GAMMA is the adiabatic coefficient of air and is equal to 1.4, co is a real maximum rotation speed of the fan rotor 9,

[0087] GR is the prescribed reduction ratio of the reduction mechanism 19 of the fan shaft 20 towards the first shaft 11, n is the total number of stages between the inlet of the second compressor 5 and the outlet of the second turbine 8. The number n is therefore the number of stages of the high pressure body.

[0088] In equations and inequalities, multiplication is denoted by the * sign, while the . sign denotes the fixed-point representation of numbers. Note that since Figures 1 and 2 are partial views, the diameter D is only partially visible.

[0089] The distance L between the inlet of the high-pressure compressor 5 and the outlet of the high-pressure turbine 8 corresponds to the length L of the high-pressure body. The high-pressure body comprises the high-pressure compressor 5, the high-pressure turbine 8 and the high-pressure shaft 10.

[0090] The actual maximum rotation speed XN of the first shaft 11 was measured in the endurance test of the propulsion system 1. The actual maximum rotation speed XN of the first shaft 11 is expressed in revolutions per minute.

[0091] The rotation speed XN of the first shaft 11 corresponds to a maximum rotation speed of the low-pressure turbine 7. The rotation speed XN is equal to the actual maximum rotation speed co of the fan rotor 9, multiplied by the reduction ratio GR of the reduction mechanism 19 of the fan shaft 20 towards the first shaft 11. The actual rotation speed of the first shaft 11 during operation of the propulsion system 1 may be different from the actual maximum rotation speed XN of the first shaft 11 and may be less than or equal to the actual maximum rotation speed XN of the first shaft 11. The actual maximum rotation speed XN of the first shaft 11 may therefore be reached by the actual rotation speed of the first shaft 11 of the actual propulsion system in operation.

[0092] The actual maximum rotational speed of the fan rotor 9 is expressed in revolutions per minute.

[0093] The actual rotational speed of the fan rotor 9 during operation of the propulsion system 1 may be different from the actual maximum rotational speed co of the fan rotor 9 and may be less than or equal to the actual maximum rotational speed co of the fan rotor 9. The actual maximum rotational speed co of the fan rotor 9 may therefore be achieved by the actual rotational speed of the fan rotor 9 of the actual propulsion system in operation.

[0094] The actual maximum rotational speed XN of the first shaft 11 is defined by regulation and is part of the data, which are declared and certified for the actual propulsion system in a TCDS certificate (in English: Type Certificate Data Sheet) and which have been demonstrated in the endurance test having been carried out on the actual propulsion system.

[0095] The actual maximum rotational speed XN of the first shaft 11 is a data declared in advance for any propulsion system and is therefore prescribed.

[0096] This first configuration of the invention makes it possible to size the high pressure shaft 10 so that the dynamics of the low pressure shaft 11 are controlled.

[0097] In a second configuration of the invention, the diameter D of the fan rotor 9, expressed in millimeters and measured in the plane normal to the axis of rotation X at the level of the intersection between the tip 21 and the leading edge 22 of blades 14 of the fan rotor 9, is chosen such that where FN is the thrust generated by the fan rotor 9, which is measured when the propulsion system 1 is stationary in takeoff mode in a standard atmosphere and at sea level and is expressed in Newtons. In equations and inequalities, multiplication is denoted by the * sign, while the . sign denotes the fixed-point representation of numbers.

[0098] This second configuration of the invention makes it possible to dimension the torque parameter ^ gr on the low pressure shaft 11 (or first shaft 11) compared to the other parameters BPR, D and T e .

[0099] The second configuration of the invention can be provided without the first configuration of the invention.

[0100] The first configuration of the invention can be provided without the second configuration of the invention.

[0101] The first configuration of the invention can be provided with the second configuration of the invention.

[0102] In the non-limiting numerical examples described below in Table 1:

[0103] Engine 1 is a comparative example (reference engine) formed by a double-body propulsion system comprising a fan section ducted and driven in rotation by means of a gear train reduction mechanism corresponding to a state of the art according to a current technical standard which is sought to be improved by engines 2, 3 and 4, which are according to the invention. The low-pressure shaft of engine 1 is supported by a front ball bearing and two rear roller bearings.

[0104] Engines 2, 3 and 4 are examples according to the invention, each formed by a double-body propulsion system comprising a shrouded fan section driven in rotation by means of a gear train reduction mechanism.

[0105] Engines 2 and 3 are propulsion systems combining the first configuration of the invention with the second configuration of the invention, where: o engines 2 and 3 allow better overall performance than engine 1 due to the improvement in thermopropulsive efficiency and the greater compactness of the fan section, o engine 3 differs from engine 2 by a more compact low-pressure turbine and a higher rotational speed of the low-pressure turbine. This is made possible by an increase in the temperature at the high-pressure turbine inlet associated with an increase in the fuel flow rate.

[0106] Engine 4 is a propulsion system corresponding to the first configuration of the invention without the second configuration of the invention, where: o engine 4 has a fan section of diameter equivalent to engine 1 and allows better overall performance than engine 1 due to the improvement in thermopropulsive efficiency.

[0107] The low-pressure shaft of engines 2, 3 and 4 is supported by a front ball bearing including a mode damper and two rear roller bearings, the most downstream roller bearing also including a mode damper. The reduction mechanism and the second downstream bearing do not include a mode damper. Table 1:

[0108] It can be seen that in Table 1 the actual maximum rotation speed XN of the low pressure turbine of the reference engine 1 of the comparative example does not respect the inequalities Eq2 and Eq3.

[0109] It can be seen that in Table 1 the actual maximum rotation speed XN of the low pressure turbine of engines 2, 3 and 4 of the examples according to the invention complies with inequalities Eq2 and Eq3.

[0110] It can be seen that in Table 1 the diameter D of the fan rotor of the reference engine 1 of the comparative example and of the example of the engine 4 according to the invention does not respect the inequalities Eq5 and Eq6.

[0111] It can be seen that in Table 1 the diameter D of the fan rotor of the examples of engines 2 and 3 according to the invention respects the inequalities Eq5 and Eq6.

[0112] In the above, the actual maximum rotation speed XN of the first shaft 11 (redline speed in English) 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, may be between 8500 rpm and 12000 rpm, preferably between 9000 rpm and 11000 rpm. The limit speed XN corresponds to the maximum rotation speed when the propulsion system is healthy. The actual maximum rotation speed XN of the first shaft 11 is therefore likely to be reached by the low pressure shaft 11 in flight conditions. This limit speed XN is part of the data declared in the engine certification (“type certificate data sheet” in English).Indeed, this actual maximum rotation speed XN of the first shaft 11 is usually used as the reference speed XN for the dimensioning of propulsion systems 1 and in certain certification tests (such as blade loss or rotor integrity tests).

[0113] The propulsion system 1 may be configured to provide a thrust FN of between 80068 N and 222411 N, preferably between 88964 N and 155688 N.

[0114] A manufacturing method of the propulsion system may provide for dimensioning D (step E1) and / or dimensioning L (step E2).

[0115] Step E1 can be before step E2 (figure 6) or after step E2 or simultaneous with step E2.

[0116] In the above, the person skilled in the art, in order to increase the actual maximum rotational speed XN of the first shaft 11, can lower the gas flow rate in the first turbine 7 or increase the flow rate of the gas through the turbine 7. In both cases this results in a reduction in the diameter of the first turbine 7.

[0117] The first possibility of lowering the gas flow in the first turbine 7 can be achieved by the fact that the temperature T e at the inlet of the first turbine 7 is increased. This first possibility of reducing the gas flow rate in the first turbine 7 can be achieved by increasing the fuel flow rate in the combustion chamber 6. To reduce the actual maximum rotation speed XN of the first shaft 11, a person skilled in the art can proceed in the opposite manner.

[0118] The person skilled in the art may choose not to modify the actual maximum rotation speed co of the fan rotor 9, by modifying the reduction ratio GR of the reduction mechanism 19 to adapt to a modification of the actual maximum rotation speed XN of the first shaft 11, by the formula XN = co * GR. The person skilled in the art knows how to modify the reduction ratio GR by modifying the geometry of the reduction mechanism 19.

[0119] A person skilled in the art can also modify the actual maximum rotational speed co of the fan rotor 9 by modifying in another way the reduction ratio GR of the reduction mechanism 19.

Claims

CLAIMS 1. Aeronautical propulsion system (1) comprising: a first turbine (7) configured to drive a first compressor (4) via a first shaft (11) about an axis (X) of rotation; a second turbine (8) 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) about the axis (X) of rotation; 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 rotation speed lower than an actual maximum rotation speed XN of the first shaft (11), which was measured in an endurance test;wherein an inlet of the second compressor (5) corresponds to the most upstream part of a movable blade wheel (5a) most upstream of the second compressor (5) and to a leading edge (5c) at the blade root of a first stage most upstream of the second compressor (5), and an outlet of the second turbine (8) corresponds to the most downstream part of a movable blade wheel (8b) most downstream of the second turbine (8) and to a trailing edge (8d) at the blade root of a last stage most downstream of the second turbine (8), characterized in that; XN = a) * GR (Eq1 ), q4), where L is a distance between the inlet of the second compressor (5) and the outlet of the second turbine (8) in millimeters, D is a diameter of the fan rotor (9), expressed in millimeters and measured in a plane normal to the axis (X) of rotation at an intersection between a tip (21) and a leading edge (22) of blades (14) of the fan rotor (9), n is the total number of stages between the inlet of the second compressor (5) and the outlet of the second turbine (8), the actual maximum rotation speed XN of the first shaft (11) is expressed in revolutions per minute; BPR is a propulsion system bypass ratio (1), which is measured when the propulsion system (1) is stationary in takeoff mode in a standard atmosphere and at sea level, T e is the temperature at the inlet of the first turbine (7) 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; T ref = 273 K ; GAMMA is the adiabatic coefficient of air and is equal to 1.4, co is a real maximum rotation speed of the fan rotor (9), GR is a prescribed reduction ratio of the reduction mechanism (19) from the fan shaft (20) to the first shaft (11).

2. Aeronautical propulsion system (1) according to claim 1, in which the diameter D of the fan rotor (9) is chosen such that where FN is a thrust generated by the fan rotor (9), which is measured when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and which is expressed in Newton.

3. Propulsion system (1) according to any one of the preceding claims, in which the fan rotor (9) is shrouded and the BPR dilution 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.

4. Propulsion system (1) according to any one of the preceding claims, in which the fan rotor (9) is shrouded and a peripheral speed at the tip (21) of blades (14) of the fan rotor (9), when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level, is between 260 m / s and 400 m / s.

5. Propulsion system (1) according to claim 1 or 2, in which the fan rotor (9) is unducted and the BPR dilution ratio of the propulsion system (1) is greater than or equal to 40, for example between 40 and 80 inclusive.

6. Propulsion system (1) according to any one of claims 1, 2 and 5, in which the fan rotor (9) is unducted and a peripheral speed at the tip (21) of the blades (14) of the fan rotor (9), when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level, is between 210 m / s and 260 m / s.

7. Propulsion system (1) according to any one of the preceding claims, in which the reduction ratio GR of the mechanism (19) for reducing the fan shaft (20) towards the first shaft (11) is greater than or equal to 2.5 and less than or equal to 11.

8. Propulsion system (1) according to any one of claims 1 to 4, in which the fan rotor (9) is shrouded and the reduction ratio GR of the reduction mechanism (19) of the fan shaft (20) towards the first shaft (11) is greater than or equal to 2.7 and less than or equal to 3.5, typically around 3.

0.

9. Propulsion system (1) according to any one of claims 1, 2, 5, 6, in which the fan rotor (9) is unducted and the reduction ratio GR of the reduction mechanism (19) of the fan shaft (20) towards the first shaft (11) is between 9.0 and 11.

0.

10. Propulsion system (1) according to any one of the preceding claims, in which the actual maximum speed XN of rotation of the first shaft (11) is between 8500 revolutions per minute and 12000 revolutions per minute, preferably between 9000 revolutions per minute and 11000 revolutions per minute.

11. Propulsion system (1) according to any one of the preceding claims, wherein the propulsion system 1 is configured to provide a thrust FN of between 80068 N and 222411 N, preferably between 88964 N and 155688 N.

12. Propulsion system (1) according to any one of the preceding claims, in which the second turbine (8) is two-stage.

13. Propulsion system (1) according to any one of the preceding claims, wherein the second compressor (5) comprises at least 8 stages and at most 11 stages.

14. Propulsion system (1) according to any one of the preceding claims, in which the first turbine (7) comprises at least 3 stages and at most 5 stages.

15. Propulsion system (1) according to any one of the preceding claims, in which the first compressor (4) comprises at least 2 stages and at most 4 stages.

16. Aircraft comprising at least one propulsion system (1) according to any one of the preceding claims, fixed to the aircraft by means of a mast.

17. Method for manufacturing a propulsion system (1) comprising: a first turbine (7) configured to drive a first compressor (4) via a first shaft (11) around an axis (X) of rotation; a second turbine (8) configured to drive a second compressor (5) via a second shaft (10), the second shaft (10) being configured to rotating at a higher speed than the second shaft (10) around the axis (X) of rotation; 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 an actual maximum rotational speed XN of the first shaft (11), which was measured in endurance testing; wherein an inlet of the second compressor (5) corresponds to the most upstream part of a movable blade wheel (5a) most upstream of the second compressor (5) and to a leading edge (5c) at the blade root of a first stage most upstream of the second compressor (5), and an outlet of the second turbine (8) corresponds to the most downstream part of a movable blade wheel (8b) most downstream of the second turbine (8) and to a trailing edge (8d) at the blade root of a last stage most downstream of the second turbine (8), characterized in that. XN = a) * GR (Eq1 ), q4), where L is a distance between the inlet of the second compressor (5) and the outlet of the second turbine (8) in millimeters, D is a diameter of the fan rotor (9), expressed in millimeters and measured in a plane normal to the axis of rotation (X) at an intersection between a tip (21) and a leading edge (22) of blades (14) of the fan rotor (9), n is the total number of stages between the inlet of the second compressor (5) and the outlet of the second turbine (8), the actual maximum rotation speed XN of the first shaft (11) is expressed in revolutions per minute; BPR is a propulsion system bypass ratio (1), which is measured when the propulsion system (1) is stationary in takeoff mode in a standard atmosphere and at sea level, T e is the inlet temperature of the first turbine (7) when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at the level of the sea ​​and is expressed in degrees Celsius; T ref = 273 K ; GAMMA is the adiabatic coefficient of air and is equal to 1.4, co is an actual maximum rotational speed of the fan rotor (9), GR is a prescribed reduction ratio of the reduction mechanism (19) of the fan shaft (20) to the first shaft (11).

18. Method of manufacturing a propulsion system (1) according to claim 17, in which the diameter D of the fan rotor (9) is chosen such that where FN is a thrust generated by the fan rotor (9), which is measured when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level and which is expressed in Newton.

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