Controlling dynamic behavior of aviation propulsion system

Abstract

An aviation propulsion system (1) comprising bearings (11a-11c) supporting a drive shaft of a fan rotor (9) in the propulsion system (1), the position of the bearings satisfying the following formula (I): wherein: d2 is the distance between the centers of gravity (Gnb and Gnc) of the rear bearings (11b, 11c) of the drive shaft (11); rm is the average radius of the turbine drive unit (8); xN is the limit speed of the drive shaft (11); d1 is the distance between the center of gravity (G11a) of the front bearing (1ia) and the center of gravity (Gnb) of the rear bearing (11b) closest to the front bearing (1ia); r11b is the average radius of the rear bearing (11b) closest to the front bearing (1ia); c = 170 (mm * Tr / min)-1, and E1 = 0.18.

Description

Technical Field

[0001] The present application relates generally to the field of propulsion systems, and more particularly to aviation propulsion systems including ducted or unducted fans and having a high or very high bypass ratio. Background Art

[0002] The propulsion system generally includes, from upstream to downstream along the direction of gas flow, a fan section, a compressor section, a combustor, and a turbine section. The compressor section may include a low-pressure compressor and a high-pressure compressor, and the turbine section may specifically include a high-pressure turbine and a low-pressure turbine. The high-pressure compressor is rotationally driven by the high-pressure turbine via a high-pressure shaft. The fan and the low-pressure compressor (where applicable) are rotationally driven by the low-pressure turbine via a low-pressure shaft.

[0003] Technological research has already significantly improved the environmental performance of aircraft. The applicant has taken into account these factors that influence all stages of design and development in order to obtain aviation components and substances that consume less energy, are more environmentally friendly, and have a moderate impact on the environment when integrated and used in civil aviation, with the aim of improving the energy efficiency of these aircraft.

[0004] Therefore, in order to improve the propulsion efficiency of the propulsion system and reduce the unit fuel consumption of the propulsion system and the noise emitted by the fan section, a propulsion system with a high bypass ratio (Bypass Ratio, BPR) (corresponding to the ratio of the flow rate of the secondary air flow to the flow rate of the main air flow) has been proposed. In order to achieve such a bypass ratio, the fan section can be separated from the low-pressure turbine, so that the corresponding rotational speeds of the fan section and the low-pressure turbine can be optimized independently. Typically, the separation is achieved using a reduction mechanism arranged between the upstream end of the low-pressure shaft and the 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 speed lower than the speed of the low-pressure shaft.

[0005] However, an increase in the rotational speed of the low-pressure turbine has an impact on the dynamics of the low-pressure body, in particular on the deformation modes of the low-pressure shaft. This impact is even greater when one seeks to increase the efficiency of the low-pressure turbine. Summary of the Invention

[0006] One object of the present application is to optimize the performance of an aeronautical propulsion system, in particular in terms of specific fuel consumption, while controlling the dynamic behavior of the low-pressure body.

[0007] To this end, according to a first aspect, an aviation propulsion system is provided, the aviation propulsion system comprising:

[0008] - a fan rotor, the fan rotor being configured to be driven by the fan shaft;

[0009] a drive turbine configured to drive the fan rotor about a rotation axis via a drive shaft;

[0010] - a speed reduction mechanism coupling the drive shaft and the fan shaft so as to drive the fan shaft at a rotational speed lower than that of the drive shaft;

[0011] - bearings configured to center the drive shaft relative to the axis of rotation, the bearings comprising a front bearing extending upstream of a combustion chamber of the propulsion system and two rear bearings extending downstream of the combustion chamber, wherein the position of the bearings in the propulsion system complies with the following formula:

[0012]

[0013] Where: d2 is the distance between the centers of gravity of the rear bearings of the drive shaft, in millimeters (mm);

[0014] R m is the average radius of the driving turbine in millimeters (mm);

[0015] XN is the limit speed of the drive shaft, in revolutions per minute (tr / min);

[0016] d1 is the distance between the center of gravity of the front bearing and the center of gravity of the rear bearing closest to the front bearing, in millimeters (mm);

[0017] R 11b is the average radius of the rear bearing closest to the front bearing, in millimeters (mm); and

[0018] C=170(mm*tr / min) -1 , E1=0.18.

[0019] Some preferred but non-limiting features of the propulsion system according to the first aspect are the following features taken alone or in combination:

[0020] -The bearings are also positioned in the propulsion system to comply with the following formula:

[0021]

[0022] Among them, E2 = 0.378;

[0023] -The limiting speed of the drive shaft complies with the following formula:

[0024]

[0025] Where, A=5500mm*tr / min;

[0026] -The limiting speed of the drive shaft complies with the following formula:

[0027]

[0028] wherein B = 8500 mm*tr / min;

[0029] - all or part of the bearings of the drive shaft are associated with a bearing mode damper;

[0030] - the bearing mode damper comprises a damping by pressurized fluid film;

[0031] - each bearing mode damper comprises a deformable cage mounted between the ring of the associated bearing and the stator part of the propulsion system;

[0032] - the bearings of the drive shaft comprise at least one front bearing and exactly two rear bearings;

[0033] - the propulsion system further comprises an additional turbine configured to drive an additional compressor through an additional shaft configured to rotate around an axis of rotation at a higher speed than the speed of the drive shaft, the additional turbine being a two-stage turbine;

[0034] - the average radius of the bore of the additional turbine complies with the following formula:

[0035] R m_a ≤ E*L HP 2 * XN*10 -9 + F1

[0036] wherein: R m_a is the average radius of the bore of the additional turbine in millimeters (mm); L HP is the distance between the inlet of the additional compressor and the outlet of the additional turbine in millimeters (mm); and E = 3.15 (mm.tr / min) -1 and F1 = 23 millimeters (mm);

[0037] - the average radius of the bore of the additional turbine complies with the following formula:

[0038] R m_a ≥ E*L HP 2 * XN*10 -9 + F2

[0039] wherein: R m_a is the average radius of the bore of the additional turbine in millimeters (mm); L HP is the distance between the inlet of the additional compressor and the outlet of the additional turbine in millimeters (mm); and E = 3.15 (mm.tr / min) -1 and F2 = 13 millimeters (mm);

[0040] - The average radius of the hole for the additional turbine is in accordance with the following formula:

[0041]

[0042] Where: D is the diameter of the fan rotor in millimeters and is measured in a plane perpendicular to the axis of rotation at the intersection between the tip and the leading edge of the fan rotor blades; BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level; T e is the temperature at the inlet of the drive turbine when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level, and is expressed in degrees Celsius (°C); T ref = 273K; n is the number of stages of the additional turbine and the additional compressor; GAMMA is the adiabatic coefficient of air; and K = 6.76, L = 153.6m -1 .(℃) -1 / 2 .(tr / min) -1 , M=421mm.(℃) -1 / 2 and N1 = -11 millimeters (mm);

[0043] - The average radius of the hole for the additional turbine is in accordance with the following formula:

[0044]

[0045] Where: D9 is the diameter of the fan rotor in millimeters and is measured in a plane perpendicular to the axis of rotation at the intersection between the tip and the leading edge of the fan rotor blades; BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level; T e is the temperature at the inlet of the drive turbine when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level, and is expressed in degrees Celsius (°C); T ref = 273K; n is the number of stages of the additional turbine and the additional compressor; gamma is the adiabatic coefficient of air; and K = 6.76, L = 153.6m -1 .(℃) -1 / 2 .(tr / min) -1 , M=421mm.(℃) -1 / 2 and N2 = -21 millimeters (mm);

[0046] - The average radius of the hole for the additional turbine is in accordance with the following formula:

[0047] R m_a ≥G*FN*BPR*10 -4 +H

[0048] Where: R m_a is the mean radius of the bore of the additional turbine, in millimeters; FN is the thrust of the fan rotor measured when the propulsion system is at rest, at takeoff rating, at standard atmospheric pressure and at sea level, and is expressed in Newtons (N); BPR is the bypass ratio of the propulsion system, and is measured when the propulsion system is at rest, at takeoff rating, at standard atmospheric pressure and at sea level; and G = 0.16 millimeters per Newton (mm / N) and H = 18 millimeters (mm);

[0049] - the average radius of the bore of the high-pressure turbine is at most equal to:

[0050] R m_a ≤I*FN*BPR*10 -4 +J

[0051] Where: R m_a is the mean radius of the bore of the additional turbine, expressed in millimeters (mm); FN is the thrust of the fan rotor measured when the propulsion system is at rest, at takeoff rating, at standard atmospheric pressure and at sea level, and is expressed in Newtons (N); BPR is the bypass ratio of the propulsion system, and is measured when the propulsion system is at rest, at takeoff rating, at standard atmospheric pressure and at sea level; and I = 0.16 millimeters / Newton (mm / N) and J = 28 millimeters;

[0052] - the additional compressor comprises at least eight stages and at most eleven stages;

[0053] - the driving turbine comprises at least three stages and at most five stages; and / or

[0054] - The reduction ratio of the reduction mechanism is greater than or equal to 2.5 and less than or equal to 11.

[0055] According to a second aspect, there is provided an aircraft comprising at least one propulsion system according to the first aspect, the propulsion system being attached to the aircraft via a pylon.

[0056] According to a third aspect, a method for determining the size or manufacturing a propulsion system is provided, the propulsion system comprising a reduction mechanism and bearings, the reduction mechanism coupling a drive shaft and a fan rotor so as to drive the fan rotor at a rotational speed lower than that of the drive shaft; the bearings being configured to center the drive shaft relative to the axis of rotation, the bearings comprising a forward bearing extending upstream of a combustion chamber of the propulsion system and two aft bearings extending downstream of the combustion chamber,

[0057] The propulsion system is dimensioned so that the position of the bearings in the propulsion system complies with the following formula:

[0058]

[0059] Where: d2 is the distance between the centers of gravity of the rear bearings of the drive shaft, in millimeters (mm);

[0060] R m is the average radius of the driving turbine in millimeters (mm);

[0061] XN is the limit speed of the drive shaft, in revolutions per minute (tr / min);

[0062] d1 is the center of gravity of the front bearing (G 11a ) and the centre of gravity of the rear bearing nearest to the front bearing, in millimetres (mm);

[0063] R 11b is the average radius of the rear bearing closest to the front bearing, in millimeters; and

[0064] C=170(mm*tr / min) -1 , E1=-0.18.

[0065] According to some preferred but non-limiting features of the method for dimensioning or manufacturing a propulsion system of the third aspect, the propulsion system is dimensioned such that the bearings are also positioned in the propulsion system to comply with the following formula:

[0066]

[0067] Among them, E2 = 0.378;

[0068] The propulsion system is also dimensioned so that the limiting speed of the drive shaft complies with the following formula:

[0069]

[0070] Where, A=5500mm*tr / min;

[0071] The propulsion system is also dimensioned so that the limiting speed of the drive shaft also complies with the following formula:

[0072]

[0073] Wherein, B = 8500mm*tr / min;

[0074] The propulsion system further comprises an additional turbine configured to drive the compressor via an additional shaft configured to rotate about the axis of rotation at a speed higher than that of the drive shaft, the average radius of the bore of the additional turbine being in accordance with the following formula:

[0075] R m_a ≤E*L HP 2 *XN*10-9 +F1

[0076] wherein: R m_a is the average radius of the hole of the additional turbine, in millimeters (mm); L HP is the distance between the inlet of the additional compressor and the outlet of the additional turbine, in millimeters (mm); and E = 3.15 (mm.tr / min) -1 and F1 = 23 millimeters (mm);

[0077] - the propulsion system further comprises an additional turbine configured to drive the compressor through an additional shaft configured to rotate around an axis of rotation at a higher speed than the speed of the driving shaft, the average radius of the hole of the additional turbine complying with the following formula:

[0078] R m_a ≥ E*L HP 2 * XN* 10 -9 +F2

[0079] wherein: R m_a is the average radius of the hole of the additional turbine, in millimeters (mm); L HP is the distance between the inlet of the additional compressor and the outlet of the additional turbine, in millimeters (mm); and E = 3.15 (mm.tr / min) -1 and F2 = 13 millimeters (mm);

[0080] - the propulsion system further comprises an additional turbine configured to drive the compressor through an additional shaft configured to rotate around an axis of rotation at a higher speed than the speed of the driving shaft, the average radius of the hole of the additional turbine complying with the following formula:

[0081]

[0082] wherein: D9 is the diameter of the fan rotor, in millimeters, and is measured in a plane perpendicular to the axis of rotation, at the intersection between the tip and the leading edge of the blades of the fan rotor; BPR is the bypass ratio of the propulsion system, and is measured when the propulsion system is at rest, at the takeoff rating, at standard atmospheric pressure and at sea level; T e is the temperature at the inlet of the driving turbine when the propulsion system is at rest, at the takeoff rating, at standard atmospheric pressure and at sea level, and is expressed in degrees Celsius (°C); T ref = 273 K; n is the number of stages of the additional turbine and of the additional compressor; gamma is the adiabatic coefficient of the air; and K = 6.76, L = 153.6 m -1 (°C) -1 / 2 (tr / min) -1 , M = 421 mm. (°C)-1 / 2 and N1 = -11 millimeters (mm);

[0083] The propulsion system further comprises an additional turbine configured to drive the compressor via an additional shaft configured to rotate about the axis of rotation at a speed higher than that of the drive shaft, the average radius of the bore of the additional turbine being in accordance with the following formula:

[0084]

[0085] Where: D9 is the diameter of the fan rotor in millimeters and is measured in a plane perpendicular to the axis of rotation at the intersection between the tip and the leading edge of the fan rotor blades; BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level; T e is the temperature at the inlet of the drive turbine when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level, and is expressed in degrees Celsius (°C); T ref = 273K; n is the number of stages of the additional turbine and the additional compressor; gamma is the adiabatic coefficient of air; and K = 6.76, L = 153.6m -1 .(℃) -1 / 2 .(tr / min) -1 , M=421mm.(℃) -1 / 2 and N2 = -21 millimeters (mm);

[0086] - the propulsion system further comprises an additional turbine configured to drive the compressor via an additional shaft, the additional shaft configured to rotate about the axis of rotation at a speed higher than the speed of the drive shaft, the mean radius of the bore of the additional turbine being at least equal to:

[0087] R m_a ≥G*FN*BPR*10 -4 +H

[0088] Where: R m_a is the mean radius of the bore of the additional turbine, expressed in millimeters (mm); FN is the thrust of the fan rotor measured when the propulsion system is at rest, at takeoff rating, at standard atmospheric pressure and at sea level, and is expressed in Newtons (N); BPR is the bypass ratio of the propulsion system, and is measured when the propulsion system is at rest, at takeoff rating, at standard atmospheric pressure and at sea level; and G = 0.16 millimeters per Newton (mm / N) and H = 18 millimeters (mm); and / or

[0089] - the propulsion system further comprises an additional turbine configured to drive the compressor via an additional shaft, the additional shaft configured to rotate about the axis of rotation at a speed higher than the speed of the drive shaft, the average radius of the bore of the high-pressure turbine being at most equal to:

[0090] R m_a ≤I*FN*BPR*10 -4 +J

[0091] Where: R m_a is the mean radius of the hole of the attached turbine, expressed in millimeters (mm); FN is the thrust of the fan rotor measured when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level, and is expressed in Newtons (N); BPR is the bypass ratio of the propulsion system, and is measured when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level; and I = 0.16 millimeters / Newton (mm / N) and J = 28 millimeters.

[0092] According to a fourth aspect, a method for manufacturing an aviation propulsion system is provided, comprising the following steps:

[0093] - determining the dimensions of the aerospace propulsion system according to the dimensioning method of the third aspect; and

[0094] -Manufacturing of aviation propulsion systems. BRIEF DESCRIPTION OF THE DRAWINGS

[0095] Other features, objects and advantages will appear from the following description which is merely illustrative and non-limiting and which should be read with reference to the accompanying drawings, in which:

[0096] Figure 1 is a schematic partial cross-sectional view of an example of a propulsion system according to a first embodiment, wherein the fan section is ducted;

[0097] Figure 2 is a schematic partial cross-sectional view of an example of a propulsion system according to a first embodiment, wherein the fan section is unducted;

[0098] Figure 3 is a schematic cross-sectional view of an example of a speed reduction mechanism of a first modification;

[0099] Figure 4 is a schematic cross-sectional view of an example of a speed reduction mechanism of a second modification;

[0100] Figure 5 is an example of an aircraft that may include at least one propulsion system according to the first or second embodiment;

[0101] Figure 6 is a flow chart illustrating an example of steps in a sizing or manufacturing method.

[0102] In all figures, similar elements have the same reference numerals. DETAILED DESCRIPTION

[0103] The propulsion system 1 has a main direction extending along a longitudinal axis X. When the propulsion system 1 is in operation, the propulsion system 1 comprises, from upstream to downstream in the direction of gas flow in the propulsion system 1, a fan section 2 and a body 3 (commonly referred to as a "gas generator"), which comprises compressor sections 4 and 5, a combustion chamber 6, and turbine sections 7 and 8. The propulsion system 1 is here an aircraft propulsion system 1 that is configured to be attached to an aircraft 100 via a pylon (or mast).

[0104] The compressor sections 4 and 5 include a series of stages, each of which includes a movable impeller (rotor) 4a and 5a rotating in front of a fixed impeller (stator) 4b and 5b. The turbine sections 7 and 8 also include a series of stages, each of which includes a fixed impeller (stator) 7b and 8b, with the movable impeller (rotor) 7a and 8a rotating behind the fixed impeller (stator) 7b and 8b.

[0105] In the present application, the axial direction corresponds to the direction of the longitudinal axis X, which corresponds to the rotation of the shaft of the gas generator, and the radial direction is a direction perpendicular to and passing through this axis X. Furthermore, the circumferential (or transverse or tangential) direction corresponds to a direction perpendicular to the longitudinal axis X but not passing through the longitudinal axis X. Unless otherwise specified, the terms inner (inner) and outer (outer) are used with reference to radial directions, respectively, such that an inner part or inner surface of an element is closer to the axis X than an outer part or outer surface of the same element.

[0106] During operation, the air flow F entering the propulsion system 1 is divided into a primary air flow F1 and a secondary air flow F2 , and the primary air flow F1 and the secondary air flow F2 circulate in the propulsion system 1 from upstream to downstream.

[0107] A secondary air flow F2 (also called "bypass air flow") flows around the body 3. The secondary air flow F2 is used to cool the periphery of the body 3 and to generate most of the thrust provided by the propulsion system 1.

[0108] The primary air flow F1 flows through the primary air path within the main body 3, thereby passing through the compressor sections 4 and 5, the combustion chamber 6, and the turbine sections 7 and 8 in sequence. In the combustion chamber 6, the primary air flow F1 is mixed with fuel to be used as a combustion product. The primary air flow F1 flows through the turbine sections 7 and 8, which receive energy from the combustion chamber 6. This causes the rotors of the turbine sections 7 and 8 to rotate. The rotation of the rotors of the turbine sections 7 and 8 in turn causes the rotors of the compressor sections 4 and 5 and the rotor portion 9 of the fan section 2 to rotate.

[0109] In a two-body propulsion system 1, compressor sections 4 and 5 may include a low-pressure compressor 4 and a high-pressure compressor 5. Turbine sections 7 and 8 may include a high-pressure turbine 7 and a low-pressure turbine 8. The rotor of high-pressure turbine 7 rotationally drives the rotor of high-pressure compressor 5 via a high-pressure shaft 10. The rotor of low-pressure turbine 8 rotationally drives the rotor of low-pressure compressor 4 and the rotor portion 9 of fan section 2 via a low-pressure shaft 11. Thus, main body 3 includes a high-pressure main body, which includes high-pressure compressor 5, high-pressure turbine 7, and high-pressure shaft 10, and a low-pressure main body, which includes fan section 2, low-pressure compressor 4, high-pressure turbine 8, and low-pressure shaft 11. The rotational speed of the high-pressure main body is greater than that of the low-pressure main body. In a three-body propulsion system 1, turbine sections 7 and 8 also include an intermediate turbine, located between high-pressure turbine 7 and low-pressure turbine 8 and configured to drive the rotor of low-pressure compressor 4 via the intermediate shaft. The fan rotor 9 and the rotor of high-pressure compressor 5 are still driven by the low-pressure shaft 11 and high-pressure shaft 10, respectively.

[0110] The low-pressure shaft 11 is typically housed within the high-pressure shaft 10 over a portion of its length and is coaxial with the high-pressure shaft 10. The low-pressure shaft 11 and the high-pressure shaft 10 can rotate in the same direction, i.e., they can be driven in the same direction about the longitudinal axis X. In a variant, the low-pressure shaft 11 and the high-pressure shaft rotate in opposite directions, i.e., they can be driven in opposite directions about the longitudinal axis X. Optionally, an 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 can rotate in the same direction or in opposite directions.

[0111] The fan section 2 comprises at least a fan rotor 9 adapted to be driven in rotation relative to the stator portion of the propulsion system 1 by the turbine sections 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 setting. In this case, the root of each of the blades 14 of each rotor 9 may be pivotally mounted along a setting axis and connected to a pitch changing mechanism 15 installed in the propulsion system 1, the setting being adjusted by the pitch changing mechanism 15 depending on the flight phase. The pitch changing mechanism 15 is arranged in a manner such that the blades 14 of the rotor 9 are ... blades 14 of the rotor 9 may be fixed relative to the hub 13 or have a variable setting. In this case, the root of each of the blades 14 of the rotor 9 may be pivotally mounted along a setting axis and connected to a pitch changing mechanism 15 installed in the propulsion system 1, the setting being adjusted by the pitch changing mechanism 15 depending on the flight phase. Figure 1 Shown in dashed lines to illustrate that this feature is optional.

[0112] The fan section 2 may also include a fan stator 16 or straightener, which includes blades 17 mounted on the hub 16 of the fan stator and has the function of guiding the secondary air flow F2 flowing at the outlet of the fan rotor 9. The blades 17 of the fan rotor 18 may be fixed relative to the hub or have a variable setting. Similar to the blades 14 of the rotor, the root of each of the blades 17 of the stator may be pivotally mounted along the setting axis X and connected to a pitch changing mechanism 15a, which is generally different from the pitch changing mechanism 15 of the fan rotor 9, and the setting is adjusted by the pitch changing mechanism 15a according to the flight phase.

[0113] The diameter D9 of the fan rotor 9 can be between 2.032 meters (m) (80 inches) and 4.699 meters (m) (185 inches), inclusive, which makes it possible to obtain thrust within the above range. When the fan rotor 9 is ducted, the diameter D9 is, for example, between 2.159 meters (m) (85 inches) and 3.048 meters (m) (120 inches), inclusive, for example, equal to 2.286 meters (m) (90 inches), which makes it possible to conventionally integrate the propulsion system 1, in particular, under the wing of an aircraft. When the fan rotor 9 is non-ducted, the diameter D9 is, for example, greater than or equal to 2.54 meters (m) (100 inches), for example, between 3.048 meters (m) (120 inches) and 3.962 meters (m) (156 inches). The diameter D9 of the fan rotor 9 is measured in a plane perpendicular to the axis of rotation X at the intersection between the tip 21 and the leading edge 22 of the blades 14 of the fan rotor 9 and is expressed in meters. Figure 1 and Figure 2 If this is a partial view, diameter D9 is only partially visible.

[0114] Furthermore, the fan rotor 9 comprises at least twelve blades 14 and at most twenty-four blades 14, for example at least sixteen blades 14 and at most twenty-two blades 14. The number of blades 16 in the fan stator 17 depends on the acoustic standards defined for the propulsion system 1 and is at least equal to the number of blades 14.

[0115] To improve the propulsion efficiency of the propulsion system 1 and reduce its specific fuel consumption and the noise emitted by the fan section 2, the propulsion system 1 has a high bypass ratio. The term "high bypass ratio" is understood herein to mean 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 at rest, not installed, at takeoff rating, at standard atmospheric pressure (as defined by the 3rd edition of the International Civil Aviation Organization (OACI) Manual, Document 7488 / 3), and at sea level conditions (referred to as sea level standard (SLS) conditions). Note that in this application, parameters (pressure, flow, thrust, speed, etc.) are systematically determined under these conditions. The term "not installed" is understood herein to mean that measurements are taken when the propulsion system 1 is in a test stand (and not installed on the aircraft 100), where such measurements are more easily performed. On the other hand, when the propulsion system 1 is cooled, that is, when the propulsion system 1 has been stopped long enough so that the components of the propulsion system are at ambient temperature, distances (lengths, radii, diameters, etc.) are measured at ambient temperature (approximately 20°C), and it should be understood that these dimensions vary very little compared to the conditions when the propulsion system 1 is at take-off rating.

[0116] The fan rotor 9 is separated from the low-pressure shaft 11 using a reduction gear 19, which is arranged between the upstream end of the low-pressure shaft 11 and the fan rotor 9, so as to independently optimize the respective rotational speeds of the fan rotor 9 and the low-pressure shaft 11. In this case, the propulsion system 1 also includes an additional shaft (i.e., a so-called fan shaft 20). The low-pressure shaft 11 connects the low-pressure turbine 8 to the inlet of the reduction gear 19, while the fan shaft 20 connects the output of the reduction gear 19 to the fan rotor 9. Therefore, the fan rotor 9 is driven by the low-pressure shaft 11 via the reduction gear 19 and the fan shaft 20 at a speed lower than the speed of the low-pressure turbine 8.

[0117] This separation enables a reduction in the rotational speed and pressure ratio of the fan rotor 9 and an increase in the power extracted by the high-pressure turbine 8. Specifically, the overall efficiency of the propulsion system is the propulsive efficiency in first-order terms, which favorably influences propulsive efficiency by minimizing the change in the kinetic energy of the air passing through the propulsion system 1. In a propulsion system 1 with a high bypass ratio, the majority of the flow that generates propulsion is formed by the secondary air flow F2 of the propulsion system 1, the kinetic energy of which is primarily influenced by the compression experienced by the secondary air flow F2 as it passes through the fan section 2. Therefore, propulsive efficiency and the pressure ratio of the fan section 2 are related: the lower the pressure ratio of the fan section 2, the higher the propulsive efficiency. To optimize the propulsive efficiency of the propulsion system 1, the fan pressure ratio (corresponding to the ratio of the average pressure at the outlet of the fan stator 17 (or, in the absence of a stator, the fan rotor 9) to the average pressure at the inlet of the fan rotor 9) is less than or equal to 1.70, for example, less than or equal to 1.50, for example, between 1.05 and 1.45. Here, the average pressure is measured along the height of the blade 14 (surface delimiting the air flow path radially inwardly at the inlet of the fan rotor 9 to the tip 21 of the fan blade 14 ).

[0118] The propulsion system 1 is configured to provide a thrust between 18,000 lbf (80,068 N) and 51,000 lbf (222,411 N), such as between 20,000 lbf (88,964 N) and 35,000 lbf (155,688 N).

[0119] The fan section 2 may be of ducted or unducted type. In the case of a ducted fan section 2, the fan section 2 comprises a fan casing 12 in which the fan rotor 9 is accommodated.

[0120] The ducted fan section 2 includes a fan rotor 9 extending upstream of the fan stator 16. The blades of the fan stator are generally referred to as outlet guide vanes (OGV) and have a fixed setting relative to the hub of the fan stator. In addition, the bypass ratio of the propulsion system 1 is, for example, greater than or equal to 10, for example, between 10 and 35 (including 10 and 35), for example, between 10 and 18 (including 10 and 18). In addition, the circumferential speed at the tip 21 of the blades of the fan rotor 9 can be between 260m / s and 400m / s. The blades 14 of the fan rotor 9 can be fixed or have a variable setting. The pressure ratio of the fan can be between 1.20 and 1.45.

[0121] In the unducted fan section 2, the fan section 2 (which can also be called a propeller) is not surrounded by a fan casing. The fan section 2 is unducted, the blades 14 of the fan rotor 9 having a variable setting. The propulsion system comprising at least one unducted fan rotor 9 is called an “open rotor” or an “unducted fan”. The propulsion system 1 can comprise two unducted and contra-rotating fan rotors 9. Such a propulsion system 1 is called a “Contra-Rotating Open Rotor” (CROR) or an “Unducted Double Fan” (UDF). The fan rotor(s) 9 can be arranged at the rear of the main body 3 to form a pusher-type structure, or at the front of the main body 3 to form a tractor-type structure. In a variant, the propulsion system 1 can comprise a single unducted fan rotor 9 and an unducted fan stator 16 (straightener). Such a propulsion system 1 is called an “Unducted Single Fan” (USF). In the case of a USF-type propulsion system 1, the blades 17 of the straightener 16 are rotationally fixed relative to the rotation axis X of the upstream fan rotor 9, and therefore the blades 17 of the straightener 16 are not subjected to any centrifugal force. The blades 17 of the straightener 16 also have a variable setting.

[0122] The elimination of the duct surrounding the fan section 2 makes it possible to very significantly increase the bypass ratio, without the propulsion system 1 being adversely affected by the mass of a casing or nacelle intended to surround the fan section 2. Thus, the bypass ratio of the propulsion system 1 comprising the unducted fan section 2 is greater than or equal to 40, for example between 40 and 80 inclusive. Furthermore, the peripheral speed at the tip 21 of the blades 14 of the fan rotor(s) 9 can be between 210 m / s and 260 m / s. The pressure ratio of the fan can for example be between 1.05 and 1.20.

[0123] The reduction gear 19 can for example comprise a reduction gear 19 with a planetary gear train, for example a reduction gear 19 of one or two stages of the “planetary” or “star” type. According to a first variant, the reduction gear 19 can be of the “star” type, comprising a sun gear 22, a ring gear 23 and a plurality of planet gears 24. According to a second variant, the reduction gear 19 can be of the “planetary” type, comprising a sun gear 22, a ring gear 23 and a plurality of planet gears 24. According to a third variant, the reduction gear 19 can be of the “star” type, comprising a sun gear 22, a ring gear 23 and a plurality of planet gears 24. According to a fourth variant, the reduction gear 19 can be of the “planetary” type, comprising a sun gear 22, a ring gear 23 and a plurality of planet gears 24. Figure 3) and comprises a sun gear 19a (input of the reduction gear 19), a ring gear 19b (output of the reduction gear 19) and a series of satellite portions (planet wheels) 19c; the sun gear 19a is centred on the rotation axis X of the reduction gear 19 (substantially collinear with the longitudinal axis X) and can be configured to be driven in rotation by the low-pressure shaft 11 ; the ring gear 19b is coaxial with the sun gear 19a and is configured to drive the fan shaft 20 in rotation about the rotation axis X; the series of planetary portions 19c is distributed circumferentially between the sun gear 19a and the ring gear 19b about the rotation axis X, each satellite portion 19c being engaged internally with the sun gear 19a and externally with the ring gear 19b. This series of satellite portions 19c is mounted on a planet carrier 19d which is fixed with respect to the stator portion 19e of the propulsion system 1, for example with respect to the casing of the compressor section 4, 5. In a second variant, the reduction gear 19 can be epicyclic Figure 4 ) in which case the ring gear 19b is fixedly mounted on the stator portion 19e of the propulsion system 1 and the fan shaft 20 is driven in rotation by the planet carrier 19d (which can therefore move in rotation with respect to the stator portion 19e of the propulsion system 1, for example with respect to the casing of the compressor section 4, 5).

[0124] Regardless of the configuration of the reduction gear 19, the diameter of the ring gear 19b and of the planet carrier 19d is 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.

[0125] The reduction ratio of the reduction gear 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 ducted fan rotor 9, the reduction ratio can be greater than or equal to 6.0, typically of the order of 3.0. In the case of a propulsion system 1 comprising an unducted fan rotor, the reduction ratio can be comprised between 9.0 and 11.0.

[0126] The low-pressure shaft 11 is supported by three or four bearings 11a, 11b, and / or 11c to control its deformation mode. Thus, the low-pressure shaft 11 may include one or two forward bearings 11a extending upstream of the combustion chamber 6, and two aft bearings 11b, 11c extending downstream of the combustion chamber 6. The bearings 11a, 11b, 11c comprise an inner ring mounted on the low-pressure shaft 11 and an outer ring mounted on the stator portion of the propulsion system 1, typically on the housing of the propulsion system 1, through which forces are transmitted through the propulsion system 1. Thus, the first forward bearing 11a may be mounted on the low-pressure shaft 11, on one hand, and on the inlet casing 26 of the propulsion system 1, which extends between the fan rotor 9 and the low-pressure compressor 4, on the other hand. Where applicable, the second forward bearing may be mounted on the low-pressure shaft 11, on the other hand, and on the inter-compressor casing 23 (or intermediate casing) or inlet casing 26 of the propulsion system 1, or between the low-pressure compressor 4 and the high-pressure compressor 5. The first rear bearing 11b can be mounted on the low-pressure shaft 11 on the one hand and on the inter-turbine casing 24 upstream of the low-pressure turbine 8 (i.e., the casing extending between the high-pressure turbine 7 and the low-pressure turbine 8) on the other hand. In a variant, the first rear bearing 11b can be mounted on the exhaust casing 27, which extends downstream of the low-pressure turbine 8. The first rear bearing 11b extends downstream of the bearing 12b, at the extreme downstream of the high-pressure shaft 10. The second rear bearing 11c can be mounted on the exhaust casing 27 or, if integration is possible, on the inter-turbine casing. Where applicable, the first rear casing 11b and the second rear casing 11c can be mounted on the same cylindrical casing, which is itself attached to the exhaust casing 27.

[0127] In one form of embodiment, the low-pressure shaft 11 is supported by one front bearing 11 a and two rear bearings 11 b , 11 c .

[0128] The twin-body propulsion system 1 may particularly include a two-stage high-pressure turbine 7 , a high-pressure compressor 5 of at least eight stages and at most eleven stages, a low-pressure turbine 8 of at least three stages and at most five stages, and a low-pressure compressor 4 of at least two stages and at most four stages.

[0129] Compared to direct-drive engines (in which the fan shaft 20 is directly driven by the low-pressure shaft 11 and rotates at the same speed as the low-pressure shaft 11), the presence of a speed reduction mechanism 19 between the low-pressure shaft 11 and the fan shaft 20 significantly increases the speed of the low-pressure body (which includes the low-pressure turbine 8, the low-pressure shaft 11, and, in the case of a dual-body propulsion system, the low-pressure compressor 4, where applicable) and reduces the number of stages in the low-pressure turbine 8. This reduces the radial and axial volume of the low-pressure turbine 8, as well as its mass, improving the specific fuel consumption of the propulsion system 1. Simultaneously, reducing the fan's pressure ratio reduces the amount of energy required to drive the fan rotor 9, allowing the flow rate at the inlet of the high-pressure compressor 5, and therefore the inlet section of the high-pressure compressor 5, to be reduced. However, this does have the consequence of limiting the space available for the low-pressure shaft 11 (since it is housed within the high-pressure shaft 10), thus contributing to a reduction in the diameter of the low-pressure shaft 11.

[0130] These modifications to the propulsion system 1 have the effect of altering the deformation pattern of the low-pressure shaft 11. Consequently, the low-pressure shaft 11 is more susceptible to exceeding the limit speed and entering resonance within the operating range of the low-pressure body. During this resonance, which occurs when the low-pressure shaft 11 transitions to its limit speed, the low-pressure shaft 11 experiences an overvoltage phenomenon that amplifies the deformations and forces caused by the (unavoidable) imbalances of the low-pressure body.

[0131] In order to control the dynamic behavior of the low-pressure shaft 11 and move the deformation mode of the low-pressure shaft 11 into the operating range of the low-pressure body that is not likely to damage the propulsion system 1, the bearings 11a, 11b, 11c of the low-pressure shaft 11 are arranged radially and axially so as to meet the following formula:

[0132]

[0133] Where: d2 is the center of gravity G of the rear bearings 11b and 11c of the low-pressure shaft 11 11b and G 11c The distance between them, in millimeters (mm);

[0134] R m is the average radius of the low-pressure turbine 8, in millimeters (mm);

[0135] XN is the limit speed of the low-pressure shaft (11), in revolutions per minute (tr / min);

[0136] d1 is the center of gravity G of the front bearing 11a 11a The center of gravity G of the rear bearing 11b closest to the front bearing 11a 11b The distance between them, in millimeters (mm);

[0137] R 11bis the average radius of the rear bearing 11 b closest to the front bearing 11 a , in millimeters (mm); and

[0138] C=170(mm*tr / min) -1 , E1=0.18 (dimensionless).

[0139] d1, d2, R m 、R 11b and D9 is the distance, as described above, these parameters are determined when the propulsion system 1 (and fan section 2) are cold.

[0140] The average radius R of the low-pressure turbine 8 m This corresponds to the average value of the average radius of the rotor 8a (movable blade wheel) of the low-pressure turbine 8. In a given stage, the average radius R1 of the rotor 8a is equal to the average value between the outer radius of the rotor 8a and the inner radius of the rotor 8a, wherein, when the propulsion system 1 is cold, the outer radius and the inner radius are measured between the axis of rotation X and the inner radial surface of the rotor 8a (which radially delimits the blades of the rotor 8a on the inside). Note that both the outer radius and the inner radius are measured in a plane perpendicular to the axis of rotation X of the low-pressure turbine 8, at the mid-distance between the leading edge 8c at the tip of the movable blade and the trailing edge 8d at the tip of the movable blade (i.e., at 50% of the chord at the blade tip), between the axis of rotation X and the outer radial surface of the rotor 8a (which radially delimits the blades of the rotor 8a on the outside).

[0141] Average radius R of rear bearing 11b 11b (or the average radius R of the front bearing 11a 11a ) corresponds to the axis X and the center of gravity G closest to the front bearing 11a, that is, the rear bearing 11b most upstream 11b (or the center of gravity G of the rear bearing 11a 11a ), the distance being measured in a plane perpendicular to the longitudinal axis X. Note here that the bearing comprises an outer ring and an inner ring, which are coaxial, and a bearing (ball, roller, etc.), the bearing being mounted between the inner ring and the outer ring and being configured so that the inner ring can move relative to the outer ring. The inner ring is mounted on the low-pressure shaft 11 and is rotationally driven by the low-pressure shaft 11; the outer ring is mounted on the stator part of the propulsion system 1 (inlet casing 26, compressor inter-casing 23, turbine inter-casing 24, exhaust casing 27, etc.). The center of gravity of the bearing corresponds to the center of gravity of the assembly formed by the inner ring, the outer ring and the bearing. When the inner ring and / or the outer ring are integrally formed as one piece with a support (flange, housing, enclosure, etc.) of the propulsion system 1 so that the ring cannot be distinguished from the support, the center of gravity is determined without taking into account the ring in question.

[0142] When the low-pressure shaft is supported by the two front bearings upstream of the combustion chamber, the distance d1 is measured between the front bearing 11a closest to the combustion chamber 6 (i.e., the most downstream front bearing 11a) and the rear bearing 11b closest to the combustion chamber 6 (i.e., the most upstream rear bearing 11b). In other words, the distance d1 corresponds to the minimum distance between the center of gravity of the front bearing 11a and the center of gravity of the rear bearing 11b.

[0143] When the low-pressure shaft 11 is supported by the two rear bearings 11b and 11c (and one or two front bearings 11a), the center of gravity G of the two bearings 11b and 11c is 11b and G 11c When the low-pressure shaft 11 is supported by exactly three rear bearings in addition to the one or more front bearings 11 a , the distance d2 is measured between the most upstream bearing and the most downstream bearing of the rear bearings.

[0144] The limit speed XN corresponds to the absolute maximum speed that the low-pressure shaft 11 may encounter throughout the entire flight (according to the European certification regulation EASACS-E 740 (or according to the US certification regulation 14-CFR Part 33)). The limit speed corresponds to the maximum rotational speed when the propulsion system is healthy (and possibly at the end of its life). Therefore, the low-pressure shaft 11 is likely to reach the limit speed under flight conditions. This limit speed is part of the data stated in the type certificate data sheet. Specifically, this rotational speed is often used as a reference speed for determining the size of the propulsion system 1 and in certain certification tests (such as blade loss or rotor integrity tests).

[0145] The applicant has noticed that the position of the bearings 11a-11c of the low-pressure shaft 11 has an impact on the rigidity of the fixation of the low-pressure shaft 11. In particular, the closer the rear bearings 11a-11c are to each other (the smaller d2), the greater the rigidity of the fixation of the low-pressure shaft 11, the lower the rigidity of the low-pressure shaft 11, and the smaller the deformation mode of the low-pressure shaft 11. Therefore, these parameters are more likely to be within the operating range of the low-pressure body. Similarly, the larger the diameter of the shaft (the diameter is the same as the average radius R of the rear bearings), the greater the rigidity of the low-pressure shaft 11. 11b(proportional to) and / or the distance (d1) between the front bearing and the rear bearing is smaller, the deformation mode is smaller. Therefore, when the positions of the bearings 11a-11c comply with formula (1), the stiffness of the low-pressure shaft 11 is increased to be sufficient to fix the low-pressure shaft 11 more rigidly. Typically, for a propulsion system including a reduction mechanism 19, the first deformation mode of the low-pressure shaft 11 occurs at a transient speed corresponding to, for example, a speed less than idle speed, and is therefore not prone to damage to the propulsion system 1. Therefore, a transient speed that occurs only for a very short time during the operation of the propulsion system 1 is selected. For example, the size of the propulsion system is determined so that the limit speed corresponds to a speed that occurs only at the beginning of startup or at the end of shutdown of the propulsion system 1, such as a speed less than or greater than idle speed but outside the normal operating range or within an operating range that is rarely used and short-lived during the operation of the propulsion system 1. Therefore, the low-pressure shaft 11 can be supercritical while having a controlled dynamic condition.

[0146] The limit speed XN of the low-pressure shaft 11 can be varied by acting on the following elements of the propulsion system 1 :

[0147] 1. Shaft geometry (outer and inner diameters). The outer diameter of the low-pressure shaft influences the design of the high-pressure body, as the low-pressure shaft 11 passes through the high-pressure shaft 10 and limits the bore radius of the high-pressure turbine 7. Furthermore, variations in the internal shape of the low-pressure shaft 11 that increase the limiting speed are limited by the manufacturing constraints of the low-pressure shaft 11.

[0148] 2. The material of the low-pressure shaft 11 (the influence of the Young's modulus / density ratio). Typically, steel with a low Young's modulus variation is used. Alternatively, the low-pressure shaft 11 can be made of a metal matrix composite material, which has a significantly higher Young's modulus / density ratio but a lower interfacial torque transmission capacity than conventional steel.

[0149] 3. The number and location of the bearings 11a-11d of the low-pressure shaft 11. Specifically, the configuration in which the low-pressure shaft 11 is supported by four bearings increases the limit speed XN of the low-pressure shaft 11. However, compared to a three-bearing configuration, the additional bearings do increase integration difficulties. Where applicable, the rear bearing of the low-pressure shaft 11 can be integrated upstream of the low-pressure turbine 8, which increases the limit speed XN of the low-pressure shaft 11 at the expense of engine length.

[0150] 4. The rigidity of the trunnions between the rotors of the low-pressure body. Specifically, increasing the rigidity of the trunnions enables an increase in the limit speed XN of the low-pressure shaft 11. However, increasing the rigidity of the trunnions also increases the difficulty of production and integration.

[0151] In the case of a propulsion system with a high bypass ratio (BPR) and a high overall pressure ratio (OPR), the distance between the outer diameter of the low-pressure shaft 11 and the low-pressure bearing is severely limited by the integration of the high-pressure body, which leads to supercritical dynamic conditions that must be controlled. In order to avoid the length loss associated with the integration of the bearing for the low-pressure shaft 11 upstream of the low-pressure turbine, the distance d2 between the centers of gravity of the two rear bearings 11b, 11c of the low-pressure shaft 11 is adjusted according to formulas (1) and (2) so that the desired dynamic conditions can be obtained. In addition, the dimensioning of the limit speed XN is determined to comply with formulas (3) and (4) (see below) so that the integration of the high-pressure body can be optimized while remaining within the accessible indicators of the other parameters mentioned above.

[0152] Preferably, the bearings 11a, 11b, 11c of the low-pressure shaft 11 are arranged radially and axially so as to meet the following formula:

[0153]

[0154] Where: E2 = 0.378 (dimensionless).

[0155] When the position of the bearings 11a-11c complies with formula (2), the bearings are kept close enough to enable the bearings to be integrated below the low-pressure turbine 8 and / or exhaust casing 27 of the propulsion system 1 and limit the axial length of the low-pressure body and therefore the weight and volume of the propulsion system 1.

[0156] Preferably, the propulsion system 11 is dimensioned so that the limiting speed of the drive shaft 11 complies with the following formula:

[0157]

[0158] Where: A = 5500mm*tr / min.

[0159] The limit speed XN of the low-pressure shaft 11 may be, for example, between 8500 rpm and 12 000 rpm, preferably between 9000 rpm and 11 000 rpm.

[0160] When the limit speed XN complies with formula (3) and the fixed stiffness of the bearings 11a-11c (i.e. the radial and axial positions of the bearings 11a-11c) complies with formula (1) and, where applicable, formula (2), the limit speed is high enough to ensure that the first deformation mode of the low-pressure shaft 11 will be within the speed range of the low-pressure shaft 11, within which the dynamic conditions of the low-pressure shaft 11 can be controlled.

[0161] Preferably, the limit speed XN also conforms to the following formula:

[0162]

[0163] Where: B = 8500mm*tr / min.

[0164] When the low-pressure shaft 11 does not comply with the formula (4), the limit speed of the low-pressure shaft 11 becomes too high and causes an unfavorable dynamic condition of the turbine shaft.

[0165] Where applicable, to further improve the control of the dynamic conditions and vibration levels of the low-pressure shaft 11, the bearings 11a-11c of the low-pressure shaft 11 include bearing mode dampers 25, which include a pressurized oil film configured to damp mode transitions (viscous damping). This type of damping is commonly referred to as a "squeeze film." All or part of the bearings 11a-11c may also include a flexible cage mounted between an outer ring and a rigid support member fixedly secured to a stator portion of the propulsion system 1 (inlet casing 26, compressor inter-casing 23, turbine inter-casing 24, exhaust casing 27, etc.) and radially deformable studs mounted between the rigid support member and the outer ring of the bearing to allow radial displacement of the substantially cylindrical wall, and therefore the bearings 11a-11c, relative to the rigid support member. Documents WO 2021 / 001610 and WO 2022 / 195198 in the name of the applicant describe an example of a bearing having a flexible cage and a pressurized fluid film damper including the flexible cage. Thus, these bearing modal dampers 25 damp the response of the low-pressure shaft 11, thereby limiting the resonance of the low-pressure shaft when its frequency reaches the first deformation mode.

[0166] Preferably, the bearings 11 a - 11 c of the low-pressure shaft 11 include exactly three bearings, namely a front bearing 11 a and two rear bearings 11 b , 11 c .

[0167] Note that formula (1) is applicable to any type of bearing. As an example, the front bearing 11a may include a ball bearing, a tapered roller bearing, or an angular contact ball bearing, and the rear bearings 11b, 11c may each include a roller bearing.

[0168] The low-pressure shaft 11 is housed in the high-pressure shaft 10 , the diameter of which is determined by the mean radius R of the bore of the high-pressure turbine 7 . m_aHowever, the larger the diameter of the low-pressure shaft 11, the more modes of the low-pressure shaft 11 appear at high rotational speeds (or in other words, the more the deformation mode of the shaft shifts to high frequencies). In order to obtain a compromise between the average bore radius of the high-pressure turbine 7 with a small diameter (which makes it possible to improve the total pressure ratio of the propulsion system and the bypass ratio BPR of the propulsion system) and to control the dynamic conditions of the low-pressure shaft 11, in particular when the fan section 2 is of the bypass type, the average radius R of the bore of the high-pressure turbine 7 is preferably 0. m a Complies with the following formula:

[0169] R m_a ≤E*L HP 2 *XN*10 -9 +F1 (5)

[0170] Where: R m_a Expressed in millimeters (mm);

[0171] L HP is the length of the high-pressure body in millimeters (mm);

[0172] E=3.15(mm.tr / min) -1 And F1 = 23 millimeters (mm).

[0173] Specifically, as described above, the deformation frequency of the low-pressure shaft 11 is proportional to the square of the diameter of the low-pressure shaft 11 and the distance between the bearings 11a-11b of the low-pressure shaft 11 (d1 2 ) is proportional to the ratio of the bearings 11a-11b supporting the low-pressure shaft 11. Therefore, the further apart the bearings 11a-11b supporting the low-pressure shaft 11 are, the lower the deformation frequency (and deformation mode) of the low-pressure shaft 11. However, the position of the bearings 11a-11b of the low-pressure shaft 11 depends on the length of the high-pressure body. By dividing the average radius R of the hole of the high-pressure turbine 7 by m_a Defined as the high pressure body L HP (High-pressure main body L HP formed by the high-pressure compressor 5 and the high-pressure turbine 7), the dynamic behavior of the low-pressure shaft 11 can therefore be better controlled.

[0174] Preferably, the average radius R of the holes of each disk of the high-pressure turbine 7 is m_a It complies with formula (5).

[0175] The average radius R of the hole of the high-pressure turbine 7 m_a The propulsion system 1 that complies with formula (5) can achieve a total pressure 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 root of the fan rotor 9), and the total pressure ratio is 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.

[0176] The rotation speed of the high pressure turbine 7 can be between 15000 rpm and 27000 rpm. When the rotation speed of the high pressure turbine (and the high pressure shaft 10) is included in this interval and the average radius R of the high pressure turbine is m_a When the formula (4) is met, the mechanical load of the rotor 7a is acceptable for the high-pressure turbine 7.

[0177] For a high pressure main body including 10 to 13 compressor stages and turbine stages, the high pressure main body L HP The length of can be between 950mm and 1450mm, for example.

[0178] In particular, when the fan section 2 is of the ducted type, the average radius R of the bore of the high-pressure turbine 7 is m_a It also conforms to the following formula:

[0179] R m_a ≥E*L HP 2 *XN*10 -9 +F2 (6)

[0180] Where: F2 = 13 millimeters (mm).

[0181] Sizing the mean radius of the holes to comply with formula (6) ensures a minimum radius of the high-pressure turbine and therefore ensures the ability of the high-pressure turbine to withstand centrifugal forces.

[0182] In some cases, it is possible to determine the high pressure body L HP Before determining the length of the hole R of the high pressure turbine 7 m_a In this case, the average radius of the hole can be determined based on basic dimensional parameters: the diameter D9 of the fan rotor 9, the bypass ratio BPR of the propulsion system 1 or the limit speed XN of the low-pressure shaft 11, instead of the length of the low-pressure body (as defined in formulas (5) and (6)):

[0183]

[0184] Where: BPR is the bypass ratio of the propulsion system;

[0185] T e is the temperature at the inlet of the low-pressure turbine 8 in degrees Celsius;

[0186] T ref =273K;

[0187] n is the number of stages of the high-pressure body (i.e., the high-pressure turbine 7 and the high-pressure compressor 5);

[0188] Gamma is the adiabatic coefficient of air;

[0189] K = 6.76 and L = 153.6 m -1 (°C) -1 / 2 (tr / min) -1 (°C) -1 / 2 and N1 = -11 millimeters (mm).

[0190] As specified in equation (6), the average radius R m_a of the holes of the high-pressure turbine 7 complies with equation (7) when

[0191] Preferably, the average radius R m_a of the holes of each disc of the high-pressure turbine 7 complies with equation (7).

[0192] Moreover, the average radius of the holes also complies with the following equation:

[0193]

[0194] where N2 = -21 millimeters (mm).

[0195] Preferably, the average radius R m_a of the holes of each disc of the high-pressure turbine 7 complies with equation (8).

[0196] Preferably, the dimensions of the propulsion system 1 are also determined so that, in particular when the fan section 2 is ducted, the average radius R m_a of the holes of the high-pressure turbine 7 also complies with the following equation:

[0197] R m_a ≤ I * FN * BPR * 10 -4 + J (9)

[0198] where: FN is the thrust of the fan rotor 9, in Newton (N);

[0199] I = 0.16 millimeters per Newton (mm / N) and J = 28 millimeters (mm).

[0200] In particular, the torque transmitted by the low-pressure shaft 11 to the fan rotor 9 is related to the thrust of the fan rotor 9 and to the duct ratio BPR of the propulsion system 1. Moreover, increasing the duct ratio BPR of the propulsion system 1 makes it possible to reduce the diameter of the high-pressure body and, therefore, the average radius R m_a of the holes of the high-pressure turbine 7. Therefore, by determining the dimensions of the propulsion system so that the average radius R m_aAlso in accordance with formula (9), a propulsion system 1 is obtained, the dimensions of the high-pressure body and the diameter of the low-pressure shaft 11 of which are determined as a function of the thrust of the fan rotor 9 (mainly the total thrust of the propulsion system 1 ) and the bypass ratio BPR.

[0201] Preferably, the average radius R of the holes of each disk of the high-pressure turbine 7 is m_a It complies with formula (9).

[0202] For example, the average radius of the bores of a high-pressure turbine is at most equal to 300 mm.

[0203] In the form of embodiment, in order to maintain the possibility of inserting the low-pressure shaft 11 into and passing through the high-pressure shaft 10, while ensuring the ability of the low-pressure shaft 11 to transmit the torque required by the fan rotor 9 and the low-pressure compressor 4, especially when the fan section 2 is a bypass type, the average radius R of the hole of the second turbine 7 is m_a Complies with the following formula:

[0204] R m_a ≥G*FN*BPR*10 -4 +H (10)

[0205] Where: G = 0.16 mm / Newton (mm / N) and H = 18 mm.

[0206] Preferably, the average radius R of the holes of each disk of the high-pressure turbine 7 is m_a It complies with formula (10).

[0207] Comparison example:

[0208] The engine 1 is a twin-body propulsion system comprising a ducted fan section 2 corresponding to current technical standards (at the filing date of the present application) with a view to improvement.

[0209] The engine 2 is a twin-body propulsion system 1 including a ducted fan section 2 according to the teachings of the present application, the drive shaft (ie, low-pressure shaft) of the twin-body propulsion system 1 being supported by bearings conforming to formulas (1) and (2) defined above.

[0210] Both Engine 1 and Engine 2 have a low-pressure shaft supported by one bearing just upstream of the combustion chamber and two bearings downstream of the combustion chamber. In addition, the front bearing includes damping by a pressurized fluid film.

[0211]

[0212]

[0213]

[0214] The ratio of the mean radius of the low pressure turbine to the bearing-to-bearing distance d2 is greater than the maximum value defined by equation (2) as a function of the limit speed of the low pressure shaft, the bearing-to-bearing distance dl and the radius R of the upstream rear bearing of the low pressure shaft 11b The result is that the first deformation mode of the low pressure shaft in the supercritical regime occurs at transient ratings close to the take-off rating. Thus, despite the bearing mode damper, the vibratory loads experienced during the mode change are high. This also appears from the fact that the limit speed of the low pressure shaft of the engine 1 is less than the minimum limit speed defined by equation (3) as a function of the radius R of the upstream rear bearing of the low pressure shaft 11b and the bearing-to-bearing distance dl.

[0215] By comparison, the ratio of the mean radius of the low pressure turbine to the bearing-to-bearing distance d2 of the engine 2 complies with equation (1) and equation (2). Thus, the stiffness of the low pressure shaft is increased so that the first deformation mode of the low pressure shaft in the supercritical regime occurs closer to the idle transient operating range. The bearing mode damper also enables damping of the mode change with moderate vibratory loads in the engine 2 in this operating range compared to the vibratory loads experienced by the engine 1. Furthermore, since the low pressure shaft is more rigidly fixed, the diameter of the low pressure shaft can be limited despite the fact that the low pressure shaft drives seven stages in total (three compressor stages 4 and four turbine stages 8). However, the bearings supporting the low pressure shaft remain close enough to enable the bearings to be integrated under the low pressure turbine. Finally, despite the greater number of stages, the mean radius of the low pressure turbine 8 of the engine 2 is less than the mean radius of the low pressure turbine 8 of the engine 1 so that the low pressure turbine 8 of the engine 2 is more compact radially than the low pressure turbine of the engine 1.

[0216] Furthermore, the limit speed of the low pressure shaft of the engine 2 is greater than the minimum limit speed defined by equation (3) so that the second deformation mode of the low pressure shaft is outside the operating range of the engine 2. However, the limit speed of the low pressure shaft of the engine 2 remains less than the maximum limit speed defined in equation (4) to limit the vibratory loads experienced during the mode change. Thus, in particular from the mechanical load parameter XN 2 It can be seen in the mechanical stresses in the low pressure shaft and in the low pressure turbine S8 that the mechanical stresses remain acceptable.

[0217] To change from the engine 1 (reference) to the engine 2 (according to the content of the disclosure), the diameter of the fan D9 and the bypass ratio BPR are increased, which enables the propulsive efficiency to be improved and the equivalent fan thrust to be maintained with a drop in the pressure ratio of the fan section 2. Furthermore, the total pressure ratio and the temperature at the inlet of the high pressure turbine 7 are increased, which enables the thermal efficiency of the propulsion system 1 to be increased.

[0218] Furthermore, the outer diameter of the shaft of the low-pressure turbine 11 is reduced in order to simplify the integration of the high-pressure body, in particular by conforming to the mechanical dimensions of the disks of the high-pressure turbine.

[0219] By adjusting the distance d2 between the centers of gravity of the rear bearings of the low-pressure shaft 11 and by optimizing the internal geometry of the low-pressure shaft 11 while complying with the manufacturing stresses of the shaft 11, adjustments are made to obtain the limiting speed of the first deformation mode of the low-pressure shaft 11 within the rated range defined by formulas (1) to (4).

[0220] Taking into account the location of the air path of the low-pressure compressor, the number of bearings supporting the low-pressure shaft remains equal to 3 (and is not increased to 4) due to the integration of the bearing upstream of the low-pressure shaft.

[0221] Solutions of moving one of the two bearings forward or moving both bearings from downstream of the low-pressure turbine to upstream of the low-pressure turbine are also excluded in order to avoid elongation of the engine 2 and misalignment problems between the two bearings.

Claims

1. An aviation propulsion system (1), comprising: - a fan rotor (9) configured to be driven by a fan shaft (20); a drive turbine (8) configured to drive the fan rotor (9) about an axis of rotation (X) via a drive shaft (11); - a speed reduction mechanism (19) coupling the drive shaft (11) and the fan shaft (20) so as to drive the fan shaft (20) at a rotation speed lower than that of the drive shaft (11); - bearings (11a-11c) configured to center the drive shaft (11) relative to the axis of rotation (X), the bearings (11a-11c) comprising a front bearing (11a) extending upstream of a combustion chamber of the propulsion system (1) and two rear bearings (11b, 11c) extending downstream of the combustion chamber, wherein the positions of the bearings (11a-11c) in the propulsion system (1) comply with the following formula: Wherein: d2 is the center of gravity (G 11b and G 11c ) in millimeters (mm); R m is the average radius of the driving turbine (8), in millimeters (mm); XN is the limit speed of the drive shaft (11), in revolutions per minute (tr / min); d1 is the center of gravity of the front bearing (11a) (G 11a ) and the center of gravity (G 11b ) in millimeters (mm); R 11b is the average radius of the rear bearing (11b) closest to the front bearing (11a), in millimeters (mm); and C=170(mm*tr / min) -1 , E1=0.18 and E2=0.

378.

2. The propulsion system (1) according to claim 1, wherein: The limit speed of the drive shaft (11) complies with the following formula: Where: A = 5500mm*tr / min.

3. The propulsion system (1) according to claim 1 or 2, wherein: The limit speed of the drive shaft (11) complies with the following formula: Where: B = 8500mm*tr / min.

4. The propulsion system (1) according to any one of claims 1 to 3, wherein: All or some of the bearings (11a-11c) of the drive shaft (11) are associated with bearing modal dampers (25).

5. The propulsion system according to claim 4, wherein: The bearing mode damper (25) comprises a squeeze film.

6. The propulsion system according to claim 4 or 5, wherein: Each bearing mode damper (25) comprises a deformable cage mounted between the ring of the associated bearing and a stator part (23, 24, 26, 27) of the propulsion system (1).

7. The propulsion system according to any one of claims 1 to 6, wherein: The bearings of the drive shaft (11) include at least one front bearing (11a) and exactly two rear bearings.

8. The propulsion system according to any one of claims 1 to 7, wherein: The propulsion system further comprises an additional turbine (7) configured to drive an additional compressor (5) via an additional shaft (10), the additional shaft (10) being configured to rotate about the axis of rotation (X) at a higher speed than the speed of the drive shaft (11), the additional turbine being a two-stage turbine.

9. The propulsion system (1) according to claim 8, wherein: The additional compressor (5) comprises at least eight stages and at most eleven stages.

10. The propulsion system (1) according to any one of claims 1 to 9, wherein: The driving turbine (8) comprises at least three stages and at most five stages.

11. The propulsion system (1) according to any one of claims 1 to 10, wherein: The reduction ratio of the reduction mechanism is greater than or equal to 2.5 and less than or equal to 11.

12. An aircraft comprising at least one propulsion system (1) according to any one of claims 1 to 11, the propulsion system being attached to the aircraft by means of a pylon.

13. A method for determining the dimensions of a propulsion system (1), the propulsion system comprising a reduction mechanism (19) coupling a drive shaft (11) and a fan rotor (9) so as to drive the fan rotor (9) at a rotational speed lower than that of the drive shaft (11), the bearings being configured to center the drive shaft (11) relative to a rotational axis (X), the bearings (11a-11c) comprising a front bearing extending upstream of a combustion chamber of the propulsion system (1) and two rear bearings extending downstream of the combustion chamber, The dimensions of the propulsion system are determined so that the positions of the bearings (11a-11c) in the propulsion system (1) comply with the following formula: in: d2 is the center of gravity (G 11b and G 11c ) in millimeters (mm); R m is the average radius of the driving turbine (8), in millimeters (mm); XN is the limit speed of the drive shaft (11), in revolutions per minute (tr / min); d1 is the center of gravity of the front bearing (11a) (G 11a ) and the center of gravity (G 11b ) in millimeters (mm); R 11b is the average radius of the rear bearing (11b) closest to the front bearing (11a), in millimeters (mm); and C=170(mm*tr / min) -1 , E1 = -0.18 and E2 = 0.

378.

14. Method for dimensioning a propulsion system (1) according to claim 13, wherein The propulsion system (1) is also dimensioned so that the limiting speed of the drive shaft (11) complies with the following formula: Where: A = 5500mm*tr / min.

15. The size determination method according to claim 13 or 14, wherein: The dimensions of the propulsion system are also determined so that the limiting speed of the drive shaft (11) also complies with the following formula: Where: B = 8500mm*tr / min.

16. A method for manufacturing an aerospace propulsion system, the method comprising the steps of: - Determining the dimensions of the aerospace propulsion system according to the method of any one of claims 13 to 15; as well as - Manufacturing of said aviation propulsion system.

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