Improving dynamic behavior of drive shaft of fan of aviation propulsion system

Abstract

An aviation propulsion system (1) comprising a drive shaft (11) of a fan rotor (9) supported by a front bearing and two rear bearings, the limit speed of the drive shaft satisfying formula (I): wherein: XN is the limit speed of the drive shaft (11) in the unit of rotational speed per minute (tr / min); r is the average radius of the front bearing, and the unit is millimeter (mm); and d1 is the distance between the center of gravity (G11a) of the front bearing (11a) and the center of gravity (G11b) of the rear bearing (11b) closest to the front bearing (11a) in millimeter (mm); and A = 5500 mm * tr / min.

Description

Technical Field

[0001] The present application relates generally to the field of propulsion systems and more particularly to aviation propulsion systems comprising ducted or unducted fans and having a high or even 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 rotated by the high-pressure turbine via a high-pressure shaft. The fan and, where applicable, the low-pressure compressor are rotated by the low-pressure turbine via a low-pressure shaft.

[0003] Technological research has significantly improved the environmental performance of aircraft. Applicants consider influencing factors at all stages of design and development to obtain aviation components and products 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 goal of improving the energy efficiency of aircraft.

[0004] In order to improve the propulsion efficiency of the propulsion system and reduce the specific consumption of the propulsion system and the noise emitted by the fan section, a propulsion system with a high bypass ratio (BPR, the bypass ratio corresponds to the ratio between the secondary air flow and 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. Usually, the separation is achieved by 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] The current trend is to increase the overall pressure ratio of the propulsion system, which corresponds to the ratio between the outlet pressure of the high-pressure compressor and the inlet pressure of the fan. This allows the inlet pressure of the gases in the combustion chamber to be increased, thus further improving the overall efficiency of the propulsion system. Therefore, increasing the overall pressure ratio requires increasing the pressure ratio of the high-pressure compressor and / or the low-pressure compressor, particularly since attempts are being made to reduce the pressure ratio of the fan for the reasons mentioned above.

[0006] One of the results is that the torque transmitted by the low-pressure shaft to the reduction mechanism is reduced compared to a direct-drive propulsion system. Therefore, the low-pressure shaft can have a smaller diameter, which makes it possible to integrate the high-pressure body into the propulsion system. However, reducing the diameter of the low-pressure shaft has the effect of lowering the frequency of the natural modes of the low-pressure shaft. In addition, increasing the rotational speed of the low-pressure shaft expands the operating range of the low-pressure shaft. Therefore, the low-pressure shaft may exceed the critical speed and may resonate (unless design precautions are taken into account). Under resonance, which occurs when the critical speed of the low-pressure shaft is exceeded, the shaft is subjected to overstress phenomena, which amplify the deformations and forces caused by the (unavoidable) imbalance of the low-pressure shaft. Under these conditions, the low-pressure shaft is said to be supercritical.

[0007] Therefore, changing the diameter of the low-pressure shaft has an impact on the dynamic behavior of the low-pressure shaft and changes the location of the supercritical mode of the low-pressure shaft. A propulsion system rotating at a critical speed under steady-state conditions risks rapid degradation, which must be avoided. Summary of the Invention

[0008] One object of the present application is to optimize a propulsion system to increase the efficiency of the propulsion system while controlling the dynamic behavior of the fan rotor drive shaft.

[0009] To this end, an aviation propulsion system is proposed, which includes:

[0010] - a fan rotor, the fan rotor being connected to the fan shaft;

[0011] a drive turbine configured to drive the fan rotor about a rotation axis via a drive shaft;

[0012] - 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;

[0013] - 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,

[0014] Among them, the limit speed of the drive shaft satisfies the following formula:

[0015]

[0016] Where: XN is the limit speed of the drive shaft, in revolutions per minute (tr / min);

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

[0018] 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); and

[0019] A=5500mm*tr / min.

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

[0021] -The limit speed of the drive shaft also satisfies the following formula:

[0022]

[0023] Where: B = 8500mm*tr / min;

[0024] - Each bearing is associated with a bearing mode damper;

[0025] - Bearing mode dampers include damping by a pressurized fluid film;

[0026] - The bearing mode damper comprises a deformable cage mounted between the ring of each bearing and the stator part of the propulsion system;

[0027] - the bearing comprises an additional front bearing extending upstream of the front bearing;

[0028] -The limit speed of the drive shaft also satisfies the following formula:

[0029]

[0030] Where: Rm is the mean radius of the drive turbine in millimeters (mm); d2 is the distance between the centers of gravity of the rear bearings in millimeters (mm); and C = 170 (mm*tr / min) -1 And E1=378.

[0031] -The limit speed of the drive shaft (11) also satisfies the following formula:

[0032]

[0033] Where: Rm is the mean radius of the drive turbine in millimeters (mm); d2 is the distance between the centers of gravity of the rear bearings in millimeters (mm); and C = 170 (mm*tr / min) -1 And E2=180;

[0034] - the bearings comprise at least one front bearing and exactly two rear bearings;

[0035] - the propulsion system further comprises an additional turbine configured to drive the additional 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 additional turbine being a two-stage turbine;

[0036] The average radius of the hole for the additional turbine wheel satisfies the following formula:

[0037] R m_a ≤E*L HP 2 *XN*10 -9 +F1

[0038] Where: R m_a is the average radius of the hole to which the turbine is attached, 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 mm;

[0039] - The average radius of the hole of the additional turbine (7) satisfies the following formula:

[0040] R m_a ≥E*L HP 2 *XN*10 -9 +F2

[0041] Where: R m_a is the average radius of the hole to which the turbine is attached, 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);

[0042] - The average radius of the hole of the additional turbine (7) satisfies the following formula:

[0043]

[0044] 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 speed, standard atmospheric pressure, and sea level; T e is the inlet temperature of the drive turbine when the propulsion system is at rest, takeoff speed, standard atmospheric pressure and sea level, and is expressed in degrees Celsius; 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);

[0045] - The average radius of the hole of the additional turbine (7) satisfies the following formula:

[0046]

[0047] 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 speed, standard atmospheric pressure, and sea level; T e is the inlet temperature of the drive turbine when the propulsion system is at rest, take-off speed, 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);

[0048] - The average radius of the hole of the additional turbine (7) satisfies the following formula:

[0049] R m_a ≥G*FN*BPR*10 -4 +H

[0050] 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 speed, 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 speed, at standard atmospheric pressure, and at sea level; and G = 0.16 millimeters / Newton (mm / N) and H = 18 millimeters (mm);

[0051] - the mean radius of the bore of the high-pressure turbine (7) is at most equal to:

[0052] R m_a ≤I*FN*BPR*10 -4+J

[0053] 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 speed, 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 speed, at standard atmospheric pressure, and at sea level; and I = 0.16 millimeters / Newton (mm / N) and J = 28 millimeters (mm);

[0054] - the mean radius of the holes of the additional impeller is at most equal to 300 mm;

[0055] - the additional compressor (5) comprises at least eight stages and at most eleven stages;

[0056] - the driving turbine (8) comprises at least three stages and at most five stages;

[0057] - The reduction ratio of the reduction mechanism is greater than or equal to 2.5 and less than or equal to 11.

[0058] According to a second aspect, an aircraft is proposed, comprising at least one propulsion system according to the first aspect, the propulsion system being attached to the aircraft via a mast.

[0059] According to a third aspect, a method for determining the dimensions of a propulsion system is provided, the propulsion system comprising a reduction system and bearings, the reduction system 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 front bearing extending upstream of a combustion chamber of the propulsion system, and two rear bearings extending downstream of the combustion chamber,

[0060] The propulsion system is dimensioned so that the limiting speed of the drive shaft satisfies the following formula:

[0061]

[0062] Where: XN is the limit speed of the drive shaft, in revolutions per minute (tr / min);

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

[0064] 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); and

[0065] A=5500mm*tr / min.

[0066] Some preferred but non-limiting features of the sizing method or manufacturing method according to the second aspect are the following features taken alone or in combination:

[0067] The propulsion system is dimensioned so that the limiting speed of the drive shaft also satisfies the following formula:

[0068]

[0069] Where: B = 8500mm*tr / min;

[0070] The propulsion system is dimensioned so that the limiting speed of the drive shaft also satisfies the following formula:

[0071]

[0072] Where: Rm is the mean radius of the drive turbine in millimeters (mm); d2 is the distance between the centers of gravity of the rear bearings in millimeters (mm); and C = 170 (mm*tr / min) -1 And E1=378;

[0073] The propulsion system is dimensioned so that the limiting speed of the drive shaft also satisfies the following formula:

[0074]

[0075] Where: Rm is the mean radius of the drive turbine in millimeters (mm); d2 is the distance between the centers of gravity of the rear bearings in millimeters (mm); and C = 170 (mm*tr / min) -1 And E2=180;

[0076] 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 that of the drive shaft, the average radius of the bore of the additional turbine satisfying the following formula:

[0077] R m_a ≤E*L HP 2 *XN*10 -9 +F1

[0078] Where: R m_a is the average radius of the hole to which the turbine is attached, 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 mm;

[0079] - the propulsion system further comprises an additional turbine (7) configured to drive the compressor via an additional shaft, the additional shaft being 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 satisfying the following formula:

[0080] R m_a ≥E*L HP 2 *XN*10 -9 +F2

[0081] Where: R m_a is the average radius of the hole to which the turbine is attached, 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);

[0082] 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 satisfying the following formula:

[0083]

[0084] Where: D9 is the diameter of the fan rotor in millimeters (mm) 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 speed, standard atmospheric pressure, and sea level; T e is the inlet temperature of the drive turbine when the propulsion system is at rest, take-off speed, 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);

[0085] 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 satisfying the following formula:

[0086]

[0087] 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 speed, standard atmospheric pressure, and sea level; T e is the inlet temperature of the drive turbine when the propulsion system is at rest, take-off speed, 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);

[0088] - 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 additional turbine being at least equal to:

[0089] R m_a ≥G*FN*BRP*10 -4 +H

[0090] 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 speed, 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 speed, at standard atmospheric pressure, and at sea level; and G = 0.16 millimeters / Newton (mm / N) and H = 18 millimeters (mm);

[0091] - 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:

[0092] R m_a ≤I*FN*BPR*10 -4 +J

[0093] Where: R m_ais the average radius of the hole 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 speed, 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 speed, at standard atmospheric pressure, and at sea level; and I = 0.16 millimeters / Newton (mm / N) and J = 28 millimeters (mm).

[0094] According to a fourth aspect, a method for manufacturing an aviation propulsion system is provided, comprising the following steps:

[0095] - determining the dimensions of the aerospace propulsion system according to the dimensioning method of the third aspect; and

[0096] -Manufacturing of aviation propulsion systems. BRIEF DESCRIPTION OF THE DRAWINGS

[0097] Other features, objects and advantages will appear from the following description which is intended to be illustrative only and non-limiting and which should be read in conjunction with the accompanying drawings, in which:

[0098] 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;

[0099] 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;

[0100] Figure 3 is a schematic cross-sectional view of an example of a speed reduction mechanism according to a first alternative;

[0101] Figure 4 is a schematic cross-sectional view of an example of an epicycloid speed reduction mechanism according to a second alternative;

[0102] Figure 5 is an example of an aircraft according to the first or second embodiment that may include at least one propulsion system;

[0103] Figure 6 is a flow chart illustrating an example of steps in a sizing or manufacturing method according to one embodiment.

[0104] In all figures, similar elements have the same reference numerals. DETAILED DESCRIPTION

[0105] 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 main body 3 (commonly referred to as a "gas generator"), which includes compressor sections 4 and 5, a combustion chamber 6, and turbine sections 7 and 8. Here, the propulsion system 1 is an aircraft propulsion system 1, which is configured to be fixed to an aircraft 100 via a pylon (or mast).

[0106] The compressor sections 4 and 5 include a series of stages, each of which includes a moving impeller (rotor) 4a and 5a rotating in front of a stationary impeller (stator) 4b and 5b. The turbine sections 7 and 8 also include a series of stages, each of which includes a stationary impeller (stator) 7b and 8b, with the moving impeller (rotor) 7a and 8a rotating behind the stationary impeller (stator).

[0107] 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. On the other hand, 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), respectively, are used with reference to radial directions, 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.

[0108] In operation, the air flow F entering the propulsion system 1 is divided into a primary air flow F1 and a secondary air flow F2 , which flow from upstream to downstream in the propulsion system 1 .

[0109] A secondary air flow F2 (also called "bypass air flow") flows around the body 3. The secondary air flow F2 cools the periphery of the body 3 and is used to generate most of the thrust provided by the propulsion system 1.

[0110] The main air flow F1 flows in the main duct within the main body 3, thereby passing through the compressor sections 4, 5, the combustion chamber 6, and the turbine sections 7, 8 in sequence. In the combustion chamber 6, the main air flow F1 is mixed with fuel to be used as an oxidant. The main air flow F1 passes through the turbine sections 7, 8, which receive energy from the combustion chamber 6. This causes the rotors of the turbine sections 7, 8 to rotate. The rotation of the rotors of the turbine sections 7, 8 in turn rotates the rotors of the compressor sections 4, 5 and the rotor portion 9 of the fan section 2.

[0111] In a two-body propulsion system 1, the compressor sections 4 and 5 may include a low-pressure compressor 4 and a high-pressure compressor 5. The turbine sections 7 and 8 may include a high-pressure turbine 7 and a low-pressure turbine 8. The high-pressure turbine 7 rotates the rotor of the high-pressure compressor 5 via a high-pressure shaft 10. The low-pressure turbine rotor 8 rotates the rotor of the low-pressure compressor 4 and the rotor portion 9 of the fan section 2 via a low-pressure shaft 11. Thus, the main body 3 includes a high-pressure main body, which includes the high-pressure compressor 5, the high-pressure turbine 7, and the high-pressure shaft 10, and a low-pressure main body, which includes the fan section 2, the low-pressure compressor 4, the low-pressure turbine 8, and the low-pressure shaft 11. The rotational speed of the high-pressure main body is greater than that of the low-pressure main body. In a three-body propulsion system 1, the turbine sections 7 and 8 also include an intermediate turbine, located between the high-pressure turbine 7 and the low-pressure turbine 8 and configured to drive the rotor of the low-pressure compressor 4 via the intermediate shaft. The fan rotor 9 and the rotor of the high-pressure compressor 5 are still driven by the low-pressure shaft 11 and the high-pressure shaft 10, respectively.

[0112] 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., be driven in the same direction about the longitudinal axis X. Alternatively, the low-pressure shaft 11 and the high-pressure shaft can rotate in opposite directions, i.e., be driven in opposite directions about the longitudinal axis X. Where applicable, 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.

[0113] The fan section 2 includes at least one fan rotor 9, which is rotatable relative to the stator part of the propulsion system by means of the turbine sections 7, 8. Each fan rotor 9 includes a hub 13 and blades 14 extending radially from the hub 13. The blades 14 of each rotor 9 can be fixed relative to the hub 12 or have a variable pitch. In this case, the root of the blades 14 of each rotor 9 can be pivotally mounted along the pitch axis and connected to a pitch changing mechanism 15 installed in the propulsion system 1, the pitch of which is adjusted by the pitch changing mechanism 15 according to the flight phase. The pitch changing mechanism 15 is Figure 1 Shown in dashed lines to illustrate that this feature is optional.

[0114] The fan section 2 may also include a fan stator 16 or flow straightener, which includes blades 17 mounted on a hub 18 of the fan stator 16. The blades 17 have the function of straightening the secondary air flow F2 flowing out of the fan rotor 9. The blades 17 of the fan stator 16 may be fixed relative to the hub or have variable pitch. In a manner similar to the rotor blades 14, the roots of the stator blades 17 are pivotally mounted along the pitch axis X and connected to a pitch change mechanism 15a, which is generally separate from the pitch change mechanism of the fan rotor 9, and the pitch is adjusted by the pitch change mechanism according to the flight phase.

[0115] Thus, the diameter D9 of the fan rotor can be between 80 inches (203.2 cm) and 185 inches (469.9 cm), inclusive. When the fan rotor 9 is ducted, the diameter D9 is, for example, between 85 inches (215.9 cm) and 120 inches (304.8 cm), inclusive, for example, about 90 inches (228.6 cm), which enables the propulsion system 1 to be integrated in a conventional manner, in particular under the wing of an aircraft. When the fan rotor 9 is non-ducted, the diameter D9 is, for example, greater than or equal to 100 inches (254 cm), for example, between 120 inches (304.8 cm) and 156 inches (396.2 cm). Here, the diameter 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. Note that Figure 1 and Figure 2 This is a partial view, diameter D9 is only partially visible.

[0116] 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.

[0117] To improve the propulsion efficiency of the propulsion system 1 and reduce the specific consumption of the propulsion system 1 and the noise emitted by the fan section 2, the propulsion system 1 has a high bypass ratio. The term "high bypass ratio" should refer to 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, standard atmospheric pressure (as defined by the 3rd edition of the International Civil Aviation Organization (OACI) Manual, Document 7488 / 3) and sea level conditions (referred to as sea level standard (SLS) conditions). It should be noted that in this application, unless otherwise specified, parameters (pressure, flow, thrust, speed, etc.) are systematically determined under these conditions. The term "not installed" is used herein to refer to measurements taken when the propulsion system 1 is on a test bench (and not installed on the aircraft 100), in which case the measurements are easier to perform. On the other hand, distances (lengths, radii, diameters, etc.) are measured at ambient temperature (approximately 20° C.) when the propulsion system is cool, ie when the propulsion system has been stationary for a sufficient period of time that the components of the propulsion system are at ambient temperature.

[0118] The fan rotor 9 is separated from the low-pressure shaft 11 by a reduction gear 19, which is arranged between the upstream end of the low-pressure shaft 11 and the fan rotor 9, so that the respective rotational speeds of the fan rotor 9 and the low-pressure shaft 11 can be independently optimized. In this case, the propulsion system 1 also includes an additional shaft (referred to as a fan shaft 20). The low-pressure shaft 11 connects the low-pressure turbine 8 to the input 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.

[0119] 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 low-pressure turbine 8. The overall efficiency of the propulsion system is primarily determined by the propulsive efficiency, which is favorably influenced by minimizing the change in kinetic energy of the air as it passes through the propulsion system 1. In a propulsion system 1 with a high bypass ratio, the majority of the flow that generates propulsion is composed of 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, the propulsive efficiency and the pressure ratio of the fan section 2 are correlated: 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. In this case, the mean pressure is measured at the height of the blade 14 (from the surface radially delimiting the inner flow duct of the fan rotor 9 to the tip 21 of the fan blade 14 ).

[0120] The propulsion system 1 is configured to deliver a thrust between 18,000 lbf (80,068 N) and 51,000 lbf (222,411 N), for example, between 20,000 lbf (88,964 N) and 35,000 lbf (155,688 N).

[0121] 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.

[0122] The ducted fan section 2 includes a fan rotor 9 extending upstream of the fan stator. The stator blades of the fan are generally referred to as outlet guide vanes (OGV) and have a fixed pitch 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 variable pitch. The pressure ratio of the fan can be between 1.20 and 1.45.

[0123] In the unducted fan section 2, the fan section 2 is not surrounded by a fan casing. Since the fan section 2 is unducted, the blades 14 of the fan rotor 9 have variable pitch. A propulsion system comprising at least one unducted fan rotor 9 is referred to as an "open rotor" or "unducted fan". The propulsion system 1 may include two unducted and counter-rotating fan rotors 9. Such a propulsion system 1 is referred to as a "Contra-Rotating Open Rotor (CROR)" or an "Unducted Double Fan (UDF)". One or more fan rotors 9 may be arranged at the rear of the main body 3 to form a push-type structure, or at the front of the main body 3 to form a traction-type structure. Alternatively, the propulsion system 1 may include a single unducted fan rotor 9 and an unducted fan stator 16 (straightener). Such a propulsion system 1 is referred to as an "Unducted Single Fan (USF)". In the case of a propulsion system 1 of the USF type, the blades 17 of the flow straightener 16 are rotationally fixed relative to the axis of rotation X of the upstream fan rotor 9 and are therefore not subject to centrifugal forces. The blades 17 of the flow straightener 16 also have a variable pitch.

[0124] Eliminating the fairing surrounding the fan section 2 allows for a very significant increase in the bypass ratio, without the propulsion system 1 being adversely affected by the mass of the casing or nacelle designed to surround the fan section 2. Consequently, the bypass ratio of the propulsion system 1 including the unducted fan section 2 is greater than or equal to 40, for example, between 40 and 80 inclusive. The peripheral velocity at the tip 21 of the blades 14 of the one or more fan rotors 9 may also be between 210 m / s and 260 m / s. For example, the pressure ratio of the fan may be between 1.05 and 1.20.

[0125] For example, the reduction mechanism 19 may comprise a reduction mechanism having an epicycloid gear train, for example a single-stage or two-stage reduction mechanism of the epicycloid or planetary type, according to the terminology sometimes encountered by those skilled in the art. According to a first alternative, the reduction mechanism 19 may be of the "star type" ( Figure 3 ), and comprises a sun gear 19a (inlet of the reduction mechanism 19), a ring gear 19b (outlet of the reduction mechanism 19) and a series of planetary parts 19c; the sun gear 19a is centered on the rotation axis X of the reduction mechanism 19 (approximately coinciding with the longitudinal axis X) and is configured to be rotated 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 to rotate around the rotation axis X; a series of planetary parts 19c are circumferentially distributed between the sun gear 19a and the ring gear 19b around the rotation axis X, and each planetary part 19c meshes with the sun gear 19a internally and meshes with the ring gear 19b externally. This series of planetary parts 19c is mounted on a planet carrier 19d, which is fixed relative to the stator part 19e of the propulsion system 1, for example relative to the housing of the compressor sections 4, 5. According to a second variant, the reduction mechanism 19 can be a planetary type ( 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 (the planet carrier 19d can therefore rotate relative to the stator part 19e of the propulsion system 1, for example, relative to the housing of the compressor sections 4, 5).

[0126] Regardless of the configuration of the speed reduction mechanism 19 , the diameters of the ring gear 19 b and the planetary carrier 19 d are larger than the diameter of the sun gear 19 a , so that the rotation speed of the fan rotor 9 is lower than the rotation speed of the low-pressure shaft 11 .

[0127] The reduction ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11. In the case of a propulsion system 1 including a ducted fan rotor 9, the reduction ratio may be greater than or equal to 2.7 and less than or equal to 6.0, and is typically approximately 3.0. In the case of a propulsion system 1 including an unducted fan rotor, the reduction ratio may be between 9.0 and 11.0.

[0128] The low-pressure shaft 11 is supported by three or four bearings 11a, 11b and / or 11c to better control the deformation mode of the low-pressure shaft 1 and optionally move the deformation mode of the low-pressure shaft 11 into a transient state of the propulsion system 1, which has a safety margin relative to the steady-state state.

[0129] Applicants have recognized that the position of the bearings 11a-11c of the low-pressure shaft 11 may also have an impact on the deformation pattern of the low-pressure shaft 11. Therefore, the low-pressure shaft 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 first forward bearing 11a may be mounted on the low-pressure shaft 11 and, on the inlet casing 26 of the propulsion system 1, extending between the fan rotor 9 and the low-pressure compressor 4. Where appropriate, the second forward bearing may be mounted on the low-pressure shaft 11 and, on the inter-compressor casing 23 (or intermediate casing) of the propulsion system 1, between the low-pressure compressor 4 and the high-pressure compressor 5. The first aft bearing 11b may be mounted on the low-pressure shaft 11 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). Alternatively, the first aft bearing 11b may be mounted on the exhaust casing 27, extending downstream of the low-pressure turbine 8. The first rear bearing 11b extends downstream of the most downstream bearing 12b of the high-pressure shaft 10. The second rear bearing 11c may be mounted on the exhaust housing 27. Where appropriate, the first and second rear bearings 11b and 11c may be mounted on the same cylindrical sleeve, which is itself fixed to the exhaust housing 27.

[0130] The inlet casing 26 , the intermediate casing 23 , the exhaust casing 27 and, where applicable, the inter-turbine casing 24 form the structural casing of the propulsion system 1 , through which the forces generated by the propulsion system pass.

[0131] The separation of the low-pressure shaft 11 and the fan rotor 9 by the reduction mechanism 19 results in an efficient propulsion system 1 with a pressure ratio of the fan of less than 1.45. As a result, the amount of energy required to drive the fan rotor 9 is reduced, so that the inlet flow to the high-pressure compressor 5 and therefore the inlet cross-section of the high-pressure compressor 5 can be reduced. However, this has the result of limiting the space available for the low-pressure shaft 11 (because the low-pressure shaft 11 is accommodated within the high-pressure shaft 10). In addition, reducing the flow in the high-pressure compressor 5 also has the effect of reducing the flow in the low-pressure turbine 8, which can increase the rotational speed of the low-pressure shaft 11 and can reduce the torque transmitted by the low-pressure shaft 11 and the diameter of the low-pressure shaft 11. However, reducing the diameter of the low-pressure shaft 11 makes it necessary to optimize the dynamic behavior of the low-pressure shaft 11 to prevent the low-pressure shaft 11 from becoming more supercritical.

[0132] In order to optimize the propulsion system 1 while controlling the dynamic behavior of the low-pressure shaft 11 , the dimensions of the propulsion system 11 are determined so that the limiting speed of the drive shaft 11 satisfies the following formula:

[0133]

[0134] Wherein: XN is the limit speed of the drive shaft (11), the unit is revolutions per minute (tr / min);

[0135] R 11b is the average radius of the rear bearing 11 b closest to the front bearing 11 a , in millimeters (mm); and

[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); and

[0137] A=5500mm*tr / min.

[0138] The limit speed XN corresponds to the absolute maximum speed that the low-pressure shaft 11 can encounter in flight (according to European certification regulation EASA CS-E 740 (or according to US certification regulation 14-CFR Part 33)). The limit speed corresponds to the maximum rotational speed when the propulsion system is healthy (and potentially 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 engine certification (Type Certificate Data Sheet). In practice, 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).

[0139] 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 of the rear bearing 11b closest to the front bearing 11a, that is, the 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. It should be noted here that the bearing comprises an outer ring and an inner ring, which are coaxial, and rolling elements (balls, rollers, etc.), which are mounted between the inner ring and the outer ring and are configured so that the inner ring can move relative to the outer ring. The inner ring is mounted on and rotated 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 therefore corresponds to the center of gravity of the assembly formed by the inner ring, the outer ring and the rolling elements. When the inner ring and / or the outer ring are integrally formed with a support (flange, shroud, enclosure, etc.) of the propulsion system 1 and are formed as a single piece so that the ring cannot be distinguished from the support, the center of gravity is determined without taking into account the ring in question.

[0140] 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.

[0141] When the limit speed XN satisfies formula (1), the limit speed is sufficiently high to ensure that the first deformation mode of the low-pressure shaft 11 is within the rotational speed range of the low-pressure shaft 11, within which the dynamic conditions of the low-pressure shaft 11 can be controlled. Typically, for a propulsion system including a reduction gear 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 unlikely to damage the propulsion system 1. Therefore, a transient speed is selected that occurs only for a very short period of time during operation of the propulsion system 1. For example, the propulsion system is sized so that the critical 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 a transient operating range that is rarely used during operation of the propulsion system 1. Thus, the low-pressure shaft 11 remains supercritical, but the dynamic conditions of the low-pressure shaft 11 are controlled.

[0142] Preferably, the limit speed XN also satisfies the following formula:

[0143]

[0144] Where: B = 8500mm*tr / min.

[0145] When the low-pressure shaft 11 does not comply with the formula (2), the critical speed of the low-pressure shaft 11 becomes too high and involves mechanical and dimensional limitations that are difficult to control.

[0146] The limit speed XN of the low-pressure shaft 11 may typically be between 8,500 rpm and 12,000 rpm, for example, between 9,000 rpm and 11,000 rpm.

[0147] The limit speed XN of the low-pressure shaft 11 can be varied by acting on the following elements of the propulsion system 1 :

[0148] 1. Shaft geometry (outer and inner diameters). The outer diameter of the low-pressure shaft influences the design of the high-pressure casing, 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 that increase the critical speed of the low-pressure shaft 11 are limited by the manufacturing constraints of the low-pressure shaft 11.

[0149] 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 much higher Young's modulus / density ratio than conventional steel, but has a lower ability to transmit torque at the interface.

[0150] 3. The number and location of the bearings 11a-11d of the low-pressure shaft 11. The configuration of the low-pressure shaft 11 supported by four bearings increases the critical speed XN of the low-pressure shaft 11. However, compared to a configuration with three bearings, the additional bearings increase the integration difficulty. If appropriate, the rear bearing of the low-pressure shaft 11 can be integrated upstream of the low-pressure turbine 8. This increases the critical speed XN of the low-pressure shaft 11 at the expense of engine length.

[0151] 4. The stiffness of the journals between the rotors of the low-pressure body. Increasing the stiffness of the journals increases the critical speed XN of the low-pressure shaft 11. However, increasing the stiffness of the journals also increases the difficulty of design and integration.

[0152] 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 strongly limited by the integration of the high-pressure body, which leads to supercritical dynamic conditions that must be controlled. Therefore, determining the value of the limit speed XN in a manner consistent with formulas (1) and (2) makes it possible to optimize the integration of the high-pressure body while remaining within the accessible range of the other parameters mentioned above. In addition, in order to avoid the length loss associated with the integration of the bearing for the low-pressure shaft upstream of the low-pressure turbine, adjusting the distance d2 between the centers of gravity of the two rear bearings 11b, 11c of the low-pressure shaft 11 according to formulas (3) and (4) (see below) makes it possible to obtain the desired dynamic conditions.

[0153] Where appropriate, to further improve the control of supercriticality 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 generally referred to as "squeeze film" damping. The bearings may include a flexible cage mounted between an outer ring and a rigid support member fixed to a stator portion of the propulsion system 1 (inlet casing 26, inter-compressor casing 23, inter-turbine casing 24, exhaust casing 27, etc.) to control the stiffness of the bearing and fine-tune the position of the first deformation mode. To this end, the cage includes: a generally cylindrical wall mounted between the rigid support member and the outer ring of the bearing, and radially deformable studs, so that the generally cylindrical wall, and therefore the bearings 11a-11c, can be radially displaced relative to the rigid support member. Examples of bearings with flexible cages and pressurized fluid film dampers including the flexible cages are described in WO 2021 / 001610 and WO 2022 / 195198 in the name of the applicant. Thus, damping the response of the low-pressure shaft 11 using these bearing modal dampers 25 can limit resonance of the low-pressure shaft when its frequency reaches the first deformation mode.

[0154] In one embodiment, to further improve the dynamic control of the low-pressure shaft 11, the radial position and embedment of the rear bearings 11b and 11c of the shaft 11 are also considered when determining the dimensions of the low-pressure shaft 11. The farther the rear bearings 11b and 11c are from the axis X and / or the farther the rear bearings 11b and 11c are from the axis X, the more embedded the rear bearings 11b and 11c are, and thus the stiffness of the low-pressure shaft 11 increases.

[0155] Therefore, in order to increase the embedding of the low-pressure shaft 11 and further improve the dynamics of the low-pressure shaft 11, the limit speed also satisfies the following formula:

[0156]

[0157] Where: R m is the average radius of the low-pressure turbine 8, in millimeters (mm);

[0158] d2 is the distance between the centers of gravity of the rear bearings 11b, 11c of the low-pressure shaft 11, in millimeters (mm); and

[0159] C=0.17(mm*tr / min) -1 , E1=378.

[0160] For example, the limit speed also satisfies the following formula:

[0161]

[0162] Where: E2 = 180. When the low-pressure shaft 11 does not comply with formula (4), the critical speed of the low-pressure shaft 11 becomes too high and involves mechanical and dimensional limitations that are difficult to control.

[0163] The average radius R of the low-pressure turbine 8 m Corresponds to the arithmetic mean 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 corresponds to the arithmetic mean 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, halfway between the leading edge 8c at the tip of the moving blade and the trailing edge 8d at the tip of the moving blade (i.e., at 50% of the chord of the blade tip), between the axis of rotation X and the outer radial surface of the rotor 8a (which radially delimits the rotor blades 8a on the outside).

[0164] In the case where the low-pressure shaft 11 is supported by 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 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.

[0165] 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_a The larger the diameter of the low-pressure shaft 11, the more natural modes of the low-pressure shaft 11 appear at higher rotational speeds (or in other words, the more the deformation mode of the shaft shifts to higher frequencies). Optionally, in order to obtain a compromise between the average hole 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) while controlling 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 hole of the high-pressure turbine 7 can be 1 / 4 of the diameter of the low-pressure shaft 11. m_a The following formula can be met:

[0166] R m_a ≤E*L HP 2 *XN*10 -9 +F1 (5)

[0167] Where: R m_a Expressed in millimeters (mm);

[0168] L HP is the length of the high-pressure body in millimeters (mm);

[0169] E=3.15(mm.tr / min) -1 And F1 = 23 millimeters (mm).

[0170] In fact, 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. The position of the bearings 11a-11b of the low-pressure shaft 11 depends on the length of the high-pressure body. By setting the average radius R of the hole of the high-pressure turbine 7 to 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.

[0171] For example, the average radius R of the hole of each disk of the high-pressure turbine 7 is m_a Satisfies formula (5).

[0172] The average radius R of the hole of the high-pressure turbine 7 m_a The propulsion system 1 that satisfies formula (5) can achieve a total pressure ratio, which corresponds to the pressure ratio between the outlet pressure of the high-pressure compressor 5 and the inlet pressure 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, for example, greater than or equal to 44 and less than or equal to 55.

[0173] 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 within this range and the average radius R of the high pressure turbine is m_a When formula (4) is satisfied, the mechanical load of the rotor 7 a is acceptable for the high-pressure turbine 7 .

[0174] 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.

[0175] 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 The following formula can also be met:

[0176] R m_a ≥E*L HP 2*XN*10 -9 +F2 (6)

[0177] Where: F2 = 13 millimeters (mm).

[0178] The size of the average radius of the holes in accordance 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.

[0179] In some cases, it is possible to determine the high pressure body L HP Before determining the length of the hole of the high pressure turbine 7, determine the average radius R m_a In this case, the average radius of the hole can be determined based on the basic dimensional parameters, which are 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, rather than the length of the high-pressure body (as defined in formulas (5) and (6)):

[0180]

[0181] Where: BPR is the bypass ratio of propulsion system 1, expressed in pounds of thrust;

[0182] D9 is the diameter of the fan rotor 9 in millimeters (mm) and is measured in a plane perpendicular to the axis of rotation (X) at the intersection between the tips 21 and the leading edges 22 of the blades 14 of the fan rotor (9);

[0183] T e is the temperature at the inlet of the low-pressure turbine 8 in degrees Celsius;

[0184] T ref =273K;

[0185] n is the number of stages of the high-pressure body (i.e., the high-pressure turbine 7 and the high-pressure compressor 5);

[0186] Gamma is the adiabatic coefficient of air;

[0187] K=6.76,L=153.6m -1 .(℃) -1 / 2 .(tr / min) -1 , M=421mm.(℃) -1 / 2 And N1 = -11 mm

[0188] (mm).

[0189] As specified in formula (6), when the average radius R of the hole of the high-pressure turbine 7 is m_a When formula (7) is satisfied, the dynamic behavior of the low-pressure shaft 11 can therefore be better controlled.

[0190] For example, the average radius R of the hole of each disk of the high-pressure turbine 7 is m_a Satisfies formula (7).

[0191] In addition, the average radius of the hole can also meet the following formula:

[0192]

[0193] Wherein, N2 = -21 millimeters (mm).

[0194] For example, the average radius R of the hole of each disk of the high-pressure turbine 7 is m_a Satisfies formula (8).

[0195] For example, the propulsion system 1 is also dimensioned so that, 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 The following formula is also satisfied:

[0196] R m_a ≤I*FN*BPR*10 -4 +J (9)

[0197] Where: FN is the thrust of the fan rotor 9, in Newtons (N); and

[0198] I = 0.16 millimeters / Newton (mm / N) and J = 28 millimeters (mm).

[0199] 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 the bypass ratio BPR of the propulsion system 1. In addition, increasing the bypass 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 of the bore of the high-pressure turbine 7. m_a Therefore, by dimensioning the propulsion system, the mean radius R of the bore of the high-pressure turbine 7 is m_a Formula (9) is also satisfied, obtaining a propulsion system 1 whose dimensions of the high-pressure body and the diameter of the low-pressure shaft 11 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.

[0200] For example, the average radius R of the hole of each disk of the high-pressure turbine 7 is m_a Satisfies formula (9).

[0201] For example, the average radius of the bore of a high-pressure turbine is at most 300 mm.

[0202] In one embodiment, in order to maintain the possibility of the low-pressure shaft 11 being introduced 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 ism_a The following formula is also satisfied:

[0203] R m_a ≥G*FN*BPR*10 -4 +H (10)

[0204] Where: G = 0.16 mm / Newton (mm / N) and H = 18 mm.

[0205] For example, the average radius R of the hole of each disk of the high-pressure turbine 7 is m_a Satisfies formula (10).

[0206] Comparison example:

[0207] The engine 1 is a twin-body propulsion system comprising a ducted fan section 2 corresponding to the current technical standards (at the filing date of this application) to which we are seeking to improve.

[0208] The engine 2 is a two-stage propulsion system 1 including a ducted fan section 2 according to the teachings of the present application, the drive shaft of the two-stage propulsion system 1 (ie, here the low-pressure shaft) having a limit speed that complies with formulas (1) and (2) as defined above.

[0209] 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 is damped by a pressurized fluid film.

[0210]

[0211]

[0212]

[0213] The limit speed XN of the low-pressure shaft of the engine 1 is less than the minimum limit speed defined by formula (1), which is the radius R of the upstream rear bearing of the low-pressure shaft. 11b and the bearing-to-bearing distance d1. Consequently, the supercritical first deformation mode of the low-pressure shaft occurs at a transient speed close to takeoff speed. Consequently, despite the presence of the bearing mode damper, the vibration loads observed during the mode transition are high. This is confirmed by the fact that the limiting speed XN of the low-pressure shaft of engine 1 does not conform to equation (4).

[0214] In contrast, the limiting speed of the low-pressure shaft of engine 2 is greater than the minimum limiting speed defined by equation (1). Therefore, the supercritical first deformation mode of the low-pressure shaft occurs in the transient operating range close to idle speed. Assuming that the vibration loads in engine 2 are moderate compared to the vibration loads experienced by engine 1 in this operating range, the bearing mode damper can also damp the mode transition. The limiting speed XN is also kept high enough to prevent the second deformation mode from being within the operating range of engine 2. In addition, although the low-pressure shaft drives a total of seven stages (three compressor stages 4 and four turbine stages 8), the diameter of the low-pressure shaft can be limited. Finally, the average radius of the low-pressure turbine 8 of engine 2 is smaller than the average radius of the low-pressure turbine of engine 1, making the low-pressure turbine 8 of engine 2 more compact in the radial direction than the low-pressure turbine of engine 1 despite the larger number of stages.

[0215] The limit speed XN of the low pressure shaft of the engine 2 is also kept below the maximum limit speed defined in equation (2) to limit the vibration loads seen during the mode change.

[0216] To transition from Engine 1 (reference) to Engine 2 (according to the present disclosure), the fan diameter D9 and the bypass ratio BPR are increased, thereby improving propulsion efficiency and maintaining comparable fan thrust for a given reduction in the pressure ratio of the fan section 2. Furthermore, the overall pressure ratio is increased, and the inlet temperature of the high-pressure turbine 7 is also increased, which makes it possible to increase the thermal efficiency of the propulsion system 1.

[0217] Furthermore, the outer diameter of the shaft of the low-pressure turbine 11 is reduced in order to facilitate the integration of the high-pressure body while complying with the mechanical dimensions of the disks of the high-pressure turbine.

[0218] Adjustments are made to obtain the critical speed of the first deformation mode of the low-pressure shaft 11 within the speed range defined by equations (1) to (4) 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 constraints of the shaft 11.

[0219] Taking into account the location of the piping of the low-pressure compressor, the number of bearings supporting the low-pressure shaft remains at 3 (and is not increased to 4) due to the integration of the bearing upstream of the low-pressure shaft.

[0220] Solutions are also excluded that would advance one or both of the downstream bearings of the low-pressure turbine upstream of the low-pressure turbine, in order to avoid elongation of the engine 2 and to avoid misalignment problems between the two bearings.

Claims

1. An aviation propulsion system (1), comprising: - a fan rotor (9), said fan rotor being connected to 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 limit speed of the driving shaft (11) satisfies the following formula: Wherein: XN is the limit speed of the drive shaft (11), the unit is revolutions per minute (tr / min); R 11b is the average radius of the rear bearing (11b) closest to the front bearing (11a), in millimeters (mm); and d1 is the center of gravity of the front bearing (11a) (G 11a ) and the center of gravity (G 11b ) in millimetres (mm); and A=5500mm*tr / min.

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

3. The propulsion system (1) according to claim 1 or 2, wherein: Each bearing is associated with a bearing modal damper (25).

4. The propulsion system according to claim 3, wherein: The bearing mode damper (25) comprises a squeeze film.

5. The propulsion system according to claim 3 or 4, wherein: The bearing mode damper comprises a deformable cage mounted between the ring of each bearing and the stator part (23, 24, 26, 27) of the propulsion system (1).

6. The propulsion system according to any one of claims 1 to 5, wherein: The bearings include an additional front bearing extending upstream of the front bearing (11a).

7. The propulsion system according to any one of claims 1 to 6, wherein: The limit speed of the drive shaft (11) also satisfies the following formula: Wherein: Rm is the average radius of the driving turbine (8), in millimeters (mm); d2 is the distance between the centers of gravity of the rear bearings, in millimeters (mm); and C=170(mm*tr / min) -1 And E1=378.

8. The propulsion system according to any one of claims 1 to 7, wherein: The limit speed of the drive shaft (11) also satisfies the following formula: Wherein: Rm is the average radius of the driving turbine (8), in millimeters (mm); d2 is the distance between the centers of gravity of the rear bearings, in millimeters (mm); and C=170(mm*tr / min) -1 And E2=180.

9. The propulsion system according to any one of claims 1 to 8, wherein: The bearings include at least one front bearing (11a) and exactly two rear bearings.

10. The propulsion system according to any one of claims 1 to 9, 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.

11. The propulsion system (1) according to claim 10, wherein: The additional compressor (5) comprises at least eight stages and at most eleven stages.

12. The propulsion system (1) according to any one of claims 1 to 11, wherein: The mean radius of the bore of the additional turbine (7) is at most equal to 300 mm.

13. The propulsion system (1) according to any one of claims 1 to 12, wherein: The driving turbine (8) comprises at least three stages and at most five stages.

14. The propulsion system (1) according to any one of claims 1 to 13, wherein: The reduction ratio of the reduction mechanism is greater than or equal to 2.5 and less than or equal to 11.

15. An aircraft comprising at least one propulsion system (1) according to any one of claims 1 to 14, said propulsion system being fixed to the aircraft via a mast.

16. A method for determining the dimensions of a propulsion system (1), the propulsion system comprising a reduction system (19) coupling a drive shaft (11) and a fan rotor (9) so as to drive the fan rotor (9) at a speed lower than the speed of the drive shaft (11); the bearings being configured to center the drive shaft (11) relative to an axis of rotation (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 propulsion system is dimensioned so that the limiting speed of the drive shaft (11) satisfies the following formula: in: XN is the limit speed of the drive shaft (11), in revolutions per minute (tr / min); R 11b is the average radius of the rear bearing (11b) closest to the front bearing (11a), in millimeters (mm); and d1 is the center of gravity of the front bearing (11a) (G 11a ) and the center of gravity (G) of the rear bearing (11b) closest to the front bearing (11a) 11b ) in millimetres (mm); and A=5500mm*tr / min.

17. The size determination method according to claim 16, wherein: The propulsion system is dimensioned so that the limiting speed of the drive shaft (11) also satisfies the following formula: Where: B = 8500mm*tr / min.

18. The size determination method according to claim 16 or 17, wherein: The propulsion system is dimensioned so that the limiting speed of the drive shaft (11) also satisfies the following formula: Wherein: Rm is the average radius of the driving turbine (8), in millimeters (mm); d2 is the distance between the centers of gravity of the rear bearings, in millimeters (mm); and C=170(mm*tr / min) -1 And E1=378.

19. The size determination method according to any one of claims 16 to 18, wherein: The propulsion system is dimensioned so that the limiting speed of the drive shaft (11) also satisfies the following formula: Wherein: Rm is the average radius of the driving turbine (8), in millimeters (mm); d2 is the distance between the centers of gravity of the rear bearings, in millimeters (mm); and C=170(mm*tr / min) -1 And E2=180.

20. A method for manufacturing an aerospace propulsion system, the method comprising the steps of: - determining the dimensions of the aeronautical propulsion system according to the method of any one of claims 16 to 19; as well as - Manufacturing of said aviation propulsion system.

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