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

Abstract

An aviation propulsion system (1) having an average radius of the bores of a high pressure turbine (7) equal at most to Rm-a < = A * L2 * XN * 10 <-9 >-B, where: Rm-a is the average radius of the bores of the high pressure turbine (7) in millimeters (mm); and L is the distance between the inlet of the high-pressure compressor (5) and the outlet of the high-pressure turbine (7) in millimeter (mm); a = 8.3 (mm.Tr / min)-1 and B = 24.5 mm (mm); and XN is the limit speed of the low-pressure shaft (11) in the unit of rotational speed per minute (tr / min).

Description

Technical Field

[0001] The present application relates generally to the field of propulsion systems, and more particularly to aviation propulsion systems having high or very high bypass ratios. 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 driven to rotate by the high-pressure turbine via a high-pressure shaft. The fan and the low-pressure compressor (where applicable) are driven to rotate by the low-pressure turbine via a low-pressure shaft.

[0003] The technical research work has significantly improved the environmental performance of aircraft. The applicant has considered factors affecting all design and development stages to obtain less energy-intensive and more environmentally friendly aviation components and products. The integration and use of these aviation components and products in civil aviation have a moderate environmental impact, thereby achieving the goal of improving the energy efficiency of aircraft.

[0004] Therefore, in order to improve the propulsion efficiency of the propulsion system and reduce the unit 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 between the flow rate of the secondary air flow and 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] The current trend is to increase the overall compression 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 indeed allows the inlet pressure of the combustion chamber gases to be increased, thus further improving the overall efficiency of the propulsion system. Therefore, increasing the overall compression ratio requires increasing the compression ratio of the high-pressure compressor and / or the low-pressure compressor, especially since, for the reasons mentioned above, the fan compression ratio is simultaneously sought to be reduced.

[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 helps 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 eigenmode, while increasing the rotational speed of the low-pressure shaft expands the operating range of the low-pressure shaft. As a result, the low-pressure shaft can exceed the critical speed and can enter resonance. Under resonance, which occurs when the low-pressure shaft exceeds the critical speed, the low-pressure shaft is subjected to an overvoltage phenomenon, which amplifies 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 its dynamic behavior and changes the location of its supercritical mode. However, a propulsion system rotating at a critical speed in steady state runs the risk of 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, according to a first aspect, an aviation propulsion system is proposed, comprising:

[0010] a first turbine configured to drive the first compressor about a rotation axis via a first shaft;

[0011] a second turbine configured to drive the second compressor via a second shaft configured to rotate about the axis of rotation at a higher speed than the first shaft;

[0012] - a fan rotor, the fan rotor being connected to the fan shaft;

[0013] a speed reduction mechanism coupling the first shaft and the fan shaft so as to drive the fan shaft at a rotational speed lower than that of the first shaft;

[0014] The average radius of the hole of the second turbine satisfies the following formula:

[0015] R m_a_Min ≤R m_a ≤R m_a_MAX

[0016] in:

[0017]

[0018] R m_a is the average radius of the hole of the second turbine, in millimeters (mm);

[0019] 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 blades of the fan rotor;

[0020] 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;

[0021] T e is the maximum inlet temperature of the drive turbine when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C);

[0022] T ref =273K;

[0023] n is the number of stages of the second turbine and the second compressor;

[0024] Gamma (GAMMA) is the adiabatic coefficient of air;

[0025] XN is the maximum speed of the first axis, in revolutions per minute (tr / min);

[0026] K=6.76,L=153.6m -1 .(℃) -1 / 2 .(tr / min) -1 , M=421mm.(℃) 1 / 2 , N1 = -11 mm.

[0027] Some preferred but non-limiting features of the propulsion system according to the first aspect are the following features taken alone or in combination:

[0028] -The average radius of the hole of the second turbine also satisfies the following formula:

[0029] R m_a ≤E*L HP 2 *XN*10 -9 +F1

[0030] Where: L HP is the distance between the inlet of the second compressor and the outlet of the second turbine, in millimeters (mm); and E = 3.15 (mm.tr / min) -1 and F1=23 mm;

[0031] -The average radius of the hole of the second turbine also satisfies the following formula:

[0032] R m_a ≥E*L HP 2 *XN*10-9 +F2

[0033] Where: L HP is the distance between the inlet of the second compressor (5) and the outlet of the second turbine (7), in millimeters (mm); and E = 3.15 (mm.tr / min) -1 And F2 = 13 millimeters (mm).

[0034] According to a second aspect, an aviation propulsion system is provided, comprising:

[0035] a first turbine configured to drive the first compressor about a rotation axis via a first shaft;

[0036] a second turbine configured to drive the second compressor about a rotation axis via a second shaft configured to rotate at a higher speed than the first shaft;

[0037] - a fan rotor, the fan rotor being connected to the fan shaft;

[0038] a speed reduction mechanism coupling the first shaft and the fan shaft so as to drive the fan shaft at a rotational speed lower than that of the first shaft;

[0039] The average radius of the hole of the second turbine satisfies the following formula:

[0040] R m_a ≤E*L HP 2 *XN*10 -9 +F1

[0041] Where: R m_a is the average radius of the hole of the second turbine, in millimeters (mm);

[0042] XN is the maximum speed of the first axis, in revolutions per minute (tr / min);

[0043] L HP is the distance between the inlet of the second compressor and the outlet of the second turbine, in millimeters (mm); and

[0044] E=3.15(mm.tr / min) -1 And F1 = 23 millimeters (mm).

[0045] Some preferred but non-limiting features of the propulsion system according to the second aspect are the following features taken alone or in combination:

[0046] -The average radius of the hole of the second turbine also satisfies the following formula:

[0047] R m_a ≥E*LHP 2 *XN*10 -9 +F2

[0048] Where: L HP is the distance between the inlet of the second compressor (5) and the outlet of the second turbine (7), in millimeters (mm); and E = 3.15 (mm.tr / min) -1 and F2 = 13 millimeters (mm); and / or

[0049] -The average radius of the hole of the second turbine also satisfies the following formula:

[0050] R m_a_Min ≤R m_a ≤R m_a_MAX

[0051] in:

[0052]

[0053]

[0054] 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 blades of the fan rotor;

[0055] 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;

[0056] T e is the maximum inlet temperature of the drive turbine when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C);

[0057] T ref =273K;

[0058] n is the number of stages of the second turbine and the second compressor;

[0059] Gamma is the adiabatic coefficient of air; and

[0060] K=6.76,L=153.6m -1 .(℃) -1 / 2 .(tr / min) -1 , M=421mm.(℃) 1 / 2 , N1 = -11 millimeters (mm) and N2 = -21 millimeters (mm).

[0061] Some preferred but non-limiting features of the propulsion system according to the first and / or second aspects are the following features, taken alone or in combination:

[0062] -The average radius of the hole of the second turbine (7) also satisfies the following formula:

[0063] R m_a ≥G*FN*BPR*10 -4 +H

[0064] where: FN is the thrust of the propulsion system and is measured when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level conditions and is expressed in Newtons (N);

[0065] 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

[0066] G = 0.16 mm / Newton (mm / N) and H = 18 mm;

[0067] -The average radius of the hole of the second turbine (7) also satisfies the following formula:

[0068] R m_a ≤I*FN*BPR*10 -4 +J

[0069] where: FN is the thrust of the propulsion system and is measured when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level conditions and is expressed in Newtons (N);

[0070] 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

[0071] I = 0.16 mm / Newton (mm / N) and J = 28 mm;

[0072] - the hub-to-tip ratio of the second turbine is between 0.77 and 0.90;

[0073] - the propulsion system further comprises a bearing configured to center the first shaft relative to the axis of rotation, the bearing comprising a bearing mode damper;

[0074] - the propulsion system further includes an inter-turbine housing extending between the first turbine and the second turbine and configured to support the second turbine via a bearing assembly;

[0075] the inter-turbine housing comprises a plurality of guide vanes configured to straighten the air flow at the inlet of the first turbine;

[0076] - the propulsion system includes between twenty and thirty guide vanes;

[0077] - the total compression ratio of the propulsion system corresponds to the ratio between the outlet pressure of the second compressor and the inlet pressure of the fan rotor, the total compression ratio being greater than or equal to 40 and less than or equal to 70;

[0078] -The second turbine is a two-stage turbine;

[0079] - the second compressor comprises at least eight stages and at most eleven stages;

[0080] a bypass ratio of the propulsion system greater than or equal to 10, such as between 10 and 35 inclusive, such as between 10 and 18 inclusive, such as between 10 and 15 inclusive;

[0081] - the first turbine comprises at least three stages and at most five stages; and / or

[0082] - The first compressor comprises at least two stages and at most four stages.

[0083] According to a third aspect, an aircraft is proposed, comprising at least one propulsion system according to the first or second aspect, the propulsion system being fixed to the aircraft via a mast.

[0084] According to a fourth aspect, a method for determining the size of a propulsion system or manufacturing a propulsion system is provided. The propulsion system includes a reduction mechanism and a second turbine. The reduction mechanism couples a first turbine and a fan rotor so as to drive the fan rotor at a speed lower than that of the first turbine. The second turbine is configured to rotate at a speed higher than that of the first turbine. The size of the second turbine is determined so that the average radius of the bore of the second turbine satisfies the following formula:

[0085] R m_a_Min ≤R m_a ≤R m_a_MAX

[0086] in:

[0087]

[0088] R m_a is the mean radius of the bore of the second turbine, in millimeters (mm); 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 blades of the fan rotor; 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 maximum inlet temperature of the drive turbine when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C); T ref= 273K; n is the number of stages of the second turbine and the second compressor; gamma is the adiabatic coefficient of air; XN is the limit speed of the first shaft in revolutions per minute (tr / min); and K = 6.76, L = 153.6m -1 .(℃) -1 / 2 .(tr / min) -1 , M=421mm.(℃) 1 / 2 , N1 = -11 millimeters (mm) and N2 = -21 millimeters (mm).

[0089] Some preferred but non-limiting features of the sizing method or manufacturing method according to the fourth aspect are the following features taken alone or in combination:

[0090] The size of the second turbine is determined so that the average radius of the hole of the second turbine also satisfies the following formula:

[0091] R m_a ≤E*L HP 2 *XN*10 -9 +F1

[0092] Where: L HP is the distance between the inlet of the second compressor and the outlet of the second turbine, in millimeters (mm); and E = 3.15 (mm.tr / min) -1 and F1=23 mm;

[0093] The size of the second turbine is determined so that the average radius of the hole of the second turbine also satisfies the following formula:

[0094] R m_a ≥E*L HP 2 *XN*10 -9 +F2

[0095] Where: L HP is the distance between the inlet of the second compressor and the outlet of the second turbine, in millimeters (mm); and E = 3.15 (mm.tr / min) -1 and F2 = 13 millimeters (mm);

[0096] The size of the second turbine (7) is determined so that the average radius of the hole of the second turbine also satisfies the following formula:

[0097] R m_a ≥G*FN*BPR*10 -4 +H

[0098] Wherein: FN is the thrust of the propulsion system and is measured when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level conditions, 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 conditions; and G = 0.16 millimeters per Newton (mm / N) and H = 18 millimeters (mm);

[0099] The size of the second turbine is determined so that the average radius of the hole of the second turbine also satisfies the following formula:

[0100] R m_a ≤I*FN*BPR*10 -4 +J

[0101] Wherein: FN is the thrust of the propulsion system and is measured when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level conditions 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 per newton (mm / N) and J = 28 millimeters (mm).

[0102] According to a fifth aspect, a method for manufacturing a propulsion system is proposed, comprising the following steps:

[0103] - sizing of an aerospace propulsion system according to the method of the fourth aspect; and

[0104] -Manufacturing of aviation propulsion systems. BRIEF DESCRIPTION OF THE DRAWINGS

[0105] 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:

[0106] Figure 1 is a schematic partial cross-sectional view of an example of a propulsion system according to an embodiment;

[0107] Figure 2 is a schematic cross-sectional view of one example of a speed reduction mechanism according to a first modification;

[0108] Figure 3 is a schematic cross-sectional view of one example of a speed reduction mechanism according to a second modification;

[0109] Figure 4 is an example of an aircraft that may include at least one propulsion system according to an embodiment; and

[0110] Figure 5is a flow chart illustrating an example of steps in a sizing or manufacturing method according to one embodiment.

[0111] In all figures, similar elements have the same reference numerals. DETAILED DESCRIPTION

[0112] 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 includes 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, which is configured to be fixed to an aircraft 100 via a pylon (or mast).

[0113] Compressor sections 4 and 5 include a series of stages, each including impellers (rotors) 4a and 5a rotating in front of impellers (stators) 4b and 5b. Turbine sections 7 and 8 also include a series of stages, each including impellers (stators) 7b and 8b rotating behind them.

[0114] In the present application, the axial direction corresponds to the direction of the longitudinal axis X, which corresponds to the rotation of the gas generator shaft, 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.

[0115] 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.

[0116] 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.

[0117] The main air flow F1 flows in a main path within the main body 3, 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 serve as an oxidant. The main air flow F1 flows 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, which in turn drives the rotors of the compressor sections 4, 5 and the rotor portion 9 of the fan section 2 to rotate.

[0118] 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 the high-pressure turbine 7 drives the rotor of the high-pressure compressor 5 via a high-pressure shaft 10. The rotor of the low-pressure turbine 8 drives 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 high-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, 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.

[0119] 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. Alternatively, the low-pressure shaft 11 and the high-pressure shaft can rotate in opposite directions, i.e., they can 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.

[0120] The fan section 2 comprises at least a fan rotor 9 which can be driven by the turbine sections 7, 8 to rotate relative to a fan casing 12. Each fan rotor 9 comprises 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 setting. In this case, the root of the blades 14 of each rotor 9 can 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 according to the flight phase. The pitch changing mechanism 15 is Figure 1 Shown in dashed lines to illustrate that this feature is optional.

[0121] The fan section 2 may also include a fan stator 16 or straightener comprising blades 17 mounted on a hub 18 of the fan stator 16 and having the function of straightening the secondary air flow F2 flowing at the outlet of the fan rotor 9. The blades 17 of the fan stator 18 may be fixed relative to the hub 18 or may have a variable setting.

[0122] In order 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. "High bypass ratio" herein refers to a bypass ratio greater than or equal to 10, for example, a bypass ratio 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 known as sea level standard (SLS)). It should be noted that in this application, parameters (pressure, flow, thrust, speed, etc.) are systematically determined under these conditions. "Not installed" herein refers to measurements taken when the propulsion system 1 is in a test bench (and not installed on the aircraft 100), when such measurements are more easily performed. However, the distances (length, radius, diameter) are measured at normal temperature (about 20° C.) when the propulsion system 1 is cooled, ie when the propulsion system is stopped for a sufficient time so that the components of the propulsion system are at normal temperature.

[0123] 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 (referred to as a 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 outlet 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.

[0124] This separation makes it possible to reduce the rotational speed and pressure ratio of the fan rotor 9 and increase the power extracted by the low-pressure turbine 8. In fact, the overall efficiency of the propulsion system depends mainly on the propulsive efficiency, which is favorably influenced by minimizing the changes in the 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 the propulsive force is composed of the secondary air flow F2 of the propulsion system 1, the kinetic energy of the secondary air flow F2 being mainly affected by the compression experienced by the secondary air flow F2 during its passage through the fan section 2. Therefore, the 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. In order to optimize the propulsive efficiency of the propulsion system 1, the fan pressure ratio (corresponding to the ratio between the average outlet pressure of the fan stator 17 (or, in the absence of a stator, the fan rotor 9) and the average inlet pressure 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 average pressure is measured along the height of the blade 14 (from the surface delimiting the flow path radially inwardly at the inlet of the fan rotor 9 to the tip 21 of the fan blade 14 ).

[0125] 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).

[0126] The fan section 2 is of ducted type and, for this purpose, comprises a fan housing 12 in which the fan rotor 9 is accommodated.

[0127] The fan section 2 particularly includes a fan rotor 9 extending upstream of the fan stator. The blades 16 of the fan stator 17 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), for example, between 10 and 15 (including 10 and 15). It should be noted that when the bypass ratio is greater than or equal to 25, preferably, the fan rotor 9 is a fan rotor with variable setting. The peripheral speed at the tip 21 of the blades of the fan rotor 9 can also 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.

[0128] The reduction mechanism 19 may comprise, for example, a reduction mechanism 19 having a planetary gear train, such as a single-stage or two-stage reduction mechanism 19 of "planetary" or "star type" according to the terminology encountered by those skilled in the art. According to a first variant, the reduction mechanism 19 may be of star type ( Figure 2 ), and comprises a sun pinion 19a (inlet of the reduction mechanism 19), a ring gear 19b (outlet of the reduction mechanism 19) and a series of planetary gears 19c; the sun pinion 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 driven in rotation by the low-pressure shaft 11; the ring gear 19b is coaxial with the sun pinion 19a and is configured to drive the fan shaft 20 to rotate around the rotation axis X; a series of planetary gears 19c are circumferentially distributed between the sun pinion 19a and the ring gear 19b around the rotation axis X, each planetary gear 19c meshing with the sun pinion 19a internally and with the ring gear 19b externally. The series of planetary gears 19c are mounted on a planetary gear 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" ( Figure 3 ), 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 driven in rotation by the planetary gear carrier 19d (the planetary gear 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).

[0129] Regardless of the configuration of the speed reduction mechanism 19 , the diameters of the ring gear 19 b and the planetary gear carrier 19 d are larger than the diameter of the sun pinion 19 a , thereby making the rotation speed of the fan rotor 9 lower than the rotation speed of the low-pressure shaft 11 .

[0130] The reduction ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11, for example, greater than or equal to 2.7 and less than or equal to 6.0, for example, approximately 3.0.

[0131] Separating the low-pressure shaft 11 and the fan rotor 9 by means of a reduction mechanism makes it possible to obtain an efficient propulsion system 1 with a fan pressure ratio of less than 1.45. As a result, the amount of energy driving the fan rotor 9 is reduced, so that the flow at the inlet of the high-pressure compressor 5 and therefore the inlet 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, 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 requires optimizing 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 average radius R of the bore of the high-pressure turbine 7 is m_a is at most equal to:

[0133] R m_a_Min ≤R m_a ≤R m_a_MAX (1)

[0134] in:

[0135]

[0136] BPR is the bypass ratio of propulsion system 1;

[0137] 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. Figure 1 If it is a partial view, the diameter D9 is only partially visible;

[0138] T e is the maximum inlet temperature of the low-pressure turbine 8, in degrees Celsius (°C);

[0139] T ref =273K;

[0140] n is the number of stages of the high-pressure body (i.e., the high-pressure turbine 7 and the high-pressure compressor 5);

[0141] Gamma is the adiabatic coefficient of air;

[0142] K=6.76,L=153.6m -1 .(℃) -1 / 2 .(tr / min) -1 , M=421mm.(℃) 1 / 2, N1 = -11 millimeters (mm) and N2 = -21 millimeters (mm); and

[0143] XN is the limit speed of the low-pressure shaft 11, expressed in revolutions per minute (tr / min), corresponding to the absolute maximum speed that the low-pressure shaft 11 may encounter during the entire flight (according to European certification regulation EASA CS-E 740 (or according to US certification regulation 14-CFR Part 33.87)). The limit speed corresponds to the maximum rotational speed when the propulsion system is intact (and likely 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 forms 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 dimensions of the propulsion system 1 and in some certification tests (such as blade loss or rotor integrity tests).

[0144] The average radius R of the hole of the high-pressure turbine 7 m_a Corresponds to the distance, measured in a plane perpendicular to the longitudinal axis X, halfway between the leading edge 7c and the trailing edge 7d of the blades of the most upstream rotor 7a of the high-pressure turbine 7, i.e. the rotor 7a of the first stage of the high-pressure turbine 7, measured between a bore of a disk of the rotor 7a of the high-pressure turbine 7 and the axis of rotation X. The bore here corresponds to the radially inner surface of the disk of the rotor 7a.

[0145] Since the low-pressure shaft 11 is accommodated in the high-pressure shaft 10 , the diameter of the high-pressure shaft 10 is affected by the average radius R of the bore of the high-pressure turbine 7 . m_a However, the larger the diameter of the low-pressure shaft 11, the more eigenmodes appear in the low-pressure shaft 11 at high rotational speeds (or in other words, the more the deformation mode of the shaft shifts toward high frequencies). Therefore, formula (1) makes it possible to achieve a compromise between the small diameter average bore radius of the high-pressure turbine 7, which makes it possible to improve the overall compression 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.

[0146] In fact, the deformation frequency of the low-pressure shaft 11 is proportional to the ratio of the diameter of the low-pressure shaft 11 to the square of the distance between the bearings 11a-11c of the low-pressure shaft 11. Therefore, the farther apart the bearings 11a-11c 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-11c of the low-pressure shaft 11 depends on the length L of the high-pressure body. HP , which itself depends on the number of stages of the high-pressure body. By taking the average radius R of the hole of the high-pressure turbine 7 m_a Defined as a function of the number of stages of the high-pressure body (the high-pressure body is formed by the high-pressure compressor 5 and the high-pressure turbine 7 ), it is thus possible to better control the dynamic behavior of the low-pressure shaft 11 .

[0147] 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 are basic dimensional parameters of the propulsion system 1. Therefore, the average radius R of the hole of the high-pressure turbine 7 is m_a The function defined as these parameters and the number of stages of the high-pressure body enables the dimensioning of the high-pressure turbine to be carried out at an upstream stage of the dimensioning process of the propulsion system 1 .

[0148] Note that the limit speed XN is between 8500 rpm and 12000 rpm, for example, between 9000 rpm and 11000 rpm.

[0149] For example, the average radius R of the hole of each disk of the high-pressure turbine 7 is m_a Satisfies formula (1).

[0150] For example, the average radius of the bores of a high-pressure turbine is at most equal to 300 mm.

[0151] The inlet temperature of the low-pressure turbine is T e It can be between 950°C and 1230°C.

[0152] When the length L of the high-voltage body HP When applicable, the average radius of the hole of the high-pressure turbine 7 satisfies the following formula:

[0153] R m_a ≤E*L HP 2 *XN*10 -9 +F1 (2)

[0154] Where: L HP is the length of the high-pressure body in millimeters (mm);

[0155] E=3.15(mm.tr / min) -1 And F1 = 23 millimeters (mm).

[0156] Length L of the high-pressure body HP Corresponds to the distance between the inlet of the high-pressure compressor 5 and the outlet of the high-pressure turbine 7. The inlet of the high-pressure compressor 5 corresponds to the most upstream portion of the most upstream impeller 5b (rotor) of the high-pressure compressor 5 (that is, approximately the leading edge 5c at the blade root of the first stage of the high-pressure compressor 5). The outlet of the high-pressure turbine 7 corresponds to the most downstream portion of the impeller 7a (rotor) of the high-pressure turbine 7 (that is, approximately the trailing edge 7d at the blade root of the last stage of the high-pressure turbine 7).

[0157] For a high pressure main body including 10 to 13 compressor stages and turbine stages, the high pressure main body L HPThe length of can be between 950mm and 1450mm, for example.

[0158] For example, the average radius R of the hole of each disk of the high-pressure turbine 7 is m_a Satisfies formula (2).

[0159] 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:

[0160] R m_a ≥E*L HP 2 *XN*10 -9 +F2 (3)

[0161] Where: F2 = 13 millimeters (mm).

[0162] Sizing the mean radius of the holes to satisfy formula (3) ensures a minimum radius for the high-pressure turbine and therefore ensures the ability of the high-pressure turbine to withstand centrifugal forces.

[0163] For example, the average radius R of the hole of each disk of the high-pressure turbine 7 is m_a Satisfies formula (3).

[0164] The average radius R of the bore of the high-pressure turbine 7 of the propulsion system 1 m_a Satisfying all or part of formulas (1) to (3) can achieve an overall compression 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 overall compression 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.

[0165] 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 therefore the high pressure shaft 10) is within this range and when the average radius R of the high pressure turbine is m_a When formula (1) is satisfied, the mechanical load of the rotor 7 a is acceptable for the high-pressure turbine 7 .

[0166] In one embodiment, the propulsion system 1 is also dimensioned such that the average radius R of the bore of the high-pressure turbine 7 is m_a is at most equal to:

[0167] R m_a ≤I*FN*BPR*10 -4 +J (4)

[0168] Where: FN is the thrust of propulsion system 1, in Newtons (N);

[0169] BPR is the bypass ratio of the propulsion system 1; and

[0170] I = 0.16 millimeters / Newton (mm / N) and J = 28 millimeters (mm).

[0171] In practice, 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. Furthermore, 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 mean 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 (4) 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.

[0172] For example, the average radius R of the hole of each disk of the high-pressure turbine 7 is m_a Formula (4) is also satisfied.

[0173] 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, the average radius R of the hole of the second turbine 7 is m_a At least equal to:

[0174] R m_a ≥G*FN*BPR*10 -4 +H (5)

[0175] Where: G = 0.16 mm / Newton (mm / N) and H = 18 mm.

[0176] For example, the average radius R of the hole of each disk of the high-pressure turbine 7 is m_a Satisfies formula (5).

[0177] The applicant has also noted that the positions of the bearings 11a-11c of the low-pressure shaft 11 may also have an impact on the deformation mode of the low-pressure shaft 11. In this case, 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 11 and possibly change the deformation mode of the low-pressure shaft 11 in a transient state of the propulsion system 1 with a safety margin compared to the stable state.

[0178] Therefore, the low-pressure shaft can include one or two forward bearings 11a and two aft bearings 11b and 11c. The first forward bearing 11a is mounted on the low-pressure shaft 11 and on the inlet casing 26 of the propulsion system 1, which extends between the fan rotor 9 and the low-pressure compressor 4. If desired, the second forward bearing can be mounted on the low-pressure shaft 11 and on the inter-compressor casing 23 (or intermediate casing) of the propulsion system 1, that is, between the low-pressure compressor 4 and the high-pressure compressor 5. The first aft bearing 11b can be mounted on the low-pressure shaft 11 and on the inter-turbine casing 24 upstream of the low-pressure turbine 8 (that is, the casing extending between the high-pressure turbine 7 and the low-pressure turbine 8). Alternatively, the first aft bearing 11b can be mounted on the exhaust casing 27, which extends downstream of the low-pressure turbine 8. The first aft bearing 11b extends downstream of the most downstream bearing 12b of the high-pressure shaft 10. The second aft bearing 11c can be mounted on the exhaust casing 27. If desired, the first rear housing 11 b and the second rear housing 11 c may be mounted on the same cylindrical shroud, which is itself secured to the exhaust housing 27 .

[0179] The inlet casing 26 , the intermediate casing 23 , the exhaust casing 27 and, if necessary, the inter-turbine casing 24 form the structural casing of the propulsion system 1 , through which the forces generated by the propulsion system 1 are transmitted.

[0180] The high-pressure turbine 7 that satisfies formula (1) may have a hub-to-tip ratio, which corresponds to the ratio between the outer radius R2 of the high-pressure turbine 7 and the inner radius R1 of the rotor 7a (blade wheel) of the high-pressure turbine 7, and the hub-to-tip ratio is greater than 0.77 and less than 0.90. Here, the outer radius R2 and the inner radius R1 of the high-pressure turbine 7 are measured in a plane perpendicular to the axis of rotation X, and are measured at the hub of the most downstream rotor 7a of the high-pressure turbine 7 (i.e., the last stage of the high-pressure turbine 7) at the middle between the leading edge 7c and the trailing edge 7d of the blade of the rotor 7a (i.e., at 50% of the chord at the blade root). The outer radius R2 of the high-pressure turbine 7 corresponds to the distance between the tip 7e of the blade of the rotor 7a of the high-pressure turbine 7 and the axis of rotation X of the high-pressure turbine 7 in this plane. The inner radius R1 of the rotor 7a corresponds to the following distance, which is measured in a plane perpendicular to the axis of rotation X of the high-pressure turbine 7, between the outer radial surface of the hub of the rotor 7a (the outer radial surface delimits the flow path in the rotor 7a radially on the inside) and the axis of rotation X, and is measured at the hub halfway between the leading edge 7c and the trailing edge 7d of the blades of the rotor 7a (at 50% of the chord at the blade root).

[0181] This high-pressure turbine 7 has an optimized outlet cross section S sIn fact, the smaller the hub-to-tip ratio, the smaller the outer diameter of the high-pressure turbine 7 (for the same cross-section). Therefore, a hub-to-tip ratio between 0.77 and 0.90 is associated with an optimized mean radius R of the reduced bore. m_a (satisfying formula (1) and the rotational speed as described above) in combination, it is possible to obtain a more efficient high-pressure turbine 7 (due to the adapted rotational speed of the high-pressure turbine 7) in a suitable occupied space, while also optimizing the outlet surface for expanding the gas at the outlet of the combustion chamber 5. Therefore, the dimensions of the high-pressure turbine 7 are determined to obtain a hub-to-tip ratio between 0.77 and 0.90, which makes the high-pressure turbine 7 more efficient and thus reduces the specific consumption of the propulsion system 1 without compromising the dynamic behavior of the low-pressure shaft 11.

[0182] To further improve control of the supercritical state of the low-pressure shaft 11, the bearings 11a-11c of the low-pressure shaft 11 include bearing mode dampers comprising a pressurized oil film configured to damp mode switching (viscous damping). This type of damping is often referred to as a "squeeze film."

[0183] Furthermore, propulsion system 1 includes an inter-turbine casing 24 extending between high-pressure turbine 7 and low-pressure turbine 8. The inter-turbine casing defines the flow path between high-pressure turbine 7 and low-pressure turbine 8 and includes guide vanes 25 configured to straighten the air flow exiting high-pressure turbine 7, thereby improving the supply to low-pressure turbine 8 and, therefore, the efficiency of the low-pressure turbine. Thus, guide vanes 25 have an aerodynamic surface configured to redirect the air flow entering low-pressure turbine 8. In one embodiment, propulsion system 1 includes between twenty and thirty guide vanes 25.

[0184] If necessary, the inter-turbine casing 24 forms the structural housing of the propulsion system 1, thereby making it possible to improve the overall dynamics of the propulsion system 1. To this end, the inter-turbine casing comprises an inner shroud mounted on a set of bearings of the propulsion system 1 (usually the rear bearing 26 of the high-pressure shaft), an outer shroud that can be configured to absorb the mechanical forces in the propulsion system 1, and a series of arms extending radially between the inner and outer shrouds and configured to allow the passage of auxiliary equipment and to absorb the mechanical forces between the inner and outer shrouds. The guide vanes 25 can be distinct from the arms and extend between the arms and the low-pressure turbine 8.

[0185] Comparison example:

[0186] The engine 1 is a twin-body propulsion system including a ducted fan and corresponding to current technical standards (at the date of filing of the present application) for which improvements are sought.

[0187] Engine 2 is a twin-body propulsion system including a ducted fan according to the teachings of the present application and has a mean bore radius that satisfies equation (1).

[0188]

[0189]

[0190] The dynamic condition of the low pressure shaft of the engine 1 is subcritical. The engine 1 also has a minimum mean bore radius R less than as defined in equation (1). m_a_MIN The average hole radius R m_a , while engine 2 has an average bore radius R according to formula (1) m_a . As a result, engine 2 is more compact than engine 1, has a higher bypass ratio, a reduced number of stages, and the dynamic conditions of the low-pressure shaft of engine 2 are supercritical. However, this dynamic condition can be controlled by introducing flexible cages and viscous vibration dampers on all or part of the bearings of the low-pressure shaft (in particular, the most upstream bearing). In addition, the optimization of the positioning of the path of the high-pressure body associated with the improvement of the cooling efficiency of the high-pressure turbine blades and the improved active control of the clearances at the blade tips in the high-pressure turbine makes it possible in particular to reduce the number of stages of the high-pressure compressor from 10 to 9.

[0191] The reduction in the number of low-pressure turbine stages (from 4 to 3) is also associated with an increase in the reduction ratio of the reduction mechanism and an increase in the fan rotor speed. The compactness of the reduction mechanism can be maintained by improving materials (e.g., using ceramic rolling elements (bearings)).

[0192] The expected gain of Engine 2 compared to Engine 1 is reduced consumption installed on the aircraft due to reduced frictional drag (reduced diameter of the fan rotor and reduced length of the engine) and reduced mass.

Claims

1. An aviation propulsion system (1), comprising: a first turbine (8) configured to drive the first compressor (4) about an axis of rotation (X) via a first shaft (11); a second turbine (7) configured to drive the second compressor (5) via a second shaft (10), the second shaft (10) being configured to rotate about the axis of rotation (X) at a higher speed than the first shaft (11); - a fan rotor (9), said fan rotor being connected to a fan shaft (20); - a speed reduction mechanism (19) coupling the first shaft (11) and the fan shaft (20) so as to drive the fan shaft (20) at a rotation speed lower than that of the first shaft (11); The average radius of the hole of the second turbine (7) satisfies the following formula: R m_a_Min ≤R m_a ≤R m_a_MAX in: R m_a is the average radius of the hole of the second turbine (7), in millimeters (mm); 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 tip (21) and the leading edge (22) of the blade (14) of the fan rotor (9); BPR is the bypass ratio of the propulsion system (1) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions; T e is the maximum inlet temperature of the drive turbine (8) when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C); T ref =273K; n is the number of stages of the second turbine (7) and the second compressor (5); Gamma is the adiabatic coefficient of air; XN is the limit speed of the first shaft (11) in revolutions per minute (tr / min); and K=6.76,L=153.6m -1 .(℃) -1 / 2 .(tr / min) -1 , M=421mm.(℃) 1 / 2 , N1 = -11 millimeters (mm) and N2 = -21 millimeters (mm).

2. The propulsion system (1) according to claim 1, wherein: The average radius of the hole of the second turbine (7) also satisfies the following formula: R m_a ≤E*L HP 2 *XN*10 -9 +F1 Where: L HP is the distance between the inlet of the second compressor (5) and the outlet of the second turbine (7), in millimeters (mm); and E=3.15(mm.tr / min) -1 And F1 = 23 millimeters (mm).

3. The propulsion system (1) according to claim 1 or 2, wherein: The average radius of the hole of the second turbine (7) also satisfies the following formula: R m_a ≥E*L HP 2 *XN*10 -9 +F2 Where: L HP is the distance between the inlet of the second compressor (5) and the outlet of the second turbine (7), in millimeters (mm); and E=3.15(mm.tr / min) -1 And F2 = 13 millimeters (mm).

4. The propulsion system (1) according to any one of claims 1 to 3, wherein: The average radius of the hole of the second turbine (7) also satisfies the following formula: R m_a ≥G*FN*BPR*10 -4 +H wherein: FN is the thrust of the propulsion system (1) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions and is expressed in Newtons (N); BPR is the bypass ratio of the propulsion system (1) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions; G = 0.16 millimeters / Newton (mm / N) and H = 18 millimeters (mm).

5. The propulsion system (1) according to any one of claims 1 to 4, wherein: The average radius of the hole of the second turbine (7) satisfies the following formula: R m_a ≤I*FN*BPR*10 -4 +J wherein: FN is the thrust of the propulsion system (1) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions and is expressed in Newtons (N); BPR is the bypass ratio of the propulsion system (1) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions; I = 0.16 millimeters / Newton (mm / N) and J = 28 millimeters (mm).

6. The propulsion system (1) according to any one of claims 1 to 5, wherein: The hub-to-tip ratio of the second turbine (7) is between 0.77 and 0.

90.

7. The propulsion system (1) according to any one of claims 1 to 6, further comprising bearings (11a-11c) configured to center the first shaft (11) relative to the rotation axis (X), the bearings (11a-11c) comprising bearing mode dampers.

8. The propulsion system (1) according to any one of claims 1 to 7, further comprising an inter-turbine housing (24) extending between the first turbine (8) and the second turbine (7) and configured to support the second turbine (7) via a bearing assembly (11a).

9. The propulsion system (1) according to claim 8, wherein: The inter-turbine housing (24) includes a plurality of guide vanes (25) configured to straighten air flow at the inlet of the first turbine (8).

10. Propulsion system (1) according to claim 9, comprising between twenty and thirty guide vanes (25).

11. The propulsion system (1) according to any one of claims 1 to 10, wherein: The overall compression ratio of the propulsion system (1) corresponds to the ratio between the outlet pressure of the second compressor (5) and the inlet pressure of the fan rotor (9), and the overall compression ratio is greater than or equal to 40 and less than or equal to 70.

12. The propulsion system (1) according to any one of claims 1 to 11, wherein: The second turbine (7) is a two-stage turbine.

13. The propulsion system (1) according to any one of claims 1 to 12, wherein: The second compressor (5) comprises at least eight stages and at most eleven stages.

14. The propulsion system (1) according to any one of claims 1 to 13, wherein: The bypass ratio of the propulsion system (1) is greater than or equal to 10, for example, between 10 and 35 and including 10 and 35, for example, between 10 and 18 and including 10 and 18, for example, between 10 and 15 and including 10 and 15.

15. The propulsion system (1) according to any one of claims 1 to 14, wherein: The first turbine (8) comprises at least three stages and at most five stages.

16. The propulsion system (1) according to any one of claims 1 to 15, wherein: The first compressor (4) comprises at least two stages and at most four stages.

17. An aircraft comprising at least one propulsion system (1) according to any one of claims 1 to 16, said propulsion system being fixed to the aircraft via a mast.

18. A method for determining the size of a propulsion system (1), the propulsion system comprising a speed reduction mechanism (19) and a second turbine (7), the speed reduction mechanism coupling a first turbine (8) and a fan rotor (9) so as to drive the fan rotor (9) at a speed lower than that of the first turbine (8), the second turbine being configured to rotate at a speed higher than that of the first turbine (8), the size of the second turbine (7) being determined so that the average radius of the bore of the second turbine satisfies the following formula: R m_a_Min ≤R m_a ≤R m_a_MAX in: R m_a is the average radius of the hole of the second turbine (7), in millimeters (mm); 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 tip (21) and the leading edge (22) of the blade (14) of the fan rotor (9); BPR is the bypass ratio of the propulsion system (1) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions; T e is the maximum inlet temperature of the drive turbine (8) when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C); T ref =273K; n is the number of stages of the second turbine (7) and the second compressor (5); Gamma is the adiabatic coefficient of air; XN is the limit speed of the first shaft (11) in revolutions per minute (tr / min); and K=6.76,L=153.6m -1 .(℃) -1 / 2 .(tr / min) -1 , M=421mm.(℃) 1 / 2 , N1 = -11 millimeters (mm) and N2 = -21 millimeters (mm).

19. The size determination method according to claim 18, wherein: The size of the second turbine (7) is determined so that the average radius of the hole of the second turbine (7) also satisfies the following formula: R m_a ≤E*L HP 2 *XN*10 -9 +F1 Where: L HP is the distance between the inlet of the second compressor (5) and the outlet of the second turbine (7), in millimeters (mm); and E=3.15(mm.tr / min) -1 And F1 = 23 millimeters (mm).

20. The size determination method according to claim 18 or 19, wherein: The size of the second turbine (7) is determined so that the average radius of the hole of the second turbine (7) satisfies the following formula: R m_a ≥E*L HP 2 *XN*10 -9 +F2 Where: L HP is the distance between the inlet of the second compressor (5) and the outlet of the second turbine (7), in millimeters (mm); and E=3.15(mm.tr / min) -1 And F2 = 13 millimeters (mm).

21. The size determination method according to any one of claims 18 to 20, wherein: The size of the second turbine (7) is determined so that the average radius of the hole of the second turbine (7) also satisfies the following formula: R m_a ≥G*FN*BPR*10 -4 +H wherein FN is the thrust of the propulsion system (1) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions and is expressed in Newtons (N); BPR is the bypass ratio of the propulsion system (1) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions; G = 0.16 millimeters / Newton (mm / N) and H = 18 millimeters (mm).

22. A sizing method according to any one of claims 18 to 21, wherein: The size of the second turbine (7) is determined so that the average radius of the hole of the second turbine (7) satisfies the following formula: R m_a ≤I*FN*BPR*10 -4 +J wherein: FN is the thrust of the propulsion system (1) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions and is expressed in Newtons (N); BPR is the bypass ratio of the propulsion system (1) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions; and I = 0.16 millimeters / Newton (mm / N) and J = 28 millimeters (mm).

23. A method for manufacturing a propulsion system (1), comprising the following steps: - dimensioning the propulsion system (1) according to any one of claims 18 to 22; and - Manufacturing said propulsion system (1).

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